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PE RFCm.l"'vI AKCE PGTESTl:-\L or
HYDROGE.;\ F CE LED, AIHERE :\THI~C
CnL1SE AIHCHAFT
FINAL REPORT
\'OLD!YIE 2 - PHASE I SlTDIES
REPOHT ~O. GD/C-DCBGG-n().L'~
Prepared for
NATIONAL AERONAUTICS & SPACE ADMI0iiSTRATIO:\
MISSION ANALYSIS DIVISION
MOFFETT FIE LD, CAUFORN1A
Project Leader
by
CJENERAL DYNAMICS
CONVAIR DIviSION SAN DIEGO, CALJFOHNIA
CONTRACT NAS 2-3180
6 M:\ Y 1966
-----------
.' ·PEtiJ)0192 ; ~.-
\
... ; FOREWORD
This volume shows the accomplishments during Phase I (through the first quarterl, (Ica I prcsent3tion) of Contract NAS2-3180. The propulsion sccti()J1S of this \'olul11l'
;tt'l: l,!:tssified CONFIDENTIAL and are included under separate CO\'cr ~lS fkpor: ;\(1.
GD' C-DCB-G6-004/2A.
\
\ I
)
ACKNOWLEDGMENTS
The following contributed significantly to the contents of this volume:
C. J. Cohan G. Place B. W. Prior J. E. Butsko K. S. Coward R. W. Woodrey C. J. Tait A. M. Goldman C. R. Bernick M. L. French
Performance and Aerodynamics Synthesis Program and Performance Propulsion Aerodynamics Sonic Boom Sonic Boom Configuration Design and Evaluation Mission Analysis Economics Weights
ii
SUl\IlV1ARY
The purpose of Contract NAS2-3180 was to study the performance potential of a liquid
hydrogen fueled, commercial transport operating in the 1985-2000 time period. The contract was organized in two phases. Phase I (through the first quarterly oral pre
sentation) was a broad parametric study, the purpose of which was to select two configurations on which to perform more detailed studies in Phase II. This volume sho~\'s
the Phase I studies.
The main areas of study during Phase I were configuration shape, pr 9pulsion
for cruise at Mach numbers up to 12, mission analysis, sonic boom and cost.. The
main conclusions were:
1. A design range of 5,000 nautical miles and a passenger capacity of 200 passengers
are good compromises between. take-off weight, cost, projected paE'senger traffic
) and sonic boom considerations.
2. The most promising configurations for Phase II are: (a) a 70° delta \ving/circul:1r
body with integral tanks, and (b) a double delta (80c
'65) blended bod\ configur
ation. A third configuration, variable sweep "'ingicircular bod\, is retained ror
ilmIted study during Phase II.
3. The best cruise Mach number is 6-7. The best propulsion system is a turboia:l
ramjet. lvIainly because of engine cooling requirements, scramjets, at their
best cruise Mach number of 8 are not competitive with the turbofanramjet. Because of subsonic loiter requirements, liquid ox\gen burning propulsion S\'stems \". ~ - ~
(e. g., ejector ramjets) are not competitive. Propulsion data are included in a
separate volume (Report GDIC-DCB-66-004/2A).
4. For fairly small penalties in take-off weight, vehicle shaping and transonic i1":l
jectory can be chosen so that a maximum over-pressure of about 2.0 psf nee':]'''
at 1Iach 1...1 and less than 1. 0 psf during cruise.
iii
TABLE OF CONTENTS
1.0 INTRODUCTION......... .... . . . . .. .. . . . . .. .. . . . . . . .. . . . 1
1.1 OBJECTIVES ........ iii ••••••••••••••• I........................ 1 1.2 GROUND RULES ............... t • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • 1 1.3 APPROACH TO PliA.SE II .... 1.. ................................. 2 1.4 ORGANIZATION OF THIS REPORT. . . • • • • • . • . . . . . . • . • . . • . • . • • • . . . 4
2.0 DELTA WING BASELINE CONFIGURATION. • . • • • . • . • • . • . • • • • • ... 6
2.1 GROUND RULES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6 2.2 DE SIGN REQUffiE MENTS. . . . . • . • . • • • . . . • • . . . • . . • • • . • . . • . . . . • • . . 6 2.3 CONFIGURATION CONCEPTS................................... 7 2.4 CONFIGURATION EVALUATION AND SELECTION ................. 7 2.5 SELECTED CONFIGURA TION ARRANGEMENT.................... 9 2.6 AERODyNAMICS............................................... 11 2.7 PROPULSION ..........•.....•............•................ ' •... 14 2.8 WEIGHTS ..............•.....•........•.•...................... 14
3.0 SYNTHESIS PROGRAM. . . . . • • . . . • . . . • . . . . . . . . . . • . . . . . . . .. 19
3 . 1 DE SCRIPTION. . . . . . • . . . • . . . . . . . . . . . . • . . . . . . . . . . . . . . . . . . . . . . . .. 19 3.2 COMPUTATIONAL TECHNIQUES. . . . . . . . . . • . . . . . . . . . . . . . . . . . . . .. 21
3.2.1 WEIGHT/SIZING SUBROUTINE ........................... 21 3.2.2 AERODYNAMIC SUBROUTINE. . . . . . • . . . . . . . . . . . • . . . . . • . .. 22 3. 2. 3 PROfULSION SUBROUTINE. . . • . . • . . • • . • . . . . . . . . . . . . . . . •. 23 3.2.4 PERFORMANCE SUBROUTINE .....•.......•............. 24
4.0 TRAJECTORY EVALUATION. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 25
4.1 TRAJECTORY EVALUATION GROUND RULES .............•...... 25 4.2 SONIC BOOM VARIA TIONS. . . . • • . . . . . . • •. . • . . . .. . . . . . . . . . . . . . . .. 27 4.3 DYNAMIC PRESSURE VARIATIONS ....•.............•........... 27 4.4 INLET PRESSURE VARIATIONS ..................•.. : .•.•....•.. 30 4.5 CRUISE ALTITUDE VARIATIONS........... ......••....•.•...... 30 4.6 DESCENT TRAJECTORy.................................... .•. 35 4. 7 SUBSONIC LOITER AND CRUISE VARIA TIONS. . . . . . . . . . . . . . . . • . .. 35 4.8 SELECTED TRAJECTORY. .... .•... ..•.•. . ... ... ...... .. .... ... 38
5.0 PROPULSION EVALUATION, CRUISE MACH < 8 .................. 41-76 (See GD/C-DCB-66-004/2A)
iv
TABLE OF CONTE~'TS (cont'd)
6.0 DELTA WING CONFIGURATION VARIATIONS ...................... . 81
6. 1 GROUND RULES .....•........ '" . . . ... ..... . .. .. . . . . ... . ... 81 6.2 WING GEOMETRY VARIATIONS......... .......•... .. ........ 81 6.3 BODY GEOMETRY VARIATIONS.................... .......... 87 6.4 INTEGRAL TANKS .......•..•.............•............•. .'.. 87 6. S FINAL DELTA WING CONFIGURATION. . • . . . . . . . . . . . . . . . . . . .. 87
7.0 VARIABLE SWEEP CONFIGURATION....... ......... . .............. 98
7. 1 CONFIGURATION EVALUATION. . . . • . . . . . . . . . . . . . . . . . . . . . . . .. 98 7.2 SELECTED CONFIGURATION ARRANGEMENT................. 99 7.3 AERODYNAWCS. . . . . . . . . . .. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 99
7.4 WEIGHTS ...................•..........•............•....... 106 7. S WING GEOMETRY VARIATIONS ....•......................... 106 7.6 FINAL VARIABLE SWEEP CONFIGURATION ................... 110
8.0 BLENDED BODY CONFIGURA TION. . . . . . . . . . . . . . . . . . . . . . . . .. 123
8.1 BACKGROUND................................................ 123 8.2 GROUNDRULES ..•...•....•..............•................... 123 8.3. CONFIGURATION CONCEPTS ............•...................... 124
8.4 DOUBLE DELTA CONFIGURATION .............................. 124 ~ 8.5 VARIABLE SWEEP, OVERLAPPING WING CONFIGURATION ....... 127
8.6 AERODYNAlVIICS .........•........•.•.............•............ 127
8. 7 WEIGHTS ...............•..............•...................... 129
8.8 DOUBLE DELTA - GEOMETRY VARIATIONS ..................... 129 8.9 FINAL DOUBLE DELTA CONFIGURATION ....................... 129 8.10 VARIABLE SWEEP WING GEOMETRY VARIA TIONS ............... 129 8.11 FIKAL VARIABLE SWEEP CONFIGURATION ...................... 130
9.0 PROPULSION EVALUATION, CRUISE ]'vIACH >8 (SCRAMJET) ........ .l62-187 (See GD/C-DCB-66-004/2A)
10.0 COMPARISON OF CONFIGURATIONS .......................... 199
11. 0 MISSION ANALYSIS ....................•................. 202
11.1 SELECTION OF DESIGN RANGE ................................ 202 11. 1. 1 PASSENGER TRAFFIC 1985-2000 ....................... 202 11.1. 2 GEOGRA..PHIC DISTRIBUTION OF PASSENGERS .......... 204
11.1.3 RECOl\IMENDED DESIGN RANGE ....................... 209
11.2 CRUISE MACH NUMBER .............•..... ~ ................... 214
11.2.1 BLOCK TITvIE ...............•......................... 214
11.2.2 EFFECTS OF LOCAL TIME ............................ 217
11. 2.3 UTILIZATION .....•...........••.•..••................ 217
11. 2. 4 BEST CRUISE MACH NUMBER. ................•....... 223
11.3 PASSENGER CAPACITY ••........•..•...................•..... 223
11.4 CONCLUSIONS ................................................ 226
12. 0 SONIC BOOM ......................•..............••... 229
12. 1 INTRODUCTION.............................................. 229
12.2 METHOD OF ANALYSIS - FAR FIELD .......................... 230
12.3 METHOD OF ANALYSIS - NEAR FIELD ........•................ 231
12.4 DISCUSSION AND RESULTS .................................... 232
12.4.1 DELTA WING CONFIGURATION ........................ 233
12.4.2 VARIABLE SWEEP CONFIGURATION. . . . . . . . . . . . . . . . .. 233
12.4.3 BLENDED BODY CONFIGURATION.................... 233
12.4.4 SCRAMJET CONFIGURATION...................... ... 233
12.4.5 SUPERSONIC TRANSPORT ......•.................•... 243 12.4.6 COMPARISON OF CONFIGURATIONS ................... 243
12.4. 7 CON FIGURA TION VARIA TIONS. . . . . . . . . . . . . . . . . . . . • . .. 243
12.5 MISSION SONIC BOOM CHARACTERISTICS ................. 252
12.6 CONCLUSIONS .........................•..•...•......•........ 252
12.7 NOMENCL.A TURE 255
13.0 COST .............................•................. 257
13.1 COMPUTER PROGRAM ..................................•..... 257
13.2 ASSUl\IPTIONS .........•...•................................•. 259
13.3 TYPICAL COST ..•.........•.............................•.... 260 13.4 COST SENSITIVITY ............................................ 260
14.0 SE LECTION OF PHASE II CONFIGURATIONS ......•.....•.....•....... 266
14.1 PASSENGER CAPACITy •...•..............•..........•...•.... 266 14.2 DESIGN RANGE ..................•..•..............•.......... 266 14.3 CO:r-.,TFIGURATION SELECTION ....•....••...•.....•........•.... 268
15.0 LIST OF REFERENCES .....•....•...•........•.........•............ 271
vi
/
LIST OF FIGURES
Page No.
1. APPROACH TO PHASE I 3
2. DELTA WING CONFIGURATION CONCEPTS 8
3. BASELINE DELTA WING CONFIGURATION 10
-1. DELTA "'lING CONFIGURATION - AERODYNAl\IIC CHAHACTERISTICS 12
;). DELTA \\>1NG CONFIGUR~TION - DRAG BREAKDOWN 13
6. DELTA \\1NG CONFIGURATION - AERODYNAMIC CENTER 13
7. SYNTHESIS PROGRAM FLOW 20
8. TYPICAL TRAJECTORY 2 G
9. TRAJECTORY VARIATIONS 28
10. EFFECT OF SOl\TIC BOOM ON TAKE-OFF WEIGHT 29
11. EFFECT OF DYNAMIC PRESSURE ON TAKE-OFF WEIGHT 29
1 '). EFFECT OF RAM,JET AND TITRBQTET SIZE ON TAKE OFF WEIGHT 31
13. EFFECT OF INLET PRESSURE ON TAKE-OFF WEIGHT 32
14. EFFECT OF CRUISE ALTITVDE ON ASCENT MASS RATIO 33
15. EFFECT OF CRLHSE ALTITUDE ON TAKE-OFF WEIGHT 34
'16. EFFECT OF TURBOJET AND RAMJET THROTTLING 01\ TAKE-OFF 3(i WEIGHT
17. EFFECT OF DESCENT TRAJECTORY ON TAKE-OFF WEIGHT 3G
18. EFFECT OF SUBSOmC LOITER ALTITVDE AND l\IACH NO. 37 ON RA..."1\GE PARAMETERS
19. J;:FFECT OF SUBSONIC LOITER ALTITUDE AND MACH NO. 37 ON FUEL FLOW
20. SE LECTED TRAJECTORY -10
vii
'\ I
LIST OF FIGURES (Cont'd) Page No.
21 - SEE GD/C-DCB-66-004/2A 53.
54. TAKE-OFF WEIGHT VS, LEADING EDGE SWEEP 83
55. TAKE-OFF WEIGHT VS, THICKNESS RATIO 84
56. TAKE-OFF WEIGHT VS. ASPECT RATIO 85
57. TAKE-OFF WEIGHT VS, WING LOADING 86
58. TAKE-OFF WEIGHT VS, BODY FINENESS RATIO 88
59. VARIABLE SWEEP WING CONFIGURATION 100
60. VARIABLE SWEEP CONFIGURATION - SWEEP SCHEDULE 101
61. VARIABLE SWEEP CONFIGURATION - AERODYNAMIC 102 CHARACTE ruSTICS
" 62. )
VARIABLE SWEEP CONFIGURATION - DRAG BREAKDOWN 103
63. VARIABLE SWEEP CONFIGURATION - TRIM: AT MACH 0.5 104
64. VARIABLE SWEEP CONFIGURATION - TRIM AT MACH 4.5 105
65. TAKE-OFF WEIGHT VS, LEADING EDGE SWEEP 109
66. TAKE-OFF WEIGHT VS, THICKNESS RATIO 109
67. TAKE-OFF WEIGHT VS, ASPECT RATIO
68, TAKE-OFF WEIGHT VS, WING LOADING
69. BLENDED BODY CONFIGURATION CONCEPTS 125
70. BLENDED BODY DOUBLE DELTA CONFIGURATION 126
71. BLENDED BODY VARIABLE SWEEP CONFIGURATION 128
72. DOUBLE DELTA AERODYNAMIC CHARACTERISTICS 131
73. DOUBLE DELTA DRAG BREAKDOWN 132
74, VARIABLE SWEEP AERODYNAMIC CHARACTERISTICS 133
75. VARIABLE SWEEP DRAG BREAKDOWN 134
) viii
LIST OF FIGURES ICont'd)
76. DOL'BLE DELTA LONGITUDINAL STABILITY
77. DOUBLE DELTA PLA...NFORM LOADING EFFECTS
78. VARIABLE SVlEEP WING LOADING EFFECTS
7;') - SEE GD/C-DCB-66-004/2A
9;') •
Page No.
135
142
96. WORLD TOTAL ANNUAL TWO WAY TRAFFIC 203
97. GEOGRAPHIC DISTRIB"CTION OF PASSENGERS 206
98. DISTRIBVTION OF PASSENGER TRAFFIC 207
~)~). PASSENGER TRAFFIC VS. RANGE 208
lOO. CUTvIT.::LATIVE TRAFFIC VS. RANGE 21()
101. ROUTE STRUCTURE 213
102. TRIP TlliIE VS. RANGE 216
103. TRIP TThI E 222
10-1. lTTILIZA TION 224
lOS. PASSE:'\GERS LEA VTNG NEW YORK/DAY 22;:;
lOti. PASSENGERS LEAviNG NEW YORK VS. NO. OF SEATS REQ'D. 228
107. SONIC BOOM CHARACTERISTICS - DELTA WING 234
CONFIGDRATION
108, F (T) FrNCTION - DELTA WIKG CONFIGDRA TION 235
109. SONIC BOO:i\I CHARACTERISTICS - VARIABLE SWEEP 236
CONFIGU RATION
110. F (T) FUNCTION VARIABLE SWEEP CONFIGURATION
111. SO:\1:C BOO:;\I CHARACTERISTICS BLENDED BODY
CONFIGl~RATION, SINGLE DELTA
112. SONIC BOOl\l CHARACTERISTICS BLENDED BODY
CONFIGl~RATION, DOUBLE DELTA
237
238
239
113. FIT) FrNCTION BLENDED BODY CONFIGL'RATION, 240
SINGLE DELTA
114. F (T) I-TNCTION BLENDED BODY CONFIGFHATION, 241
DOCBLE DELTA
ll;,). SO~lC BOO:vr CHARACTERISTICS SCRA::\lJET CONFICrRATIO::';- 242
ix
LIST OF FIGURES (Contrd) Page No.
116. F (T) FUNCTION SCRAMJET CONFIGURATION 244
117. SONIC BOOM CHARACTERISTICS - SUPERSONIC TRANSPORT 245
118. SONIC BOOM CHARACTERISTICS - CONFIGURATION 246 COMPARISON
119. EFFECT OF BODY FINENESS RATIO ON SONIC BOOM 248
120. EFFECT OF BODY SHAPE ON TAKE-OFF WEIGHT 249
122. EFFECT OF WING LOADING ON SONIC BOOM 250
123. EFFECT OF ASPECT RATIO ON SONIC BOOM 251
124. EFFECT OF TAKE-OFF WEIGHT ON SONIC BOOM 253
125. MISSION SONIC BOOM CHARACTERISTICS 254
126. COST MODEL SCHEMATIC 258
127. COST SENSITIVITY 265
x
LIST OF TABLES
1. STCDY OBJECTIVES 1
2. STUDY GROVND RCLES 2
3. DELTA WING CONFIGURATION GROU:-iD RULES 6
4. BASELINE DELTA WING CONFIGURATION - WEIGHT & SIZLNG DATA 15
5-8. (SEE REPORT GD/C-DCB-66-004/2A)
9.
10.
11.
12.
13.
DATA FOR INTEGRAL VS NON INTEGRAL TANKS
FINA L DE L TA WING CONFIGURA TION DATA
VARIABLE SWEEP CONFIGURA. TION BASELINE WEIGHT & SIZING DATA
FI~AL VARIABLE SWEEP CONFIGURA. TION DATA
BLENDED BODY CONFIGURATION GROUND RULES
14. BLENDED BODY, DOGBLE DELTA, BASELINE WEIGHT & SIZING
15.
16.
1 .., or i.
DATA
BLENDED BODY, VARLA.BLE SWEEP, BASELINE WEIGHT & SIZING DATA
BLE~DED BODY, DOUBLE DELTA FINAL DATA
BLENDED BODY, VARIr\BLE SWEEP FINAL DATA
18-23 (SEE REPORT GD/ C-DCB-66-004/2A)
22.
23.
24.
25.
26.
?-, -I.
28.
29.
30.
COl\IPARISON OF CONFIGURATIONS, CRUISE MACH 3-8
FII'\AL CONFIGURATION DATA
GLOBAL TRANSPORT AIR TERMINALS
POSSIBLE ROUTE WITH A 5,500 N. MI. RANGE
OTHER POSSIBLE ROUTES WITH A 5,500 N. l'vII. RIl..NGE
DEFINITION OF TYPICA.L TIMES
LOGA. L TIMES
LOCAL TIl\IES
TYPICAL SCHEDULE LAX/NY/ROl\IE
xi
89
90
107
114
123
13 ~)
143
153
200
201
205
211
212
215
218
219
220
LIST OF TABLES (Continued) ~
Page
31. TYPICAL SCHEDULE LAX/TOKYO 221
32. COST PRINTOUT 261
33. SELECTION OF PASSENGER CAPACITY 267
34. SELECTION OF DESIGN RANGE 267
35. SELECTION OF PHASE II CONFIGURA TIONS 269
36. PHASE II STUDY AREAS AND SELECTED CONFIGURATION 270
xii
,
1. 0 Il\"TRODCCTION
Starting 7 Septem her 1%5, the Mission Analysis DiYision, :Moffett Field, awarded contract NAS 2-3180 to General Dynamics Convair, San Diego. This contract was to study the "Perfonnance Potential of Liquid Hydrogen Fueled, Airbreathing, Cruise Aircraft'·. The contract was for 74 man months of effort over the nine month duration of the contract. Additional background to the study and to this report is as follows:
1. 1" OBJECTIVES
The objectives of the study, as defined in RFP A-I05n (WEB 32), are shown in Table 1. Items 1 and 2 were defined as Phase I and Items 3 through 6 as Phase II. Phase 1,
which is the subject matter of this volume, was completed during the first three months of the study.
l. TO INVESTIGATE A WIDE VARIETY OF I.J-L) Fl'ELED, '- -
1 A1RBRE ATHING, CRUISE AIRCRAFT.
PHASE I
2. TO SELECT TWO PR01\IISING CONFIGt"R_A nONS.
3. TO EXAMINE, IN DETAIL, THE PROBLE1\l AREAS OF THE SE LECTED CONFIGl:RATIO:\S.
4. TO PROVIDE A DETAILED DEFINITIO?\ OF THE FINAL PHASE II CONFIGl:RA TIONS.
5. TO PERFORlIvl SENSITI\;1TY STCDIES.
6. TO DEFI:\E CRITICAL RESEARCH AREAS. ...
TABLE 1. STl'DY OBJECTIVES
1. 2 GROlJND Rl'LES
The ground rules for the study are shown in Table 2. These ground rules \vere originally specified in RFP A-I0597, but were modified as recorded in Reference 3-principally i:1 extending the upper limit of cruise ;\I2ch number from 8 to 12.
1
FUEL: LIQUID HYDROGEN
PROPULSION: AIRBRE ATHING
MISSION: COMMERCIAL TRANSPORT
OPERATIONAL: 1985-2000 TThIE PERIOD
CRUISE MACH: 3 TO 12
SONIC BOOM: ~ 2 PSF DURING CLIMB 1. 5 PSF DURING CRUISE
TAKE-OFF: 160 KNOTS & ~ 10,500 FT.
LANDING: 135 KNOTS &~ 8,000 FT.
TABLE 2. GROUND RULES FOR STUDY
1. 3 APPROACH TO PHASE I
As shown in Tables 1 and 2, Phase I was a broad parametric study of the overall characteristiC$ of a liquid hydrogen fueled, commercial transport. The end result of Phase I was to select two promising configurations for more detailed studies during Phase II.
The approach to Phase I is shown in Figure 1. To provide a firm basis for selecting the configurations for Phase II, four parallel areas were studied:
a) Configuration/performance
b) Mission analysis
c) Sonic boom
d) Cost.
The results of these studies were then combined into an overall system evaluation from which the selected configurations for the Phase II studies were obtained.
The upper portion of Figure 1 shows the configuration/perfonnance study, the purpose of which was to compare, parametrically, all reasonable combinations of trajectory, propulsion and configuration. The primary evaluation tool for these studies was an IBM 7094 computer program that combined aerodynamics, weights,
2
~
,
. riAERODY "AMICS t--
THAnE OF1"S ~PROPULS ION L
BASELINE • THA,JECTOHY r-+ ~ VEHICLE r. • PROPULSION 1\1'" K CONFIGS. HSIZE : SYNTHESIS
• VEHICLE SHAPE
~WEIGHT.." • SCHAMJET
11r
OVERALL SELECT
w SYSTEM CONFIGS
l MISSIO N ANALYSIS .. EVALUATION ~ FUR .. PHASE 11
t .t~
I SONIC BOOM
I SYSTE~ ,f ECONOMICS I I
FIGUHI :l. APPHOACII TO PHASE I
size, propulsion and performance. In order to ensure that reasonable results would be obtained from the parametric studies, the first step was to design a "baseline configuration". Aerodynamic, propulsion, size and weight data were then obtained by analysis of this design. Sizing equations were then derived that included the values obtained for the baseline vehicle. This baseline vehicle - analysis - computer equation approach ensured that the parametric study used realistic values for volume utilization, wetted areas, exposed wing areas, inlet pressure field, longitudinal and directional stability, etc. The configuration/performance evaluation comprised about 8090 of the Phase I effort.
1. 4 ORGANIZATION OF THIS REPORT
This report shows the studies performed during the first three months of the contract (Phase I of Figure 1) which culminated in an oral presentation at Convair, San Diego on 5 January 1966. The decisions made at that meeting are incorporated.
This report is organized along the lines that the study was conducted, i. e. ,
1) Early in the study a delta wing/body configuration was defined (Section 2.0). Ground rules for sizing of this configuration were to carry 200 passengers for 5,000 nautical miles at a cruise Mach number of 6: These values and the configuration were judged to provide a reasonable basis for determination of the best trajectory up to Mach 6.0 (Section 4.0) and for evaluation of best accelerator propulsion for the Mach 0 to 3 range (Section 5.0). In addition cruise Mach numbers between 3 and 8 were investigated. (Section 5.0).
2) Geometry variations of the delta wing/body configu-r'ation were then made to find the optimum wing and body shape (Section 6.0).
3) A variable sweep wing/body configuration was defined and wing planform studies were made (Section 7. 0) .
4) Blended body configurations were defined and optimized (Section 8.0).
5) An evaluation of configuration/propulsion concepts for cruise Mach numbers between 6 and 12 (Scramjets) was made. This included trajectory and cruise Mach number variations (Section 9.0).
6) The best configurations as determined in Sections 4.0 through 9.0 were then compared (Section 10.0).
7) In parallel with the vehicle studies, a mission analysis was conducted to determine, from the mission standpoint, the most promising cruise 1\1 ach num oor and design range (Section 11. 0).
4
6) Because of its significance on public acceptance, plus the effect on trajectory, engine siz ing and vehicle shape, an analysis of sonic boom was made (Section 12.0) .
9) Elementary sy"stem costs were determined (Section 13.0).
10) All of the significant factors as determined from the configuration, mission analysis, sonic boom and cost studies were then incorporated into an over-all system evaluation from which the most promising configurations for Phase II were selected (Section 14.0).
Sections 5.0 and 9.0 on Propulsion Evaluation are included under a separate cover, Report GD/C-DCB-66-004/2A (CONFIDENTIA.L).
5
i-
2.0 DELTA WING BASELINE CONFIGURATION
Of the four basic configuration types to be investigated in Phase I, the relatively conventional delta wing/body arrangement was chosen for initial definition. This configuration was defined early in the ~study and was used for the trajectory evaluation (Section 4.0) and for the propulsion evaluation, Mach < 8.0 (Section 5.0).
-2. 1 GROUND RULES
In order to commence the design of the delta wing configuration, a set of ground rules were established as shown in Table 3. These ground rules were selected as providing a vehicle size roughly in the middle of the expected variations. As indicated in Table 3, the take-off weight was expected to be about 600,000 lb. and the fuel weight to be about 40% of the take-off weight.
TAKE-OFF WEIGHT 600,000 LB. APPROX. ~\ ~Q T . '.
WEIGHT OF FUEL 240,000 LB. APPROX. ~
CRUISE MACH NO. 6.0
NUMBER OF PASSENGERS 200
WEIGHT OF CARGO 5,000 LB. PLUS 40 LB. /PASSENGER
RANGE 5,000 N.M.
TABLE 3. DELTA WING CONFIGURATION GROUND RULES
2.2 DESIGN REQUIREMENTS
In order to establish a reasonable configuration, the following aspects of the vehicle were considered:
a) Adequate control surface areas
b) Acceptable C. G. location
6
)
c) Location and stowage of the landing gear
d) Overall volume utilization (passenger, crew, cargo, fuel, systems)
e) Propulsion installation effects
f) Pilot visibility.
2. 3 CONFIGURATION CONCEPTS
Since this was the first configuration to be evolved, initial studies were made in the following areas:
BODY
A study was made to establish the most advantageous relationship of passenger, cargo and LH2 tankage with emphasis on passenger safety, drag, and volume utilization.
WING --Three basic locations of wing were studied--Iow, mid, and high relative to the body. This revealed the relationship of the propulsion system, landing gear, and wing/body structural integration.
From these basic wing/body relationships a series of configurations were evolved. The most significant combinations are shown in Figure 2.
2.4 CONFIGURATION EVALUATION AND SELECTION
For the configurations shown in Figure 2 the main areas of compromise are:
a) Passenger/cargo compartment location
b) Lengtl.l. of the main landing gear
c) Tread of the main landing gear
d) Stowage of the main landing gear
e) Inlets in the pressure field
f) Minimum inlet flow interference during M LG retraction
g) Wing carry through structure. 7
)
\\"ING LOCATION
• LOW
MIDDLE
HIGH
TOP
A iii "
T lI(QF
:Jf' 11
BOTTOM
~-----------==~~ /7 =<=:J
f
-A-
FIGtJRE 2. DELTA WING CONFIGUR.\. TION CONCEPTS
8
Pl .. SSENGER COMPARTMENT LOCATION
SIDES CENTER
\. The low wing layout allows the engines to be close to the vehicle centerline, ensuring that the inlets are in the wing pressure field. 'This allows the main landing gear to be located outboard of the engines, minimizing interference with the inlet air flow during retraction, but still providing a reasonable wing depth for gear stowage. The low wing layout also allows the wing center section to pass underneath the fuel tanks. Considering the mid wing layout with a wing carry through box incorporatedexcessive tankage volume is lost. Alternatively a heavy center section results when the wing bending moments are carried directly by the 25' diameter frames.· Both the mid wing and the high wing layouts are poor in terms of achieving adequate tread for the main landing gear. Stowage of the gear in the outer wing is not possible because of insufficient depth, stowage in the body results in excessive lost volume around the circula). fuel tanks.
Each of the columns in Figure 2 show a different locati(ln of the passenger compartment. The upper compartment, column 1, has about the best Gompromise. The low passenger compartment, Column 2, results in excessive 16st volume around the ci1(cular non-integral tanks (or requires the engines to be moved sp.~nwise). In addition, this low location of the passenger compartment interferes with the wing carry through box and is noisy because of the engine proximity. The forward compartment shown in Column 3 is probably the safest in the event of a crash, but with passengers weighiI\g 10-15% of the landing weight, a C.G. shift (zero/full passengers) of more than 20% M1A.c is not reasonable. The side compartment location shown in Column 4 is thought td be more applicable for the blended body configuration and does not appear to have any advantages over the top compartment location. Column 5 shows the passenger compartment located in mid-fuselage and on the C.G.' This is probably the lightest structural arrangement because the passenger ~compartment has the least wetted area and the unusable volume around the fuel tanks is minimized. The three-story, 10
abreast seating may be acc:eptabIe , although the separation of crew and passenger com partment does not seem desirable. This arrangement may be considered in Phase II as one of the structural trade-offs.
Of the five basic compartment locations, arranging the compartment along the upper body centerline was judged best for passenger safety, noise, and structural compatibility.
The low wing offers the best engine installation, landing gear stowage and structural integration with the body. Therefore the low wing - top passenger compartment configuration was selected.
2.5 SELECTED CONFIGURATION ARRANGEMENT
Figure 3 shows a plan and inboard profile of the baseline delta wing configuration. The body volume is dominated by the four, circular, non-integral tanks. Tanks 1, ?, and 4 are conical--closely paralleling the outside contour of the body; tank 3 is a cylindrical and less than body diameter to allow the wing box to carry through. The external lines
9
\ )
FIGURE 3.
-~-i
-----
200 PAsSENGEJ:1 CQMPT
_._---------- ---------
BA.SELINE DELTA WING CONFIGURATION
. ~;~~( "-/ /',- \ , - - ,,- -. ,-,- .- -- - ,- -, - -
"./ i :
"30'
Wing area (gross) thickness ratio aspect ratio leadi ag edge sweep
HorIZontal area (gross) thiFlmess ratio aspect ratio mOVememt
Vertical ,. r ea thibkness ratio aspect ratio
Body length (overall) breadth (max- ) depth (max. ) volume tuel volume
Landing gear - main nose tread
Engines type thrust
Passenger ~ 6 abreast • " abreast seat spacing
Carll" b.y volume
D-D
5,.'
IOe'" -~PAN
7,500 sq. ft. -04 1. 45 70°
1,410 sq. ft· 06
2.1 _150 to + 10°
1,320 sq. ft. .06 .8
335 ft. 25·3 ft. 34· 3 ft. lO6,66O cu· It· 52, 600 cu. it.
four 66 X 18 tires per oleo twin 48 X 12 tires 52 ft.
Pratt I< Whirney STFRJ-230A (4) 75,000 lbs. SLS each
24 rows 14 rows 36 inches
900 cu. ft.
!Lr:; u.u
'::
'\ )
)
of the body are roughly 3/4 power with the maximum thickness at 60(~ of the length. The passeni1.er compartment is on top of the fuel tanks with 4 and 6 abreast seating and terminates at about station 190. Aft of the cabin is a cargo bay sized for 40 lb. of baggage per passenger at 20 lb/cu. ft. plus 5,000 lb. of cargo @ 10 lb/cu. ft. The four turboramjets with two dimensional inlets are located tmder the aft center wing. The main landing gear retracts inboard and has a dual hinged strut mechanism for gear
. stowage within the wing contours. The nose is designed to droop to allow pilot visi-, bility during landing.
2.6 AERODYNAMICS
Aerodynamic characteristics of the baseline delta wing configuration were analyzed at Mach numbers from 0.5 to 6. O. Primary emphasis was placed on lift and drag since these data were required for the aerodynamic subroutine of the synthesis program. The analysis procedures outlined in USAF DATCOM were used, supported in part with experimental data on cross section shape, body camber, etc., from References 8 through 11. The resulting lift, drag and L/D characteristics are shown in Figure 4 for several flight Mach numbers. (The increasing L; D max. with increasing Mach number is the result of a relatively high zero lift transonic drag and is characteristic of bodies with large cross sectional areas: e. g., Reference 25). Figure 5 presents a breakdown of the total drag coefficient throughout the flight l\Iach number regime for a lift coefficient of . 15. These data show an appreciable percentage of the drag is due to the afterbody and friction.
A brief examination of the stability characteristics of the configuration was made. Figure 6 presents the Mach number variation of longitudinal aerodynamic center. The forward shift of aerodynamic center in going from transonic to hypersonic speeds is a result of the large body SIze relatIve to the lIftmg surface area. Since the lift on the wing and tail fails off faster with increasing Mach number than the lift on the body, increasing the body size with respect to the wing tends to increase the aerod} namic center shift.
A brief examination of the trim capability of the configuration showed that a large negative (nose down) moment exists at zero lift throughout the flight regime, due mainly to the negative lift of the forebody at zero angle of attack. This condition can be corrected by cambering of the fuselage, which will be studied more thoroughly in Phase II. The horizontal tail surfaces indicate sufficient trim power to provide trim at desired lift coefficients, once the zero-lift moment is removed. The effects of trim settings on LID will also be studied during Phase II.
I
The directional stability was checked at Mach 6.0 and it was found that the configuration was directionally stable for angles of attack up to 15 degrees.
11
l\1ACH .6
;;i:n::·~:;::k.!~~=~:~d~ 1.5 .6
.5 ......... .......... ...... .5
.5
.4 .4
C .3 .3 L
.2 .2
.1 .1
0 0 0 4 8 12 0
)
LID
o Ci.
) FIGURE 4. DELTA WING CONFIGURATION - AERODYNAMIC CHARACTERISTICS
12
)
.05
.03
.02
.61
SREF = 7,500 sq. ft. ---C--"--- -. ----;--- --------.• ---.----------------------. CCI -.cc __ ,.,..~
lit ___ i.. __ .. __ ... ! • .. ------r" .-~ ". ~ , .. ~-.o.- ___ 4.-... __ j.
! C 1S'.:
--'---+-~ ,
.. t "~ *." .. _ L. , i i
--;._---;- -- ---t·---·-T-~' ,
'-'~---.~-'--- --i- L ,,=. ----1 _ .. _____ :" __ :_:_1
"\_. .j . .1. .J
: ___ L __ ~_~_l~~j , .. -::- ['-----,1: ___ i ..... •. ,.,.::cc:-: ,.,.' :-;.i_: :cc:":,.,.: ',.,.' :--,1
_--:- t· -, -- I ., 'j
- 1--- ...:.----tc--,--...:.-i !.~
-- __ .L __ ..... _ 1. ________ 1. ____ ---Wing: ;.
,
Tail---
-l. __ ~ ! .: -~~
--- + .. _---; Lift, : ----+- ------, ,
I ,
------t-·>-- j -'----'-......--.4
I • 1
. ----_.! --- ----.~ !
I I I I
-- f t:- : ..... , Friction ,,--' , 1 1 1
O~--~--~--~--~--~L-~--~--~--~--~--~--~--.. --~
A.ERODY?-J.
CEi'\TER
(FT.
200
190
FROM NOSE)
180
170
2 .J
}lACH NO.
5 6
FIGCRE S. DELTA WING CONFIGCHA nON DR'\G BREAKDOW:\,
, -- L-----T--.~---;-- __ . ___ -+_~ __ ___i
. ' .
. ---~-------.:------~ .. ~-~ . . -. . !
L .-~ . ---.- .. - -t- L. -~ . -- ~ --L ____ -l __ L ____ .~ __ · L_· _i_:~
T t ' i ,i I i -+~. ___ ._~--- .• -~ _. ___ .,j
: , t· J
~;---~.--r-----r--:-.--~-----"''C:i--~~~------:-+----:--.... I -: ~~. j : ) , .. . ~ . i - r.--------f---~ -~ .. -. _~..L_. ____ ~ __ ~ _ _..:......: ___ : • .:. ____ ~
_, . __ 1 ~- ... +':.' . ........:_ .......... ~---i- --·r~---~. ---.. --.=-.--~.~--""oo;t:::_-~:___~ -Y--" •
-. ~_. ___ ._. L - . _-l. _ -1. . __ ~l I 1
-......;--~-=--=-.:.;..' ---~-----.---:---j I I ; i I
._~ _______ .1 ____ .. __ 1 , __ ~. ___ ~ ___ ....:...._._~ ... _: ___ ._:- .. -- L .. _. _. __
______ ._! ___ i __ · .~j __ ~ ___ : ___ j .. __ . __ ._~ __ .. _ i .1 __ . ____ .L _______ .;.._
\ : , .
o 3 6
;\L\ CH 0;()_
7
FIGLRE G. DE LTA \\'I?\G CO>.'flCLH.-\ TIt):\ .. :\ERODYl\_'\~iIC CENTER
)
2.7 PROPULSION
For purposes of vehicle sizing, four Pratt and Whitney STFRJ-230A turboramjet engines were assum ed. This engine is a twin spool, axial turbofan with a bypass ratio of 2.15 : 1. The engines have a fuel rich primary burner and the bypass flow mixes with the fuel-rich turbine exhaust in the afterburner section to produce overall stoicll:'" iometric combustion. For ramjet operation the turbofan is shut down and air is bypassed around it to the afterburner section which now functions as the ramjet combustor. The nozzle is an expansion-deflection type with variable throat area.
Each of the four turboramjets are supplied by a two dimensional, mixed compression inlet. Since these two dimensional inlets can be installed in much less spanwise width than four, separate axi-symmetric inlets, these two dimensional inlets give a more compatible installation within wing pressure field and vehicle balance constraints.
2.8 WEIGHTS
Contract NAS 2-3025, "Weight and Size Analyses of Advanced Cruise and Launch Vehicles", was used for the deterinination of the weights and sizing of the delta wing configuration. "First level" i. e., least sophisticated data, were used since these were consistent with the broad nature of the Phase I studies. Table 4 shows the weight and dimensional data for the configuration shown in Figure 3. The data shown in Table 4 are for the vehicle as drawn in Figure 3--the fuel tanks of which are sized only very approximately for 5,000 n. mi. cruise range. Vehicle C. G. shift is shown on page 18.
14
)
AERODYNAMIC CtVICES wING + WING ~OUNTED CCNTROL SURFACES VERTICAL SURFACES HORIZCNTAL SURFACES FAIRINGS,SHROUOS AND ASSOCIATED STRUCTURE
80CY STRLCTURE
54327. 8296. 8781.
o.
STRUCTURAL FLEL(OR COMMON BASIC ENCLOSING STKUCTURE PRESSURIlEC CCMPARTMENTS SECONCARY STRUCTURE
PROPELLANT) CONTAINER O. 93143.
5193. o.
INCUCEC ENVIRCNMENTAL PROTECTION COVER PANELS,NON-STRUCTURAL INSULAT lCN
lANDING GEAR
MAIN PROPULSION ENGINES ANC ACCESSORIES AIR I/'lDUCT ION NACELLES,PODS,PVLCNS,SUPPORTS FUEl(CR CCMMCN)CCNTAINERS AND SUPPORTS OXIDIZER CONTAINERS AND SUPPORTS PROPELLANT INSULATION FUEL SYSTEM OXIOIlEK SYSTEM PRESSURIZATICN SYSTEMS LUBRICATING SYSTEM
AEROCVNAMIC ,CONTROLS
PRIME POkER SOURCES ENGINE OR GAS GENERATOR LNITS n,-,,,cn c-n .r.r~ ~ ,.' rr- ." ~ 'r.,.~
POftER CCNVERSION AND DISTRIBUTION ELECTRICAL ~YCRALLIC/PNEUMATIC
GUIDANCE AND NAVIGATION
INSTRUMENTATIGN
CCMMUNICAlION
ENvIRONMENTAL CC~TRCLS EQUIPtJENT PERSCl\N EL CuOl~NT SYSTEM COMPARTMENT INSULATICN
o. o.
42282. 17400.
3719. 20955.
f).
10266. 1536.
o. 4215.
160.
2313.
4063. 1233.
202. 2150.
o. 5274.
TABLE 4. BASELINE DELTA WING CONFIGURATION - WEIGHTS
15
71403.
98336.
o.
18299.
100532.
5847.
3526.
5296.
800.
420.
2025.
7626.
)
PERSONNEL PROvISIONS ACCCMMOCATIO~S fOR PERSONNEL FIXED LIFE SLPPORT FU~NISHINGS AND CARGO HANDLING EMERGENCY EQUIPMENT
CREw STATICN CGNlROlS AND PANELS
DRY STRUCTURE
DESIGN RESERVE
PERSONNEL CREW,GEAR AND ACCESSORIES CREW LIFE SUPPORT
PAYLOAD C.AR GO PASSENGERS
USEFUL lOAD
RESIDUAL PROPElL~NT AND SERVICE ITEMS TANK PRESSURIZATION GASES TRAPPf:C FUEL TRAPPED OXIDIZER SERVICE ITEMS RESIDUALS
RESERVE P~OPElLA~T AND SERVICE ITEMS FUEL-~AIN PRCPULSION OXIDIZER-MAIN PROPULSION POwER SOURCE PROPELLANTS LUBRICANTS
WET STRUCTURE
IN-fLIGt-T LOSSES Fl..EL \lENT OXIDIZER VENl POWER SOURCE PRCPElLANTS LUBRICANTS
MAIN PROPELLANTS FUEL
TAKEOFF,ClIMB,ACCELERATE CRUISE DESCENT LO IT ER LANe
OXIDIZER TAKEOFF,CLIMB,ACCELERATE CRUISE CESCENT LeITER LAND
-0. -0. -0. -0. -0.
-0. -0. -0. -0. -0.
4700. 308.
8116. 305.
1250. 25.
13000. 35000.
230. 1213.
o. 121.
o. o.
173. 83.
I ,
1665. o.
3456. 330.
222000.
o.
13428.
300.
327838. )
o.
1275.
48000.
49275.
1565.
255.
5451.
222000.
TAKEOFF ~EIGHT 606384.)
TABLE 4 (CONT'D). BASELINE DELTA WING CONFIGURATION - WEIGHTS
16
\ I
VOLUMES BODY STRUCTURE CREW ANC PASSENGER CCMPARTMENTS CARGO CCMPARTMENT LAN 0 IN G GEAR fAYS PRCPULSICN BAY WITHIN BODY WING CENTER SECTION FUEL(CR CCMMCN PROPELLANT) OXIDIZER CCNTAINER fUEL(OR CCMMGN PROPELLANT) OTHER BODY VOLUME
TOTAL 8GOY ~OLUME
WETTEC AREAS GROSS BODY
CONTAINER
INSULATION
LOWER SURFACES(TrERMAL PROTECTION) UPPER SURFACES(TrERMAL PROTECTION) PERSCNNEL CO~PARTMENTS CARGO COMPARTMENTS
PLAN AREAS WING(CR LIFTING SURFACE) (GROSS) EXPOSED wING AREA BODY MAXIMUM CROSS SECTION BASE VERTICAL SURFACES HCRIICNTAL SURFACES AIR It\LET CAPTURE AREA
UNIT wEIGHTS WING VERTICAL SURfACES HORIZGNTAL SURFACES BODY STRUCTURE (BASIC) LIFTING SURFACE MAXIMUM LOADING
DIMENSIONAL CATA WING
STRUCTURAL SPAN ROOT CHORD LENGTH THICKNESS RATIO
EODY LENGTI-: WIDTH HEIGHT
ENGINE SCALE FACTOR
CU FT 10999. 8645. 903. 270.
o. 3993.
52386. O.
3290. 18819.
99365.
SQ. FT. 21991.85
o. o.
4831.66 442.80
SQ. FT. 7504.75 4361.79 633.54 633.54
1212.15 1350.86
116.00
L8~9ET 0.24)
6.84 6.50 ~ 80.80
FEET
169.19 143.29
0.04
323.11 24.91 33.29
5.50
CU M 311. 245. 26.
8. o.
113. 1483.
o. 93.
534.
2812.
SQ. fi. 2043.60
o. o.
448.86 41.14
SQ. M. 697.19 405.21
58.86 58.86
112.61 125.49
10.78
KG/SGM 0.33 0.31 0.29 0.19 3.61
METERS
51.7".> 43.67
98.50 7.59
10.15
TABLE 4 (CONTID). BASELINE DELTA WING CONFIGURATION - SIZE DATA
17
)
600
-CI";) 0 ~ . to
,.Q C ..j.>
-So 500 .... ~ to to 0 ~ 0
40v
30o.
~ . p:~+rt+
~h~ .'
::;t;: I±t;:: -....
----H--- ,..-,-,
~ ~.
160
+.
.......
+-1-.. -~'
'=z... ....-.. t.4;" .-l-;-,-,..,-. ~
r-~ .. +-~t;::-t
~~ .f-! -+-~ .. ---+~~
.~t::::;+ -++ .
1 ~TANK NO.
:- ,'+-1 1
. 4 ~-+!-'
~ ~.t_-'-'-'f--!ZERO PASSENGERS - '2~ '" ~==
t<
-=:tt=: . ~.- :~ AND 9.1.RGO
200 PASSENGERS,T 2 . BAC,C,AC,F. 1& CARGO ~ ,
EMPTY
3
170 180 C. G. Feet from Nose
TABLE 4 (CONT'D). BASELINE DELTA WING CONFIGURATION - C.G. SHIFT
18
)
3.0 SYNTHESIS PROGRAM
In order to provide a consistent basis for the vehicle parametric analyses, a computer program that combined performance with vehicle sizing was obviously required. A modified version of the computer program used on previous Aerospaceplane studies was used. The modifications were concerned mainly with:
a) Changing the program to compute take-off weight for a specified payload/range.
b) Simplifying the program to enable it to respond to the broad studies required in Phase I.
3.1 . DESCRIPTION
The program used a mathematical model technique in which the computation of the vehicle weight, volume, aerodynamic and propulsion characteristics were expressed as functional relationships of the vehJcle geometry, sizing and trajectory parameters. Given the appropriate sizing, aerodynamic and propulsion input, the computer program used an iterative procedure to compute the vehicle size required to perform a - ,.. specified mission. Figure~ presents a flow diagram of the synthesis program which was used for thIS study.
The mission is specified by defining, as program input, the altitude-velocity trajectories for ascent and descent, the total range, the loiter time and an initial estimate of cruise range. The configuration input includes the coefficients for use in the equations which determine weights, volume, size, aerodynamic characteristics and propulsion characteristics. For the first iteration, it is necessary to provide initial estimates of fuel weight and engine size, and this allows a determination of the take-off weight and the vehicle size.
For the vehicle thus sized, a take-off computation is made, followed by climb calculations. If at any time during the climb computation, the thrust is less than the drag, the engine size is increased and the computations are started again. Vi.'hen the desired start of cruise condition is reached, the engine is throttled to yield thrust equal drag and the vehicle cruises for the initial estimate of cruise range. A des cent computation is then made by throttling the engine so that thrust is less than drag; this is followed by a subsonic loiter, des cent and landing. At this point the total range (excluding range/during subsonic loiter or cruise) is then compared wHh the total range
19
\.
)
)
INPUT MISSION CHARACTERISTICS CONFIGURA TION CHARACTERISTICS INITIAL FUEL AND ENGINE ESTIMA TES
~,
WEIGHT/VOLUME .... .....
Wlo .... -
+ I INCREASE ENGINE SIZE I I TAKEOFF J~
t I CLIMB: IF THRUST < DRAG
t I CRUISE
L ... r'" • I DESCENT
+ I A nn TC:: 'T' f'" 0 T nc:: t;' 0 A 1I..TI":' t;' I ,J.J.j.JJUU.L '\J.L'\.\.J.&. ..... .LI.1.'\. .,L"U.L..:I
. I T I'"\T""",..,n I • I .LJ '-J~ ~!jn I
• I ADJUST FUEL
I DESCENT
+ I RANGE CHECK
I RANGE::j. RANGEREQ I
dt I LANDING
t I FUEL CHECK : FUEL USED:f FUEL ON BOARD
<tK I PRINT I
FIGURE 7. SYNTHESIS PROGRAM FLOW
20
,h
,
,
1
" )
)
required .. If the range vaiues are not equal, the cruise range is adjusted and the computation is started again at the start of cruise point. When the range values are equal (within a tolerance) the fuel is compared with the initial estimate of fuel. If not equal, a new fuel weight is assumed and the entire calculation is repeated until the fuel used is equal to the fuel on board.
3. 2 COMPUTATION TECHNIQUES
The computer program is built up from subroutines. The sections which follow outline very briefly some of the more important subroutines:
3.2.1 WEIGHT/SIZING SUBROUTINE. The weight/sizing subroutine used in the synthesis program is the "first level" weight/sizing program developed by Convair under contract NAS 2-3025.
The weight-volume subroutine contains the equations for vehicle weight, volume and size. These equations are in the following form:
W ocW -'-W +W +W -+-W -'-W -' TO fuel engines pass surfaces fuselage landing gear
v =V -'-v +V + FUS fuel pass unus
These equ::rtions can be rewritten in the following form s:
where
W
WTO = Wfuel + Wengines + C1 (No, of Pass) -'- C2 (WT~/~)
i +C V +CW +C(W )
3 FUS 4 TO' 5 TO -
C1
, k1
, 1, J, W ,W etc., are fixed at any sizing condition. fuel engine s
Due to the nonlinear nature of the weight and volume relations, it is necessary to use an iterative procedure to solve for take-off weight and size. The take-off wing loading is specified as input.
The \veightisizing subroutine also includes geometric relationships that result in \veight and aerodynamic scaling effects, e. g. ,
21
~\
I
)
)
d 1 .667
Vertical tail area = C. (bo y vo tune) . 1
Exposed wing area = Gross Area - C, (root chord x body width) . J
where C. and C, are constants determined from the baseline configurations. 1 J
3.2.2 AERODYNAMIC SUBROUTINE
The aerodynamic subroutine computes lift, drag, and angle of attack characteristics. The inputs to this subroutine are the aerodynamic coefficients; C
L ,CD < CD'
Ci wave, f etc., of the various configuration· components; body, wing, horizontal tail, etc., based upon the component reference area. Interference effects are included in component characteristics. The configuration characteristics are obtained by converting the coefficients to a reference area equal to the vehicle wing area and summing them up. For example, the following equation is used for configuration lift curve slope:
S Sb Sh w C
L =C
exp +C )I('
cross +C ... tail
)(
L S L S L S Ci
tot Ci w Ci w a
tail w
wing body
where CL
Ci • wmg
CLare component lift curve slopes
Citail .
s w
exp
Sb cross
= wing cAllosed area
= maximum body cross sectional area
= horizontal tail area
In the. performance computations the required lift coefficient is obtained from: V2
W cos y [ 1 - - ] - Thrust sin a . Rg
CL -'--------------------------req q S
w
.An iterative procedure is :used to determine the angle of attack which satisfies the above equation.
22
)
)
The total drag coefficient is given by
"Yhere
C =C +C +C D D Dfo 0 D
o rIctlOn 1
CD = pressure drag @ Cl = 0
o
CD1
= drag due to lift
3.2.3 PROPULSION SUBROUTINE. The propulsion subroutine computes the thrust and propellant flow characteristics of the propulsion system being considered. The details of the propulsion routine varied for each propulsion system. In general the propulsion subroutines were divided into a turbine engine routine and a ramjet routine. The propulsion data used as input was non-dimensionalized as discussed in Section 5. O. In general, the following form of equations are used:
a) TU REINE ENGINE . Fuel Flow Parameter = _ rr;:;--
PTvTT
w F
2 2
= f (T ) = FFP T2
where P T = inlet total pressure 2
T T 2 = turbine inlet temperature
P T 2 and T T 2 are functions of such parameters as Mach number, altitude, pres
sure recovery, etc.
Gross Thrust Parameter = TG = f (fuel flow parameter) = TGP
PT2
Ae
where Ae = nozzle exit area
~aF Air Flow Parameter = 2 = f (fuel flow parameter) = AFP
P T2
The net thrust is given by
V T = TGP x P x A - A x P - ~
T 2 e e atmos a g
23
') b) RAMJET
)
The ramjet mass flow ratio, W = F (Mach number ahead of inlet) n
Ramjet Thrust Coefficient = C t = f (Mach number and fuel/air ratio)
The net thrust is given by:
\,. T = PFC x w x C x q x ACAPT n t
where PFC = pressure field effect = f (Mach angle of attack) q = dynamic pressure
ACAPT = ramjet inlet area
Specific fuel consumption - TSFC = f (Mach, and fuel/air ratiO)
Ramjet throttling is accomplished by adjusting fuel/air ratio.
3.2.4 PERFORMANCE SUBROUTINES. The performance subroutines of take-off, climb, cruise, descent, loiter and landing use conventional aircraft performance computation techniques. The climb and descent computations assume straight line climbs between two points on a Mach number/ altitude trajectory. The performance computations require as inpu1; the vehicle weight, thrust, fuel flow, lift, drag, and angle of attack. In addition to these standard performance computations there is a subroutine which computes at a given Mach number the altitude to maximize lift to
drag/SFC. This routine is used to establish the start of cruise altitude.
24
,
)
)
4.0 TRAJECTORY EVALUATION
The first parametric study conducted in Phase I was a trajectory evaluation study. It was the objective of this study to detennine a near optimum trajectory and also to determine the performance sensitivity to trajectory variations. Figure 8 presents a typical mission trajectory and indicates the critical areas which were investigated. These areas are as follows:
1) Sonic Boom
2) Dynamic Pressure
3) Inlet Pressure
4) Cruise Altitude
5) Descent Trajectory
6) Subsonic Loiter Cruise .Mtitude and Mach Number
4.1 TRAJECTORY EVALUATION GROUND RULES
The following ground rules were applied in the trajectory evaluation study:
a) Baseline Delta Wing Configuration. This configuration was utilized since its characteristics were available at the time. It is believed that the trajectory evaluation results are essentially independent of configuration.
b) P&W STF-230 Fuel Rich Turbofan Ramjet. This propulsion system was utilized since it was typical of the airbreathing propulsion systems under consideration. It is believed that the results, at least the trends. obtained with one type of airbreathing propulsion would be applicable to other airbreathing propulsion systems. In computing the propulsion perfonnance, the engine was considered as two separate units, a turbojet and a ramjet, each with its own sizing requirements. The number of basic turbojets == ACT where the basic turbojet size is 550 lb/sec SLS. Computations were not restricted to an integer number of turbojets. The ramjet capture (cowl) area == ACAPT.
25
110
100
90
80
70
60
--;::: 0 50 0 0 ...... --
) (l) '0 ::s ;::: .... 40 <
30
20
10
o ~~!r;m~:~~~;-=:;~:~~~4m~~i~~~1r~~~~~ o 1 2 3 4 5 6
Mach No.
FIGURE S. TYPICAL TRAJECTORY
26
)
c) Cruise Mach Number = 6.0
d) Mission Range co' 5,000 n. mi.
e) Subsonic Cruise Range =-0 10 n. mi. ,Subsonic Loiter 100 sec.
f) 200 Passengers @ 175 lbs. per passenger
g) 5,000 lb. Cargo..,. 40 lb. Baggage/Passenger
h) Synthesis program \vith a preliminarj weight/sizing subroutine was used. This program was used for the entire trajectory evaluation study so that the results are consistent; however, the take-off weight characteristics obtained in the trajectory evaluation study should not be compared directly with the results of the othe'r evaluation studies; e. g., propulsion evaluation.
i) The trajectories which were considered are shown in Figure 9. 'The sonic boom levels shown on this figure are nominal values based on a preliminary sonic boom analysis. The descent trajectories are approximately constant dynamic pressure trajectories. The break in the descent curves at Mach 5.5 and 100,000 ft. is not significant. it came about since the descent trajectories were analyzed prior to completing the evaluation of cruise altitude.
4.2 somc BOOM V ARlATIONS
The effect of sonic boom level on take~off \;veight characteristics are sho\vn in Figure 10 for two turbofan engine sizes. The knee of the curves occur at an overpressure of
approximately 3.0 psf. Below 3.0 psf the take-off weight increases quite rapidly as the altitude is increased to reduce the boom overpressure. It is also necessary to increase the engine size as the altitude is increased: For lower trajectories, sonic boom overpressures greater than 3.0 psf,the increased engine perfonnance with a larger engine is more than offset by the increase in engine weight such that a smaller engine yields lower take-off weights.
A transonic and low supersonic trajectory having a nominal sonic boom o~:erpressure of 3.0 psf was selected. The detaile? sonic boom analYSiS, discussed in Section 12.0 indicates that the sonfcboom level associated with this nominal trajectory is somewhat less than 3.0 psf for most of the configurations.
4.3 DYNAMIC PRESSURE VARlATI01\18
Figure 1~ presents the effect of dynamic pressure on take-off weight indicating that there is essentially no effect on take-off weight for maximum dynamic pressure between 1500 and 2500 psf. (The weight analysis techniques used did not account for dynamic
pressure as a primary factor effecting structure weight). These dynamic pressure
27
110
100
90
80
70
60
....> ~ 31,)
3 0 0 0 ......
)~ .to - .~
...., <
30
20
10
n
)
t- . •.. ~ ~.- ~t - ,. 1- : .. ! """.-"
:=i;T-'F-'~'Il:' Son-ic "boom-overpressure 5.0 psf +---'---"-...:....:.c:...:.j-i-'F--,-:--- ~.~.·~i~-£~d~~~=;l~:$~f:2 Sonic boom overpressure 4.0 psf
o
. i . • . • . -----..... --~----...... ,
1.0 2.0 3.0 4.0
:vlach Number
Sonic boom overpressure 3.0 psf Sonic boom overpressure 2.5 psf
Dynamic pressure 3,000 psf Dynamic pressure 2,500 psf Dynamic pressure 2,000 psf Dynamic pressure 1,500 psf
330 psi
5.0 6.0
FIGCRE B. TRAJECTORY VARL!\ TIONS
2S
)
Take off weight
600,000
580,000
Ib 560,000
540,000
520,000
, " .:. , ; :::··t·: .. +·"
2.0 3.0
.... ~ ... , ...... -i . . . ; : . ; '~
Sonic boom overpressure ~ psf
ACT '·(No. of Tu r bo jet sat
7.0 SLS ai rflolA'= 550 Ib/~ec. )
6. °
4.0
FIGURE 10. EFFECT OF SOKIC BOOi\I 01\ TAKE-OF f WEIGHT
Take off weight
Ib
560,000
550, 000
-. ,~ i - ~ ~ l . •
.~~~ ~:~~~f~i:.:-.~: ~+:~ ::!·.~.~.;:~;.~.t. : .. :'.-;.: :':~~'!":-' -j':. ,-': .. - : ., ~ .• - ~. - - . ~ .-. I
•• ~ '" .•. t -~ • _ •• I .•.. ,
• :',: ::.--:~ :--:-1 .- : : ~ . : :: +' :., :::::'.:.:tc::-=.:::.:.r;..· ....... ~ __ ........................ ....., ___ .............. l_.· ...
-_ ....
1000 20n 0 ~~ooo
D:'-'113mic prl'ssure pSl
FIGGRE 11. EFFECT OF DY:·:.\.i\IIC pnESSU~E <.,);\ T:\KE~OFF WEIGHT
)
)
variations were obtained by operating over constant dynamic pressure trajectories bet\veen a sonic boom overpressure trajectory of 3. a psf and an inlet diffuser pressure trajectory of about 220 psi. From these data it is apparent that sonic boom overpressure considerations have a much larger effect on vehicle size than dynamic pressure. Since performance characteristics cannot be disassociated from propulsion size, the effects of variations in engine size, ACT, and ramjet capture area, ACAPT, were investigated.
The effect of ramjet capture area on take-off weight is presented in Figure 12. The optimum ramjet capture area is approximately 130 ft2. The effects of turbofan engine size on the variation of take-off weight with capture area are also shown. These data indicate that the optimum ramjet capture area is relatively insensitive to turbofan engine size, however increasing the turbofan from 6.0 to 7.0 increases the tak~-off weight at the optimum ramjet capture area.
It should be noted that at a later date in the Phase I studies, a more thorough investigation ofpropulsion size effects was made. This is discussed in Section 5.4.
4.4 INLET PRESSURE VARIATIONS
Figure 13 shows the effect of maximum diffuser inlet pressure on take-off weight characteristics. These pressures were achieved by operating over trajectories having these values without resorting to builtin reductions in pressure recovery. The maximum inlet pressure trajectories were used from a dynamic pressure trajectory of 2000 psf up to Mach 6. O. The corresponding inlet unit weight characteristics are also shown in Figure 13. The results indicate that the increase in inlet weight at the higher inlet pressures more than offsets the gain in propulsion performance.
4.5 CRUISE ALTITUDE VARIATIONS
In determining the best altitude for cruise, the least cruise fuel will be used when V. L/D/SFC is a maximum. Maximum values of LID can be achieved by flying at the altitude that r~sults in L/DMAX ' which is dependant on wing loading. Minimum values of SFC are dictated by the throttling characteristics of the ramjet as sized for acceleration including the constraints of engine cooling. The best cruise altitude is therefore a balance between wing loading, ramjet sizing, acceleration trajectory, ramjet throttling characteristics and engine cooling requirements.
The effect of cruise altitude was studied for a wing loading at takeoff of 90 Ibl sq. ft., an end of acceleration altitude of 91,300 ft., and for the engine cooling requirements shown in Figure 38. From Figure 14 it is seen that minimum ascent + cruise fuel occurs at a cruise altitude of about 110,000 ft. which is also the altitude for LID, X and minimum SFC. Figure 14 also shows that increasing the ramjet size to [gtf sq. ft. increases, slightly. the best cruise altitude and reduces the acceleration plus cruise fueL Figure 15 shows the effect of cruise altitude on takeoff weight for several ramjet capture areas. From these data it can be seen that the lowest takeoff weight occurs between 110,000 and 115,000 ft. , the lower altitude being associated with the smallest capture area. The minimum takeoff weight 'for the various capture
" .
30
)
)
Take off weight
Ib
570,000
560,000
550,000
540,000
530, 000
120 130 140
t,p = 3.0 psf
q = 2000 psf
150 160
Ramjet capture area - sq ft
170 180
FIGURE 12. EFFECT OF RAMJ~T SIZE AND TURBOJET SIZE ON TAKE-OFF WEIGHT
31
w I:-,J
'''---
Takeoff Weight
lb.
~
tr:-- .. .cc.~ -: ~7:7f7T;t;7-:r:-:-~·t:ili:~r ;,;,:,1,[, ., .. ,·Jttfrf1ti' UtT t..dill. ,,,, :;;~ : :: :;:: :~: :::: ~!!: 7::; ;;:, :;' .:£; ::i: ;:;;;~~ ::;,[@ ,. 't;-u~ ':~;:,~: ;;;; 520,000 :::,S '" ,,' ';,: :-:1.: :.I:: ::;:p,-:;::c:: ';~: ::::ri:;; ::r: 1: i1 iftf+ ffP~i.: .l
i: ::'
500,000 I~-~~:' ~i ~s!f~:~ ~um ~j;! ". 'J. .,j, ".,,, '" '''''' 163 " ",' ,," 203 1?2, ,,1 "; ;:; , 'f: ::;1"'1'::: ' . .:: ~~~j~:;-'.: ~:~: ; ::: :~:~ :~;l~::' ~.;:"j-" .. -1-.-, :_'::"" ''''''r'' ::+.--. ;: l:: .:,: '::; ,;":; ';; ;::,:::1:::
480 000 t:X::" ·.L ,,'.1: "1=: ::J,;:' : '" " : ': ";, ' .. " ", '" , I
100 200 300
Inlet Pressure (psi)
ACT = G. 0
ACAPT' l:lO sq. ft.
Inlet Weight Ib/ sq. ft. of capture area
FIGUH~ 13. INLET PHESSURE VS TAKE-OFF WEIGHT
" )
)
1.0
1.0
Relative sfc 1. 0
.9
.9
.S ...., :":! ... 00 00 t":j
S ill ttl ...... ::l ... 0 + ...., s:: ill 0 00
<
::i
90,000
1.6
1.5
1.4 90,000
100,000 Cruise Altitude ~ Ft.
110,000
) Cruise Altitude ~ Ft.
FIG 14. EFFECT OF CRUISE ALTITUDE h.~D RAMJET SIZE ON MASS SIZE RA TIO
33
LID
sfc
I ')
)
1.0
.9 Relative
sfc .96
.9
Take off
weight 10
5.0~~~~~~~~~~~~~~~~~~~
= 150 FT2
~~m ~lli1ill! ;1~~1 ~J::~~~ ;:lm:j ·:JT1d::n -8;';; m~mi .:*1 4.4 I
100 110 120 Cruise Altitude (1000 ft)
620,000
600,000
580,000
560,000
540,000
520,000
500,000
Cruise altitude (1000 ft)
Figure 15. Effect of Cruise Altitude on Take-off Weight
34
)
area/altitude combinations occurs at a constant fuel/air ratio of about. 023 (<P== • 8). Figure 15 also shows that the heavier inlet associated with increased inlet size more than offsets the reduced acceleration fuel weight shown in Figure 14. The results shown in Figures 14 and 15 were obtained for a constant cruise altitude. Operation at a constant L/D/SFC would give Borne slight variations in these results but not change the trends. It is concluded that the lowest takeoff weight will be achieved when cruiSing at the best value of the range parameter V. LjD/SFC and that engine oversizing specifically to increase cruise altitude is not desirable. The synthesis program __ was then modified so that the altitude for cruise would be at the maximum value of L/D/ 'SFC. . -..-,
4. 6 DESCENT TRAJECTORY
The descent from the end of cruise is accomplished by throttling the engines so that thrust is less than drag. The performance is computed along the several altitudeMach number trajectories shown previously in Figure 9.
The ramjet thrust during descent is given by:
T = CKTH x maximum ramjet thrust
The ramjet SFC during descent is constrained by cooling considerations.
The turbojet thrust during descent is given by:
T = CKD x drag
The effects of these throttling parameters on take-off weight are presented in Figure 16. On the basis of these data, throttling constants of .2 were selected for both the turbine engine and the ramjet.
The performance was analyzed along the descent trajectories sho\';'l1 in Figure 9. The effect of these trajectories on overall gross weight is shown on Figure 17 which indicates that the optimum performance is obtained for approximately the highest altitude descent trajectory shown in Figure 9 which is 910se to maximum lift/drag ratio.
4.7 SUBSONIC LOITER AND CRlJISE VARIATIONS
The subsonic loiter and cruise trajectory variations were investigated by determining the effect of Mach number and altitude on the range parameter V. L/D/SFC and the fuel flow. Figure 18 shows the effect of these parameters on the range parameter and Figure 19 shows the effect of these parameters on fuel flow.
Figure' 18 indicates that the optimum subsonic cruise condition occurs at approximately Mach. 9 at about 45,000 feet. Decreasing the Mach number decreases
35
)
)
Take off
weight Ib
Figure 16. Effect of turbojet and ramjet throttling on take off weight
Take off weight
Ib
Figure 17. Effect of Descent trajectory on take off weight
36
)
) /
V. LID SFC
6,000
4,000
2,000
o o 10 20 30 40 50
Altitude (1000 ft)
Figure 18. Effect of Subsonic Loiter Altitude & Mach-No. on Range Parameter
40
Fuel Flow 20 lb/sec
o o 10 20 30 40 50
Altitude (1000 ft)
Figure 19. Effect of Subsonic Loiter Altitude & Mach No. on Fuel Flow
37
)
the maximum range parameter and also decreases the altitude at which the maximum occurs. The relatively poor subsonic cruise performance is, in part, a result of the fact that the SFC increases when the fuel rich turbofan is throttled. The upper altitude end point on each Mach number curve is approximately the point where maximum thrust equals drag; i. e., the engine is not throttled.
Figure 19 indicates that the minimum fuel flow, which is the parameter of interest for loiter time,occurs at the higher altitude and is not very sensitive to Mach number. Essentially the same minimum fuel flow can be achieved at various Mach number-altitude combinations.
The aerodynamic data used in investigating these subsonic loiter and cruise variations had a drag divergence Mach number of .9. This was merely a preliminary estimate and was not based on detailed analysis. The effect of drag divergence Mach number on overall performance is insignificant, however the effect on the optimum Mach number for subsonic cruise would be noticeable. It W) uld change the optimum Mach number, altitude for cruise and change the range parameter somewhat. In Phase II the drag divergence characteristics will be better defined and the optimum subsonic cruis~ conditions established for each configuration.
4.8 SELECTED TRAJECTORY
On the basis of the trajectory variations discussed above a nominal trajectory was selected and is shown in Figure 20.
A transonic and low supersonic trajectory having a sonic boom overpressure level of 3. 0 psf was selected. This overpressure level is obtamed as defmed m the section above on sonic boom variations and is a nominal value based on preliminary calculations. It is expected that the more detailed sonic boom calculations will show that this trajectory has a lower sonic boom level than 3.0 psf.
Between Mach 3.5 and 4.0 the trajectory follows a 2000 psf dynamic pressure line.
Between Mach 4.0 and 6.0 the trajectory follows a 130 psi inlet pressure line.
A constant Mach nu..."'TIber climb is followed to the start of cruise condition. The start of cruise is shown at 110,000 feet on Figure 20, however this point is varied
--~~--~~----------somewhat to achieve the best value of the range parameter for each run. ----_. The initial sharp descent from Mach 6.0 to Mach 5.5 is not significant. It
merely re~;mlts from the fact that when studying cruise altitude variations it was necessary to have a descent trajectory below 100,000 feet. For cruise altitudes above 100,000 feet it was assumed that the trajectory descended from Mach 6.0 to
) the nominal descent trajectory at Mach 5.5.
38
)
)
The subsonic loiter and cruise are assumed to occur at Mach.9 at 40,000 feet.
This selected trajectory is not necessarily the optimum, however it is near optimum and is a representative trajectory for an air breathing propulsion system. As such this trajectory was used for the propulsion system comparisons and for the configuration variations. Additional trajectory variations were made during the propulsion evaluation study and during the cruise Mach number study. as discussed in Section 5.0.
39
100
90
80
70
60
E-< ~
') 0 0 50 0
>0:; Q
_I ;:J E-< SUB- ------+----. - --t-- - --->-< , E-< 40 SONIC ....:1
i
30 -----. --·i··~
;
.. - ,-
20
.. -.. --:.:T·--.~ .. - L ~--~_+~4----+--~-+_~_+i--~~I--~~--~~--~--~--------------
_l.~_~~~~_~_._ .. : _._~_~i _"~:Li·. __ ~i ... I I·" i .. ' . f . --~-'-----r-- --
: :
10 1Tt- ++1c -j--J-:: ..... , ..... i ... o
o 1 2 6
}Iach Xo.
) Figure 20. Selected Trajectory
40
5.0 PROPULSION EVALUATION CRUISE MACH < 8.0
See Report GD/C-DCB-66-004/2A.
)
) J
41
)
)
)
6.0 DELTA WING CONFIGURATION VARIATIONS
The effects of variation in the following parameters were investigated for the delta wing configu:ration:
Wing aspect ratio (70 degree leading edge sweep) Wing sweepback Wing loading Wing thickness ratio Body fineness ratio
6.1 GROUND RULES
The following ground rules were used. These ground rules were also applied to the configuration variations considered for the Variable Sweep configuration (Section 7.0) and the Blended Body configuration (Section 8.0).
1) Pratt and Whitney STF-230A Fuel Rich Turbofan Ramjet
2) Cruise Mach Number = 6.0
3) 5,000 n. mi. range
4) 200 passengers
5) 5,000 Ibs of cargo plus 40 Ibs of baggage per passenger
6) Selected Trajectory - Figure 20
7) Subsonic Loiter Time = 1,000 seconds and Subsonic Range = 100 n. m .
6.2 WING GEOMETRY VARIA TIONS
The baseline delta configuration had the following wing character i sties:
Leading Edge Sweep = 70 0
Wing Aspect Ratio = 1.45 Taper Ratio = 0 Wing Loading = 90 PSF Wing Thickness Ratio = .04
81
)
The variations considered were sweepbacks from 60° to 75 0, aspect ratio vari
ations from. 96 to 1.65 for a taper ratio of .2, wing loadings from 65 to 110 psf and wing thickness ratios from .025 to .07.
The effects of wing sweepback on take off weight are shown in Figure 54. The take-off weight is not too sensitive to changes in sweepback. The optimum sweepback is approximate ly 70 degrees.
The effects of wing thickness ratio on take-off weight are shown in Figure 55, The optimum thickness ratio is about. 06. The effects of thickness ratio on unit , wing weight and fuel weight as a percentage of take-off weight are also shown in Figure 55. Fat thickness ratios below . 06, the improved aerodynamics which yields the lower fuel weight percentage are offset by the increase in unit wing weight. For thickness ratios greater than. 06, the reverse is true.
The effects of wing aspect ratio on take-off weight are shown in F'gure 56. Increasing the aspect ratio increases the take-off weight. Also shown in Figure 5() is the effect of aspect ratio on fuel weight as a percentage of take-off weight and unit wing weight. From these curves it is apparent that the increase in wing weight with increasing aspect ratio offsets the improved aerodynamic characteristics. The unit wing weight increases due to the increase in structural span.
The effects of wing loading on take-off weight for various aspect ratios and sweepbacks are shownin Figure 57. Increasing the wing loading decreases the take-off weight due to the reduction in wing weight. In order to select a nominal wing loading, it is necessary to consider take-off and landing requirements.
The study ground rules shown in Table 2, specify the take-off velocity as 160 kts. and the landing approach velocity as 135 ~s. The mass ratios are such that take-off ~- the cti1ical speed. Take-off distance is not critical due to the high take-off thrust ~ ~ to weight characteristics. It was assumed that the take-off angle of attack was 15°. ~ 1/\ High lift devices were assumed yielding the following take-off lift coefficient characteristics:
Configuration
L. E. Sweep = 70° L.E. Sweep = 60 0
Aspect Ratio = • 96 Aspect Ratio 1.21 Aspect Ratio 1. 65
"-
0 0
.2
.2
.2
Take-off CL
1. 12 --1.21~
.99 1. 07 1.16
A ten percent velocity margin was assumed; that is, it was assumed that lift equalled weigh~ at 145 kts.,_rather than at the specified take-off velocitYOr 1601«s.
-------------------- - - --- -------------------------- ------------------
82
:0-.... 0 0 0 ...... ...., ..::: bD ..... Q)
~ ...... ...... 0 I
Q) ..:.:: cd
E-<
)
)
600
550
500 .::~ j . ~-: ~-
60
Taper Ratio, A = 0
65 70 75
Leading Edge Sweep Back Angle (Deg.)
Figure 54. Take-off Weight vs Leading Edge Sweep
83
! --"t-
: : t :: .. _. I .
. I I
80
00 ~
'-..---~
t-rj
~. "1 CD CJl CJl
..., ~ CD I o ~
~ .... ~ ..... < rn
~ ...... CJ ~
~ Ul rn !l:1 ~ ..... o
;l ..... n ?'r ;::l CD (fJ
Ul
~ \l) M ..... o
~ CJl
Fuel Weight Take-off Weight
~ 0':>
~ -:]
~ 00
~ ..... n
B (l) Ul [JJ
~ M ..... o
"----I
Unit Wing Weight ~ psf
~ 0':> 00 ..... o .....
t9
IT' I j l' r! !'~Il I Itl WI i' it nn ~ ; ; I ~ 1 ! t· J l l' 1 " -I t-· Itt- I :" t ,.! I ~ 1 i ~ . ~ t- -i I
'! 1 ...:..~ f-Ll:.. I_:":'~ U·.' j ;..;..;.'" I, " , jT·: I. :I-~--:-, ,t-'--;--r~::, t,' :.
I ,d :<:: ; Ll: :: -. : ~: ; ~'i ; I t ; Ii ; i ; ; :: :::' ,:1; ':;1 dlit: It ~,,: ~'tl ,I., .. ,).. ·t~· ;!-t~ t ~ " ,. , , •• , ' " j H, .,., ~'~f;
': I:, r I: .. :;:' t.l: ;1 :, . i . [I "'[ "I J'I ,."., 'j , iiI 1'" ~ t I' 1 t ~ ~"'! .• -1 H- t. 1 . .,.j-' . f • t l 1 I \ t· 1- t- .-:. t t-, I : ~t t ~ r r' . L... t' , II! t r It! !" ,. f; , t r t-I ~. 1 '
r~ H-t:--, , "" 'I" "C' 'I j. ~i" 'I' , I ; I '1'1 ,·tI, ;, \' ; t Et , I;, i·r' , . q, , ~ Lh trll it ; it t;c\j ,tit t:~~ ~llt lljt(tH HI! illtH,1I li'rr :1;: "~r; ~! ; .ott;"l 'Pi .:-1; q/t~Hft l:li
(Itt til' d~llf1!t:t:'li ~i;; tt[!;"); r:lt, 'f ;f,.t;",. r'i" 't't ~ It ,. ~ i ,! • j I I . " 1 -t •• J ~ r r -. \ - , - -, I
If II rrt tit; :q' :;1 rt-;;lli: j· t tn. ! II ;'1 :ilt I:,:;: 'fill Ii !ji:'
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j Itt ., , f" .. I- - ,.!. t-. I t- t I-t
t I' ~- ~:: ~ r I ~ i ;-t; i :f j; .j l t: : ~ ;. t :": I 1 I' + t I I ~ t-i l " ,-., -.j. \ .. .,. •• F-'t.l 1::l-t- ii" ",' , I [ t ~'I : I, ., t .
,.-ntti', t I,· ., j' ", • , . " , It· '! I ~ tti (+ ,;. , ,1_, I' r" 'r I , I,.! "! j .j~ I h I I t J·t' I'" t .. ~r t r" I ',' '" I ~ t '~ tTl '. t' .. r I!' 1 t I I ,
I .! .. \ r )! • I • . , , : 1 ~ -\ - " ~ ." I; j Ii!;' : i : i ;: i i ;! i f iii ;:::' ,. m+t~Tr lTi; '~;'fJ ' : ~'.: . : t :
, I" ":, ," I • I ' , , I • , ~ , I , t I, I,; f ' - , .. ~ ". - I + ' •
i ; : ~ . ~.: :; : L:; , : n ~ i .; H ~ ~ ~ i ~ ! ~!. ~:rt--:·~ -:-j; ~, ., 1" "" Tj:-:-
I • '.' r.' t! I" \ I. . j. ,. t: j ~ '", !" i . I j • ! t-! ~ t ! i i ! j't ! ~ . i: : :,1! :. t ! ! ." ,t ~ J
"It· ,II ' ,.'" t'I' ... ; 'I' '. '. .- !--I- .,.; T ,., '" t ~ I f I ' ~ I ~. I ~ I t . \. til 1 ~ ., •• l' : 't ... I •. T ' ,
f':";~~ ::;1 ~ '-.:t; :..;;- ':-:-:-:- ----:":1 ~:-: : , t , 'I 1'" \", . , "It .. I, "" ., .. ! ~.:-: I ~ r' ; ~~: :.: - '~! 1 t ~;; i_·~;
; n! :i: :\.;: ;-~~~.: :t~: '''''- ... I ---. c
..., ::T .... CJ
8" (1) Ul Ul
~ ~ .... 0
o t-.:l
o ~
. 0 ~
0 <:Jl
o 0':>
o -1
J
CJl CJl o
Tak~-off Weight (1000 Ib)
0':> o o
0) CJl o
[PI lJ! t ·tl' ftit nrn n:llH : ;~ q t :!! : I: : ~ U~; tit H 1 J I t 1 itt' ~l t i " "'<I' ''''n'f,' ct,· ,t ~l IfHr Ii i;H ;tHHhi t·· It! ~r t; ftrltlrl; .. lL!1!tHt-t-""'tH If IltHt: I' fil ;\!-' :1 tnt LfEtl!f, ·1' .. I· ._p- +ft j tt· +-
1 .1, r ~. f-ti t f t t' :t +8 " .
rHB .~. 1 ~T , IH IL -'
r f- tEfH! It. rift! f l~ ~#JtV fH- , . .t '
i H:t~ IIHtt±:tt '.1. HI . ' ii r1-ft~tn+f +r
l· I il t 'ttt ~~'H '('I' f ,:1 r . Id~l lhl tnf Ii It (r r
/ir\ fIt; HlfltH t'-Ibl t ;'I'·i:W t. j ~n it' ~Ur' ~} :f~f.·~ t .-
ill H' !ll L Ij u!fil . PI ~ [rlr! Hit nH) I t 1"
t t' I ' I r I 'I iU! '" 1 'j I, 1: tH.:I!:! fill\ itt ttl I'Ll ;~;I: + \ t1 Iii , d I ! I '1, I·d [II it! It t r ~ ~l
~f
)
...., ..t:: ..... Q)
~ -Q) ;::;s ~
o o o .....
~
650~--~~~--~~--r--r~--~-~~~-~-----~--~--~~~-----~--~~ "r : :\.::.: .~:. :;::: :.' !i;: :':l,.:::i
j::;: : ::1>:::1\: : .:>I:::.::1:::::i: :: : : t~: :: :::: '.:: :::: :::: :..: .. .. ...... .. I' . . . . .: I .. ' .:.1Tt.... ... I .'
fn 600 .... ~
~
..t:: bD ..... Q)
~ ""' ""' 0
I Q)
~ E-;
.8
v'-t-j 6 f::~r::: r
.38
. 37
.36
.35
1.0 1.2
Aspect Ratio
;
10 A speet Ratio
Aspect Ratio
1.4 1.6 1.8
.; :1 : 1.:'.
:':1' :' • • ~ 4
.. I· .
::':l""-~:'. '.:' ... I:::' :·j··I:::~~··l
.' j: .. I ..... , "j"::; . i:j - . :...:' f..::::
J
Figure 56. Take-off Weight vs Aspect Ratio
85
00 0')
Take-off Weight (1000 lbs)
800
700
600
+
500 60
~-
\---./ ~
,,',:T Take-off Speed ;:: 160 knots
- ~ 1 -'U q - ~ It 11 I l:t tt
ft It -, ,t III f ~ + ~r ~ t
If J . I t ~ , 1-
j '. till
. 1 ~t
It t
1.65 } AR (A ;:: 70°)
1.21 .Jl.LE
W[fr 1 + t tt i , J t .+ +
60 0 (A. ;:: .20)
70° }ALE (A. ;:: 0)
:I fi • j
7( 80 90 100 110 120
Take-off Wing Loading (PSF)
Figu e 1 . Take-off Weight VB Wing Loading
- - - --------- - - - -- -- --- ---
)
)
With these data, it was possible to determine the maximum values of wing loading which would meet the take-off requirement. The resulting limit lines are superimposed on Figure 57. The best configuration is the 70 degree delta wing with a wing loading of 81 psf.
6.3 BODY GEOMETRY VARIATIONS
The effect on take-off weight of body fineness ratio variations was investigated over the range of fineness ratios from 7.0 to 16. O. The baseline configuration had a body fineness ratio of 11. 5. The results are shown in Figure 58. This data indicates that the optimum fineness ratio is between 12.0 and 14. O. As the fineness ratio is decreased, the higher drag offsets the gain in structural efficiency.
6.4 INTEGRAL TANKS
USing the baseline delta wing configuration shown in Figure 3 "as a basis, the effect of changing from the five circular, non-integral tanks to integral tankage was evaluated. For Mach 6.0 cruise, the body structural concept for the non-integral tank arrangement was "hot" nickel alloy primary structure with insulated, nickel alloy tanks. The body structural concept for the integral tank arrangement was a titanium, primary load carrying structure with external insulation and nickel alloy cover panels. The integral tanks had a tank liner.
Table 9 shows the Significant differences between integral and non-integral tankage. Table 9 shows that integral tanks give a lower take-off weight. This results mainly from the spiralling effects of the smaller body volume associated with the better volume utilization of the integral tankage and from slightly lower unit body
weights for titanium construction.
6.5 FINAL DELTA WING CONFIGURATION
Based upon the configuration variations discussed above, the final delta wing configuration has the following characteristics:
Wing Loading Aspect Ratio Taper Ratio Leading Edge Sweep Thickness Ratio Body Fineness Ratio Integral Tanks
81 PSF
1.45 o 70° .06 12
The synthesis printout which includes the performance and weight/sizing characteristics is shown in Table 10. These data include all of the updating throughout Phase I and can be compared with the final data for the variable sweep, blended body and
scramjet configurations.
87
-----\ I
)
)
:0-..... 0 0 0 ......
..., .c OIl ..... ~
~ ..... ..... 0 I ~
..!<: ell ~
800 ~c:-;::-~ --- h"-7 ~ 1---
700
600
500
400 6.0 8.0 10.0 12.0 14.0 16.0
Body Fineness Ratio
Figure 58. Take-off Weight vs Body Fineness Ratio
88
I . Non-Integral Integral
Tankage Tankage ,
WEIGHTS
Aerodynamic Devices 60,079 53,593
Body 99,697 77,981 i / i
Induced Environmental Prot. 0 31,923 i I I
Main Propuls ion (Total) 105,988 67,490
Fuel Tanks 20,406 0 \
Fuel Tank Insulation 10,3;33 ° i
Fuel (Total) 218,868 196,646 I
Acceleration 77,732 69,545
Cruise 104,837 94,032 \
Descent 15,802 14,175
) Loiter 17,597 15,964
Landing 2,797 2,822
Take-off Weight 598,447 537,040
VOLUMES ,.
Body Structure 11,236 9,931 I
Fuel 51,661 46,416'
Unusable 19,072 12,513' I
. \'{ . _15'" Total Body Volume 100,381 83,418
DRAG
@M= 1.4 160,027 142,374
@ M = 6.0 (91,300 ft) 112,972 100,864
) Table 9. Data for Integral vs Non-Integral Tanks
89
-~ )'fFnr.VNA~'C DFVJCF~ . ~JNG + WING ~nUNTEO CONTROL SURfACES
\lFRTTCAI SlJRFACFS HOR 110NTAI ~URfACFS
FAIRINGS.SHROllDS AND ASSOCIATED STRUCTURE
fr.tv STRUCTlJRF
38228. 7451. 7Q14 .•
o.
STRUCTURAL FUEL{ OR COMMON PASTC ENClCSTNG STRUCTURE PRFSSlJRI7FC COMPARTMENTS SFCONDARV STRUCTURE
PROPELLANT) CONTAINER O.
fNrUCfD FNV'RO~~FNTAL PROTECTION COVFR PANFLS.NON-STRUCTURAL J,..SLJI ATtr.N
lA~DIf\G GFAR
"AJN PROPlJtSJCN fNGTNFS AND ACCFSSORIES AIR J NOIJCT leN .-~ACFtl FS.POCS.PVLONS.SUPPORTS fUFI {OR r.OMMON)CONTAINFRS AND SUPPORTS r,XIOT7FR CrNTAJNERS AND SUPPORTS PRnPFl1 ANT IN SUl AT ION fUFt SVSl-FM OXIDI7FR SYSTEM PRFSSURI1ATJON SYSTEMS UJARTCATJNG SYSTEM
AFfCorVNAp.tIC r.CNTRms
FR IME P(1W ER SOURC FS FNGINF OR GAS GENERATOR UNITS flOWFR SOURCf TANKAGE AND SYSTEMS
PO~FR cnNVFRSTn~ ANn DISTRIBUTION fl FfTRICAl ~VDRAlJtIC/P~FIJMATTC
CUJ[ANCF AND NAVIGATION
INST"U~FNTATTrN
r(l"~Uf>if(ATJON
~~~IRnNMFNTAl CCNTROlS ECUTPMENT PFRS£lNNFI rrn ANT SYSTEM CCMPARTMFNT INSUl AT ION
" •
12788. 5193.
O.
18611. 12352.
41408. 16967. . 3678.
-0. -0.
. -0. 1389.
O. 3889.
160.
2114. 1074.
3598. 1120.
181. 2150.
o. 5214.
Table 10. Final Delta Wing Configuration Data
90
53593z~
77981.
31023.
16385. --. 61490.
5230.
3248.
4718.
800.
399.
2025.
1606.
J: F J( sr:" N E I PR CV T S TON S AccrMMODATTCNS FOR PFRSONNEt fTXFO tIFF SUPPORT FURNISHJNGS AND ClRGO HANDLING E~FRGFNCY FCUIPMFNT
CRfW STATTON C[NTROlS AND PANELS
fR't STRUCTtlRF
fFC::TGPI. RFSERVF
fFfCS(~'" Fl CRFW.GEAR A~D ACCFSSORI~S CRFW LTFF SUPPORT
PA'fl(lAD rARGf: PASSENr.FRS
USFflJl lOAD
~f~TlitlAl PROPFU ANT AND SERV ICE ITEMS TANK PRFSSURIZATION GASES TRAPPEr:: FUEL TRAPPFr. eXTOl7ER SFRVICF ITEMS RESIOUAlS
PFc::FRVF pRnPFltANT AND SERVICE ITEMS FUFL-MA TN PRGPUIS ION CXJDT7FR-"ArN PROPULSION P(1W FR SOli RC E PROPFt LANT S lijRRiCAt~TS
\IF 1 STRIJCTlJR F
IN-FLIGHT IOSSFS fUEL VFNT OXIDI7FR VENT POWFR SOIJRCE PROPFllANTS UJAR ICANTS
~ATN PROPFLLANTS FlJFI
TAKFOFF.ClJMA.ACCElERATF CRUISF DFSCFNT tOTTFR LAND
DXIOI1FR TAKFnFF.CIIM8.ACCELERATE CRUJSF OESCFNT t n rTF R t.AND
T A "EOf F WF U;HT
69545. 94032. 14175. 15964.
2822.
-0. -0. -0. -0. -0.
4700. 308.
8116. 305.
(
1250. 25.
13000. 35000.
, L
204 • 1014.
o. 107.
o. -0.
170. 83~
(
1528. -0.
3397. 330.
196646.
-0. t
13428.
300.
284226.)
; o.
1275.
48000.
49275. )
1386. ,
252.
"
"
~ j') L:f9. , -
5254.
19664'6.
I'
531040.)
~\
}
)
)
~ ~~u ~ ~<'" ...... ~yt) .. , %/ ~/ , "'" c: b. {!,? "" '\
t:.:~~ "l..,? ,
..., . ,':;,'
"r.Uj~FS POLY STRLJeTlJRF rRFW ANC PASSfNCFR COMPARTMENTS CARGO COMPARTMENT IA~CING GF~R PAYS PRrpUlSTON RAY hITHtN RODY ~ING CFNTFR SFCTION flJfl{CR COlo'MON PROPFLLANT) CONTAINER extOllER CONTAINFR FUfl (f1R. Cf1MMON PROPELLANT' INSULATION rr .. fR AnCY VnUJ~F
TOTAL RnDY VOLUME
kF1TFD ARFAS CROSS RODY t flWFR SlJRFA(FS(THERMAl PROTECTION) UPPFR StJRFACES(THFRMAL PROTECTION) PFRSONNFl C(MPARTMFNTS (~RGn cnMPARTMFNTS
HAN AR fAS wINGfOR I 'FTING SllRFACE)(GROSS) FxpnSFn WING AREA PflDV MAXIMUM CRnsS SECTION PAS F "FRTICAI SURfACES ..,OR r 7 ONT At SURFAC FS AIR INlFT CAPTURE AREA
liN IT WF Ir.HTS T " .. ~I~U
"FRTJCAI SIIRFACES ~nRt7nNTAL SURFACES pcnv ~TRUCTURF (BASIC) tIFTIN!; SURfACE MAXIMUM lOADING
CI"'F~,trNAt DATA "'NG
STRlJCTURAl SPAN ,RCOT f.HORO lENGTH Th'f.KNFSS RATlO
'Frey t fNGTH wIrTH ... FIGHT
FNCI~F ,CAlF FAcrOR
CU FT 9931. 8645.
903. 270.
o. 4741.
464J6. -0. -0.
12513.
83418.
SO. FT. 19862.62 5561.53
14301.08 4831.66
442.80
SO. FT. 6630.12 3880.54
556.15 572.04
1078.71 1206.68
113.11
lB/SCFT c:: 7'7
--;;T
6.91 6.56 3.66
81.00
FEET
159.59 134.68
0.06
314.48 23.15 30.92
5.39
Table 10 (cont 'd). Final Delta Wing Configuration Data
91
CU M 281. 245.
26. 8. o.
134. 1314.
-0. -0.
354.
2361.
SQ. M. 1845.24 516.67
1328.57 448.86
41.14
se. M. 615.94 360.50
51.67 53.14
100.21 112.10
10.51
KG/SGM f'I ?h
0.31 0.30 0.17 3.67
METERS
48.64 41.05
95.85 7.06 9.43
,
,
,
!
TiMF AL T ITUDE VElOC lTV WEIGHT RANGE THRU<;T DRAG MACH Q GAMMA AI PHA (LAVAL COACT CDeDR COO
'j COL (OF AlO PR SFC
PFC WA WF wox AE FIIH OXIDflER RIC OVOT VI
,. PTl CMl J;?'
~'!.
* :0: :0: :0: :0: :0: :0: :0: :0: :0: * :0: :0: :0: :0: * :0: :0: :0: :0: :0: :0: * :0: :0: :0: :0: :0: * :0: PI
70. 50. 2<36. 5"5178. 1. 341207. 117450. C.27 104. 8.56
15.00 0.1159 0.1129 0.1129 0.0021 0.H)11 0.0098 1.03 1.000 0.980
1.Or) 314;;> • <14. -0. 3960.00 lA67. -0.00 c. -0.52 968.
15.433 1.00
67. 5000. 659. 531230. 4. V56096. 53A63. 0.60 444. 13.31
4.AR 0.1654 0.C183 0.0183 0.0021 0.0073 0.00R9 9.03 1.000 0.942
1.0(') 3110. 93. -0. 3960.00 5809. -o.on 151.0 10.<10 968.
15.599 1.00
9A. 15000. 846. 528162. 8. 319A04. 5635? 0.80 535. 19.84
4.17 0.1333 0.0159 0.0159 0.0021 0.0050 0.008A 8.39 1.000 0.874
l.N) 7581. 78. -0. 3960.00 8877. -0.00 2e7.19 5.15 968. ;; 1, • 1.-
17.649 1.00 -b~ ,.,-
') 60S
/ 178. 75000. '315. 526078. 12. _~·/-b
. / ! .
761350. 53560. C.9O 446. 19.57 . V>
4.76 C.1602 0.018i 0.0181 0.0021 0.0069 0.0091 8.84 1.000 0.828
1.00 1977. 60. -0. 3960.00 10961. -0.00 306.50 i.95 '168.
9.736 1.00
un. 37'500. S88. 523300. 20. 1/ ':-,,'; 4- I 189B70:..., 293.2]..,. 1.02 323. 9.48. / ~<I __ - 5.('H 0.2318 0.0371 o~\LUl 0.0149 0.0175 0.0097 6.25 1.000 0.785
1.N) 1341. 4h -0. 3960.00 .,
\G, '--
13740. -0.06 162.72 1.51 959. -! j)..~
, )' I
'5.9ft4 0.9A I, --. i , \/~ ,
711 • 39'500. 108'5. 522164. 25. 18AQ87. 110141. 1.12 353. 3.29 or
'5.'50 (1.2143 0.0470 0.0470 0.0274 /1 • 0.0101 0.0095 4.5p 1.000 0.778
. ::.
1.03 1336. 41. -0. 3960.00 ~?r 14876. -0.00 62.31 3.0'2 948. ~' 6.117 0.95
745. 41700. 1162. 520768. 31. 1831'56. 129'531. 1.20 365. 2.60
'5.41 0.2073 0.1)535 0.0535 0.0321 0.0170 0.0095 3.87 1.000 0.774
1.04 ]297. 39. -0. 3960.00 16271. -0.00 52.76 1.86 982.
) 6.091 0.97
Table 10 (cont'd). Final Delta Wing Configuration Data
92
)
,
•
)
TTMF THRIIST AI PHA
(f) I PFC
F!JFI PTI
Al THUOE DRAG
CLAVAl (OF lolA
nXIDJ7ER eM1
VELOr. ITY MACH
COACT ALD
WF R I(
WE IGHT o
CDCOR PR
wax OVDT
RANGE GAMMA
CDC SFC
AE VI
* * * * • • * • * * * * * * * * * * * * * * • * • * * • * ~'\6.
IA1187. ,).O'i
o.n17"'! 1. n6
19A1"'!. 6.796
411. 707391.
4."0 0.0101
1.10 7'ln9.
A.791
4AO. '-771104.
4.16 0.0086
1 • 1 1 76061. 10.6A1
,)7A. 7<;')776.
'l.Al f).006R
1.1/ 7R,)74. 14.015
~.
3.17 O.OO",:\A
1 • 1 '\ ~~7,)6.
19.,)A7
660. 417471.
7.18 o.oon
1 .14 3A<;,)7. 6?091
70" • 4f-47~6.
7."i4 O.nol,)
1. lA 43391.
17;>.977
4'iOOO. 147H4.
C.177"'1 0.0097
1335. -0.00
1.10
47000. 153800. 0.1484 O.OOAA
1450. -0.00
1.44
4AOOO. 1"6')16. C.l1/1 O.OOB5
11'>45. -0.0(1
1.66
4A100. 119014. 0.1016 f.OOAl
lAAO. -o.oe
1.R6
49,)OC. IIIILL.
0.0664 0.0077
76Rl. -0.00
7.11
"iO('l(JO. 747657.
0,.0464 0.0064 ~905.
-0.00 7.R6
50400. 769818. 0.0346 C.O(l,)A
4084. -0.00 "3.1~
1356. 1.40
0.0506 3.50 40.
34.32
1550. 1~60
0.0460 3.22
43. 75.S0
1743. 1.80
0.0413 2.96
48. 1A.Ol
1937. 2.CO
0.0312 2.73
55. 16.46
2421. Y.J<J
0.0293 2.71
76. 12.54
2901). 3.00
0.0239 1.94 108.
11.52
3370. 3.4R
(l.0201 1.72 117.
lC.81
511161. 474.
0.0506 0.991
-0. 2.01
513810. 504.
0.0460 0.915
-0. 2.51
510973. 6C8.
0.0413 o. q'i 7
-0. 3.49
508516. 726.
0.03?? 0.938
-0. 4.55
503284. LlJ., 1. •
0.0293 0.885
-0. 7.59
49848A. 1534.
0.0239 0.826
-0. 11. 15
4'B641. 2076.
0.0201 0.826
-0. 1Z.57
Table 10 (cant 'd). Final Delta Wing Configuration Data
93
50. 1.45
o. 0291 0.765
3960.00 l1Z6.
70. 0.96
0.0265 0.762
3960.00 1444.
86. 0.59
0.0243 0.763
3960.00 1656.
101. 0.49
0.0223 0.769
3960.00 1861.
-u-.-:r .J
0.0182 0.818
3960.00 2362.
1 I) 3 • 0.Z3
0.0151 0.934
3960.00 281)6.
176. 0.18
0.0128 0.906
3960.00 3326.
TrMF Al TJTtJDE VElOC lTV WEIGHT RANGE THRUST ORAG MACH 0 GAMMA
') AtPHA CLAVAL COACT CDCOR COO f.0I COF ALO PR SFC PFC WA WF WOX AE
FllFI ox r IH 1 FR RIC OVOT VI PTt CMI
* * * * • • • * * • * * * • * • • * * * * • * * * • • • * * 707. '50'500. 3390. 493460. 177.
466~n. 210034. 3.50 2039. 1.06 /.'54 0.0344 0.C200 0.0200 0.0127
0.001'5 (J.0058 1.72 0.826 0.905 l.t6 4100. 117. -0. 3960.00
4Vi80. -0.00 62.99 12.2C 3345. 1/'5.869 3.35
148. 56400. 3814. 488803. 201. 438159/. n60'51. 4.00 2009. 2.18
'/.64 0.0343 0.0111 0.0171 0.0108 0.0014 0.0056 1.94 0.826 0.900
1.19 3842. llO. -0. 3960.00 4R,/17. -0.00 147.59 12.11 3828.
tAR.R'56 3.81
199. 61'500. 4363. 483961. 236. '117887. 16899'5. 4.50 1496. 2.41
'1.13 0.04'56 0.0110 0.0170 0.0093 0.00'/0 0.0057 2.68 0.826 0.912
1.'/6 2826. 81. -0. 3960.00 '53079. -0.00 183.16 8.56 4308.
211.374 4.24
) R6'5. 76400. 4878. 479180. 286. ?I)OR09. 1~4709. 5.00 1212. 1.42
3.60 0.0'555 0.0168 0.0168 0.0081 O.OO,/A 0.0059 3.31 0.826 0.939
1.34 nIH. 65. -0. 3960.00 15781)9. -0.00 120.47 7.01 4814.
?'52.fl13 4.64
947. 84700. 5396. 474206. 356. ;lnnoo. IP82? 5.50 995. 0.94
4.13 0.0665 O.GI71 0.0171 0.0071 0.0031 0.0062 3.89 0.826 0.977
1.44 1897. 55. -0. 3960.00 62834. -0.00 88.29 5.54 5321.
'/99.'/'50 5.0'/
10'50. 91300. 5913. 468920. 451. 171/39. 100864. 6.00 873. 0.56
4.1)4 0.0744 0.0174 0.0174 0.0064 0 .. 0(41) 0.0065 4.27 0.826 1.021
1.'5'5 1666. 49. -0. 3960.00 6Al,/O. -0.00 57.46 4.52 5829.
'170.619 5.39
) Table 10 (cont td). Final Delta Wing Configuration Data
94
T J MF III T nUDE VElOC lTV WEIGHT RANGE THRlIST DRAG MACH 0 GAMMA M PHA Cl AVAL COACT CDCOR coo
rOJ r. OF AlD PR SFC PH WA wF wax AE
) f'tJ FI oxrOllER RIC aVDT VI PT1 eMl
.. * * * .. * :0< * * .. * * .. * * * * * .. .. * .. * .. * * * .. * * r ! .-- ---, P08'i. ~ 5959. 467494. 485. 1 ('96A. 90~65. 6.CO 515. 1.89
6.411 0.12'57 0.0265 0.0265 0.0064 0.0116 O.OOR4 4.75 0.826 1.021
1. A 1 11 29. 33. -0. 3960.00 (91)4'5. -0.00 197.01 0.78 5830.
20<;.'598 '5.n
[io/',f1J -~ 5959. * 313462. 4386. <:;> dDl}03. J 909A9.* <;09f.f9:- 6.00 515. *
A _ e:; 1 * 0.1266* 0.026(,* 0.0266 * 0.0064* *Start of 0.0118* C).()08e:;* 4.75* 0.826* 0.965 * Cruise
1. A?* 1131. * 24.* -0. 3960.00 * 16357A. -0.00 0.78 * 5829. • 20<;.192* 5. 13*
'5179. lC300Cl. 5463. 372819. 4494. 2007? 77390. 5.50 430. -0.42
6.25 0.1241 0.e271 0.0271 0.0071 O.DI 1 '5 0.008'5 4.57 0.826 1.016
1.70 930. 6. -0. 3960.00 1114721. -0.00 -39.61 -4.72 5340. 12'5.041 4.77
'5;>77. H12500. 4965. 372302. 4577 •
) 1 A1r.6. AD 513. 5.00 363. -0.06 6.85 0.1478 0.0334 0.0334 0.0081
0.01112 C.0092 4.42 0.826 0.977 1. (,9 1118. 5. -0. 3960.0'0
164 n8. -0.00 - 5.36 -5.34 4829. 72.7?'1 .4.32
53fl'5. 102000. 4467. 3718'n 4f..4f. 16349. 8'5165. 4.50 301. 0.08
7.'59 0.1795 0.0428 0.0428 0.0093 0.0735 C.OIOO 4.20 0.826 0.949
1 .6A 71 (1. 4. -0. 3960.00 165148. -0.00 -5.95 -5.93 4315. 40.f.76 3.86
'54'50. 99000. 3963. 371540. 4704. 1'542'5. 90478. 4.00 273. -0.53
7.85 0.1996 0.0500 0.0500 0.0108 0.0789 r.D103 3.99 0.826 0.936
1.61 660. 4. -0. 3960.00 1 6'5 '10 O. -0.00 -36.94 -6.20 3807.
24.68A "1.44
'5'114. 93000. 3453. 371198. 4756. 1'5'194. 93984. 3.50 275. -1.20
7.44 0.1992 0.0516 0.0516 0.0127 O.MAA n.OIOl 3.86 0.1'126 0.941
1.49 673. 4. -0. 3960.00 16'1A42. -D.On -72.39 -6.12 3308.
16.'167 "1.06
I /
Table 10 (cont 'd). Final Delta Wing Configuration Data
~fi
TTMF AI T nUDE VELOCITY WEIGHT RANGE THRIIST DRAG MACH 0 GAMMA AI PHA (LAVAL CDACT CDCOR COO
') r.0I COF ALD PR SFC PFr. WA WF wax AE
FIIFL nXIOT7ER RIC Dvor VI PTI r.Ml
* * * * * * • * * * * * * * * * * * * * • * * * * * * * * 5616. 86600. 2947. 370860. 4799.
1~394. 97166. 3.00 271. -1.54 7.11 0.2027 0.0541 0.0541 0.0151
n.O?90 C.OIOO 3.75 0.826 0.968 1.3R 687. 4. -0. 3960.00
166179. -0.(')0 -79.12 -6.23 2809. 10.R?~ 2.65
57!H) • 7R500. 7442. 370306. 4836. 14R77. 99107. 7.50 275. -2.29
6.60 C.200,) 0.0544 0.0544 0.0182 0.0764 C.0098 3.68 0.901 1. 421
1.7R 427. 6. -0. 3960.00 1667~3. -0.00 -<;7.50 -6.04 2313.
7.393 2.23
57R7. 69000. 1 <141. 369824. 4868. 14999. <19869. 2.00 275. -3.16
6.17 C.2004 0.0547 0.0547 0.0223 o .('In7 0.009R 3.66 0.946 1.244
1.1 q 395. 5. -0. 3960.00 1 fo771 6. -0.00 -Ul7.l6 -5.62 1815.
5.31'1 1.78
) ~RR4. 57500. 1453. 369307. 4895. 1471'4. 97981'. 1.50 268. -4.47
5.74 0.2057 0.0551 0.0551 0.0277 0.0176 0.0098 3.73 0.985 1.326
1.10 41'0. 5. -0. 3960.00 1 fl7733. -0.00 -113.10 -4.76 1309.
4.~?q 1.30
~ ~ .., 0/, , /.O"H. IU"/ • ">Vl.'V. 'uu. .
A778. 55112. 1.00 217. -3.37 6.'5'5 0.25'51 0.0384 0.0384 0.0125
('\.01'56 C.OI02 6.6'5 1.000 2.608 1.1"0 486. 6. -0. 3960.0C
168'579. -0.00 -56.94 -2.21 968. 4.0flA 1.00
610A. 4nooc. 872. 367971. 4936. 61''5'5. 41604. 0.90 223. -3.62 6.70 0.2478 0.0282 0.0282 0.0021
0.0159 r.Ol02 8.80 1.000 3.956 1.("\0 '567. 7. -0. 3960.00
16906R. -0.00 -55.05 -1.07 968. 4.618 1.00
7108. 40000. 872. 358497. 4936. 47.£,4'). 4264'-. C.90 223. -3.62
6.6'5 0.2455 0.0289 0.02fl9 0.0021 0.01'57 0.0111 8.51 1.000 0.800
1.00 418. 9. -0. 3960.00 17R'54i'. -0.00 -5'5.05 -1.07 968.
) 4.61A 1.0(1
Table 10 (cont'd). Final Delta Wing Configuration Data
9f
"-)
TrMF Al T I TUDE VELOCITY WEIGHT RANGE THRUST DRAG MACH Q GAMMA ALPHA CLAVAL CDACT CDCOR COO
Cnl COF AlD PR SFC PFC WA WF wax AE
F\lFI OXIOIlER RIC OVOl VI PTI CMI
* * • * • • • • • • * * * * * * * • * • • • • • • * * • * • 7A06. 40000. 872. 352008. 4936.
41083. 41083. 0.90 223. -3.62 6.46 0.2370 0.0278 0.0278 0.0021
0.0146 0.0111 8.52 1.000 0.803 1.00 413. 9. -0. 3960.00
18')017. -0.00 -55.05 -1.07 968. 4.618 1.00
801l. 25000. 813. 349937. 4964. 6136. 40797. 0.80 352. -5.24 4. 'i 1 0.1489 0.0175 0.0175 0.0021
0.0061 0.0094 8.52 1.000 7.760 1.00 1002. 13. -0. 3960.00
) 1 R710=\. -0.00 -74.32 -0.25 968. 8.l25 1.00
R377. 'iooo. 659. 343323. 5001. 6'i66. 43689. 0.60 444. -5.28 3.78 0.1160 0.0149 0.0149 0.0021
n nn~c n (HHIC 7 A1 1 ('Inn l~ lA"7
1.00 1883. 30. -0. 3960.00 pn717. -0.00 -60.54 -0.52 968.
l'i.'i99 1.00
8513. 50. 197. 340501. 5001. '511'J. 5257. 0.19 486. 4.00
3.78 0.1160 0.0149 0.0149 0.0021 0.0039 0.0089 7.81 1.000 1.842
1.00 600. 3. -0. 3960.00 1961i38. -0.00 o. -0.52 968.
1'5.0£>] 1.00
) Table 10 (cont 'd). Final Delta Wing Configuration I)lta
97
)
)
7.0 VARIABLE SWEEP CONFIGURATION
7.1 CONFIGURATION EVALUATION
Except for the obvious changes in the wing planform, this configuration is basically similar to the baseline delta wing configuration shown in Figure 3. Many over-all features that are good for the delta wing configuration are also good for the variable sweep configuration. Both configurations have a low wing, a top passenger compartment, circular non-integral tanks, a propulsion system installed under the aft wing center section and a main landing gear located outboard of the engines. The effect of variable sweep is therefore effectively isolated as a single parameter.
Except for the wing. all design requirements of the delta wing configuration shown in Section 2.0 apply to the variable sweep configuration.
The major wing requirement is that a geometrical pivot location be evolved that achieves the minimum changes in c. p. / c. g. travel throughout the wing sweep angle/flight regime.
Several preliminary layouts of the variable sweep configuration were made in order to study the effects of different locations of the pivot. The following general conclusions were revealed:
a) To minimize c. p. /c. g. travel. an outboard pivot (approximately 25% of extended span) is necessary (unless the wing center section trans lates fore/aft).
b) The outboard pivot allows a reasonable layou~ for the main landing gear (including tread, stowage and inlet/ engine effects).
c) Because of the high aspect ratio Wing. a center section carry through box is necessary, the low wing enables the box to pass under the fuel tanks without excessive lost volume in the body.
d) The overlapping portions of the wing and, therefore, the cut line of the inner and outer wing panels are dictated by the wing center section box and the stowed main landing gear.
98
)
7.2 SELECTED CONFIGURATION ARRANGEMENT
The final arrangement of the variable sweep configuration is shown in Figure 59. The pivot is located outboard at 25% of the extended span (22 ft from centerline) and on the 40% chord line of the outer wing panel. The leading edge sweepback varies from 20° to 70° .
Operation of the moving wing is achieved by actuatGrs mounted on the forward face of the center section carry through box. This arrangement has the advantage of reacting the major flight loads by an actuator system in tension.
The main landing gear hinges on the aft face of the wing center section box and retracts forward so that the wheels are stowed forward of the box. (Aft or inboard retraction are excluded because of the wing geometry and inlet interference, respectively. )
'7.3 AERODYNAMICS
Lift and Drag. The main emphasis of the analysis was to obtain the lift and drag characteristics for input to the aerodynamics subroutine of the synthesis program. Lift and drag characteristics were determined for Mach numbers from O. 5 to 6.0 with the wings swept aft,and for Mach 0.5 only with the wings swept forward.
Figure 60 shows a nominal schedule of wing sweep that was used for determining the performance of the variable sweep configuration. This is based on the data shown in Reference 21.
The analytical procedUres contained in USAF DATCOM were supplemented with current experimental-analytical data on variable sweep aerodynamics, such as Reference 26. The resulting lift, drag and LID characteristics are shown in Figure 61. Figure 62 presents a breakdown of the drag coefficient for a lift coefficient of .15. These data show that an appreciable percentage of the drag is due to the afterbody and friction.
The longitudinal stability characteristics of the variable-sweep configuration were determined. Although the baseline configuration was found to be statically stable throughout the flight regime, large trim deflections were found to be required to offset a large moment at zero lift, as was found for the delta-wing configuration. To demonstrate trim capability, this condition was corrected by assuming a positive rotation of the forebody of 2 0 relative to the main fuselage and wings. (Reference 27 reports a similar ~esign modification on a proposed variable sweep SST configuration to obtain a more positive pitching moment at cruise and good trim-drag charact.eristics.) Figure 63 shows the stability characteristics of the configuration for subsonic operation with full-forward wing position; deSired lift coefficients can be obtained with
j trim tail incidence angles of approximately _4 0• Figure 64 shows the stability and
99
r- -'.
~ t~D Of V!~W
SO'
-----_._--- _ sTj.' . . ~:;
200 ~S£NG.E.~ ---CMr.-----·-
Figure 59. Variable Sweep Wing Configuration
100
70·
4: I \ .... ·u·JG HINGE
44' .
.. _f~~GO ~Pi
'- M.:.. C. PJvo"l'" KoUNTING '$;",
·29~.---·
··--.L
Wlag area (gross) _epl:/fwd tIl1ckness ratio swept/fwd aspect ratio swept/fwd leading edge sweep swept/fwd
Horizontal area (gross) tb1ckness ratio aspect ratio movement
Vertical area thicklles8 ratio aspect ratio
Body length (overall) Ixeadth (max. ) depth (max. ) volume fuel volume
Lauding gear main nose tread
Engines type
thrust
Passenger 8 6 abreast .4 abreast Seat spaciag
CarlO t.y volume
120' ~PAN 70'" 50WEfP
200' SPAN 20' sweEP --.. --.-
7,350/5,000 sq. ft. .05/.098 1. 96/8.0 70°/20·
1,410 sq. ft. ·06 2. 1 -15 to + 10·
1,320 sq. ft· .06 .8
335 ft. 2S.3ft. 34'3 ft. 106,6liO cu· ft. 52,600 cu· ft.
four 66 X 18 tires/oleo twill 48 X 12 tires 35 ft.
Pratt" Whitney STFR]-2JOA (4) 75,000 Ibs. SLS each
24 rows 14 rows 36 illcbea
900 cu· ft·
=-=
~.
70
60
50
t'LE 40
:30 t-'
0 ......
20
10
0
\......J
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o 2 3 4 5
Mach Number
Figure GO Variable Sweep Configuration - Sweep Schedule
~/
)
)
gmgm MACH
.5 .8 .8 ~. f::=+:;::
~~~
::.:tr=t: ( t\ = 20°)
.&
."= j:::ti:_f--4.~t--...-,
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o
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. :
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J: ---:-.:~. e"
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",.
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.2
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6.
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:1,. v
R ;~ !; iii
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MACH
.04 .08 .12
::.c:':-: ~ .. -1-._+-- IV'OA ""TT f'~:- ; .. J. .In.\.."n
. ::1 ~
::;c" ::~. (A 20°) .. ~
10 12
Figure 61. Variable Sweep Configuration - Aerodynamic Characteristics
102
))
i I
Mach No.
SREF := 7,350 sq. ft.
ee = . 15
Figure 62. Variable Sweep Configuration Drag Breakdown
103
7
,..... o tn
C
.20 I' "li~:II:!I;I'.i!I.:III.:;!II"IIIi1llli\II:!;it:lr!l; :I!, i' " !11!!l \I: ,; I
it
00
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_8°
L .10
.05
o
o 2 4 6 8
(YO
_______________________ ~-~-•• ·o.
\~/
X :::: 200' ca
x = 60' REF
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ii:TI lIi! tdqt m:EH i nn l1 tnl If!l;J ~;L! tLTIf!ii!; IlHlWli\ !r\T~;
.06 .04 .02 o -.02 -.04 C
M
-''-.--'/ i
Figure 64. Variable sweep configuratiqn - trim at Mach 4.5
)
( )
trim capabilities of the configuration fully-swept at hypersonic speed. The directional stability was checked at Mach 6.0 and it was found that the configuration was directionally stable for angles of atta~k up to about 15 degrees.
7.4 WEIGHTS
Equations for the variable sweep configuration were determined from data generated under Contract NAS2-3025. The effect of having variable sweep was estimated from Reference 20, which showed an increase in wing weight of 25%. compared with a fixed wing. Other data were similar to the baseline delta wing configuration. Weights and dimensional data are shown in Table 11 for the configuration as shown in Figure 59 which is sized only very approximately for 5000 n. mi. range.
7.5 WINp GEOMETRY VARIA TIONS
The effects on take-off weight of variations in the following wing geometry parameters were investigated using the synthesis program:
a) Wing Leading Edge Sweep
b) Wing Thickness Ratio
c) Wing Aspect Ratio
d) Wing Loading
7.5.1 LEADING EDGE SWEEP. Leading edge s'weep variations were made for a fixed wing loading of 90 psf. As used here, the leading edge sweep refers to the maximum sweepback. Wing sweepbacks of 65 and 75 degrees were considered, in addition to the 'nominal sweep of 70 degrees. Figure 65 presents the effect of sweepback on take-off weight. The take-off weight decreases with increasing sweepback, due to the decrease in wing structural weight which results from the shorter structural span at high sweepback and due to lower drag. A check of the wing planforms showed that because of the decrease in wing span as A LE is increased, in order to in~tall
engines and landing gear under the'fixed inner wing, A LE should be less than about 70°.
7.5.2 THICKNESS RATIO. Wing thickness ratio was varied over the range from .025 to .07. Figure 66 shows the effect of thickness ratio on take-off weight. This data indicates that the optimum thickness ratio is approximately .05. As with the delta wing configuration discussed in Section 6.0, decreasing thickness ratio below about .05 improves the aerodynamic efficiency; however. this is offset by the increased wing structural weight. For thickness ratios greater than about. 05, the reverse is true. The lower wing weight due to better structural efficiency is offset by the loss in performance due to the higher drag.
106
)
AEROCYNA~IC CevICES WING t wING MOUNTED CCNTROl SURFACES VERTICAL S~RFACES HORIZCNTAL S~RFACES
FAIHI~GS,ShRCUDS ANC ASSOCIATED STRLCTUKE ,
BOey STi<LCTURE
100566. 8670. 9864. 2111.
STRUCTURAL FLEl(OR C(MMO~ BASIC ENCLCSING STRUCTURE PRESSLRIlED CGMPARfMENTS SECCN(ARY STRuCTURE
PROPELLANT) tONTAINER Q. 94881.
5193. o.
INCUCED ENVIRC\~ENrAl PRCTECTION COVER PANElS,NCN-STKLCTURAL INSULAT leN
LAf\OING fE~R
I"IA IN PROPUL S IC ~ ENGINES AN[ ACCESSORIES AIR II\CUC TIe f\ NACElLES,FCDS,PYLCNS,SUPPORTS FUEl(CR CCMMCN}CCNTAINERS AND SUPPORTS OXICIZER CO~TAI~ERS AND SUPPORTS PROPELLANT II\SLLATIGN fUEL SVSTEf'o1 OXICIlER SYSTEM PRESSLRIZATICN SYSTEMS LUBRILATING SYSTEM
~EROCVNAMIC CCNT~OLS
PRIME PCwER SOURCES ,.. ;"'. ,r nrC' rC~lcnAT,..,n rT" L. '''''~ L. V V I.... "' ... I... ,.... 'J v".L' ..J
POWER SOURCE TANKAGE AND SYSTEMS
POkER CCNVERSICN ANC DISTRIBuTION ELECTRICAL HYCRALLIC/PNEUMATIC
GUICANCE ftND ~AVIGATION
INSTRUMENTATION ~
CO~MUN ICAT ION
cNvIRONME~TAL CC~THOLS
EQUIPf'lENT PERSC f\N EL CJOLA~T SYSTEM COMPARTMENT INSLLATICN
0. o.
53813. 22500.
4259. 20968.
• "1 -..:.
10270. 1679.
o. 4217.
161).
L.. • -' u •
1358.
4550. 1339.
224. 2150.
o. 5274.
121211.
lC(lf)74.
o.
20307.
117866.
6494.
3817.
5890.
800.
417.
2025.
7648.
PE~SO~NEl P~CVISICNS ACL(~~~OAIIC~S FOR PERSC~NEL FIXEO LIfE SLFFORT FURNi5HING~ AN0 CAHGU HANULING EMERGENCY EQlIP~E~T
CRE~ STATICN C(~lRULS A~[ PANELS
CR y S TRlC TUR E
CESIGN xESEkVE
PERSCNI\EL CREw,GEAk ANL ACCtSSOkIES CRE~ lIFE SUPPuRT
PAYLOAC CARGU PA~SENfERS
USEfUL llJAC
RESIDUAL PRUPEllANT AND SEkV[CE ITEMS TA~K PkESSLKILATICN CASES TRAPPtC FUEL TRAyPEC OXILIZER SEKvICE ITE~S RESICUALS
RESERVE P~CPElLA~T ANC SERVICE ITEMS FUEL-~AI~ PR(PLlSICN OXICIlER-~AI~ PROPULSION PC~ER SOURCE. PROPELLANTS LU8RICANTS
RET STRLCTUKE
IN-FLIGhT LCSSES FUEL ~ENT
OXI!.JIlER vEt\l POwER SOURCE PRCPELLANTS LLtHHCANlS
~AIN FRCPELLANTS fUEL TAK~OFF,CLIMS,ACCELERATE
CRuISE CfSCENT LO ITER LANe
OXIDIZER TAKECFF,CLIMB,ACC~LERATE
CRUISE DESCENT LeITER LAND
TAKEOFf wEIGHT
-:;
-0. -0. -0. -0. -0.
-0. -0. -0. -0. -0.
13428. 4700.
308. 8116.
305.
300.
( 4£:0277. )
O.
1275. 125n.
25.
48)00. 130!)0. 35(00.
49275. t
1725. 231.
1358 • \ o.
136.
263. o. o.
18U. 83.
( 451539.)
5598. 1666.
c. 3602.
330.
222000. 222000.
o.
f 679137.)
Table 11 (cont'd). Variable Sweep Configuration Baseline Weight and Size Data 107
VOLUMES BO[Y STRuCTURE CREW AND PASSENGEH CCMPARTMENTS CARGO C(MPARTMENT LANCiNG GEAR BAYS PR(PULSIO~ eAY wIThIN BODY WING CENTER SECTION FUELlGR C(MM[~ PRCPELLANT) CONTAINER OXIDIZER CCNTAINER FUEL(OR C(MM(~ PROPELLANT) INSllLATION QT~ER BOCY VCLuME
TOTAL BODY VClUME
"ETTEr AREAS GROSS BODY LOwER SURFACES(TrERMAL PROTECTIO~) UPPER SLRfACES(TrERMAl PROTECTIO~) PERSGNNEl CO~PARTMENTS CARGO COMPARTMENTS
PLAN AREAS WING{GR lIFTING,SURFACE)(GROSS) EXPOSED wING AREA BODY MAXIMUM CROSS SECTION
·BASE VERTICAL SuRfA~fS HORIlCNTAL SLRfACES AIR INLET CAPTURE AREA
_ _ n'- ....... v
wING VERTICAL SURFACES HGRIZCNTAL SLRFACES eocY STRLCTuRE (BASIC) LIFTING SLRFACE ~AXIMUM LOADING
DII"ENSIONAl DATA ~ING
STRl..CTURAl SPAN ROOT ChORD LENGTH THICKi\ESS RAllO
EOCY LENGTtWIDTH rEIGHT
ENGINE SCALE FACTOR
Cll F T 10821.
8645. 903. 210.
o. 2247.
52421. o.
3292. 18438.
97041.
SQ. FT. 21653.44
o. o.
4831.66 442.80
SQ. FT. 7342.02 5347.48
623.62 623.62
1193.17 1439.04
1 c:;t) r./\ ... ..,.,v • ...,v
lBI SQF T 13.70 1.27 6.85 4.38
92.50
FEET
265.63 97.68 0.05
320.63 24.71 33.03
7.DO
eu M 306. 245. 26. f 8. o.
64. 1484.
o. 93.
522.
2146.
SQ. ~.
2011.60 o. o.
" 448.86 41.14
SQ. ~.
682.07 496.78
57.93 . 57.93 11C.85 133.69 , ""') ,,~
~::I."j
KG/SGM 0.62 0.33 0.31 C.20 4.20
METERS
80.96 2<; .. .,
S. 7. ~:
10.01
Table 11 (cont'd). Variable Sweep Configuration Baseline Weight and Size Data
108
i '
~ )
/
)
)
o o o .....
.02
60 65 70 75 80
Leading Edge Sweepback
Figure 65. Take-off Weight vs Leading Edge Sweep
.03 .04 .05 .06 .07 Thickness Ratio
Figure 66. Take-off Weight vs Thickness Ratio
109
7.5.3 ASPECT RATIO. In addition to the basic variable sweep wing, wings having aspect ratios of 1. 47 and 2.61 were also investigated. These wings had the same leading edge sweep and taper ratio as the basic variable sweep wing. Aspect ratio is computed in the swept configuration and wing area is defined as the area to the centerline exclusive of the glove. The effect of aspect ratio on take-off weight is shown in Figure 67, which indicates that increasing aspect ratio increases take-off weight. The corresponding unit wing weights and performance fuel as a percentage of take-off weight are also shown in Figure 67. As the aspect ratio increases, the aerodynamic efficiency increases, resulting in a decrease in performance fuel. This gain is offset, however, by the increase in unit wing weight. The unit wing weight increases due to the increase in structural span.
7.5.4 WING LOADING. The effect of wing loading on take-off weight for the three different aspect ratios is shown in Figure 68. The corresponding wing weights and performance fuel weights as a percentage of the take-off weight are also shown in Figure 68. Increasing the wing loading decreases the take-off weight due to the fact that the decrease in wing weight offsets the decrease in lift to drag ratio. As with the delta wing configuration discussed in Section 6.0, take-off velocity requirements place an upper limit on wing loading. For the take-off calculations, a take-off angle of attack of 15 degrees and the use of high lift devices were assumed. The following take-off lift coefficients were established for the three aspect ratio configurations
) at a wing loading of 90 psf;
)
Aspect Ratio Take-off CL
1.47 1. 96 2.61
1. 74 1. 87
1. 92
These lift coefficients were adjusted for wing loadings other than 90 psf to account for the fact that the percentage of wing available for high lift devices changes with changes in wing loading. A ten percent velocity margin was assumed. The upper limit on wing loading is shown superimposed on the data of Figure 68. This data indicate that even when considering the take-off requirements, the low aspect ratio wing yields the lowest take-off weight.
7.6 FINAL VARM.BLE SWEEP CONFIGURA TroN
Based upon the configuration evaluation data described above, a final variable sweep configuration was se lected having the follOWing characteristics:
Wing Loading Wing Sweep Wing Aspect Ratio
125 psf == 70° :::: 1.47
Wing Thickness Ratio == .05
110
12.
)
)
Body Fineness Ratio = 12 (same as delta configuration, since bodies are essentially the same)
Integral Tanks
A final synthesis and performance run for this configuration is shown in Table
111
.... .... ~
\~,
""1 ~. >-j (1)
m -l
>-3
~ , o :::: ~ ....
oq t:r' .. <! Ol
;.-en '0 (1) n ... ~ ~ ..... o
.... o
Unit Wing Weight (~ psi)
.... .... .... t.:>
.... C..:>
.... II'-
.... 01
.... .... O'l -l
...... .1~." \I ", tlt!l 1,1. H ~ . H H t. .J I, . ,q • I . ~T n~ rrmmffTTT1T'1'rTI"i Pf mmm mn rrrn " 'r rml1lf'I !~l E:'t111l !W'lii! Willi '1;;liU!:'HH1'JiJf 1J tI!!! '
:'I! ';; itl' 1 !l i'P ;:lir:ffit 1:; ,iii iii' i11i! " r 11ft j Ii ~ H::!: i!::~!1 'HI! !.,: ;11' ::JJm·':i1· ,it ; . i'l iinnJ!
t"'
. ! 't" ',ii.i I . lll' " II" IH Hj' I .I HI" I" : l~ :lE 'II I • up r m Iii' .:! I I till fmhi +it
f I ~; ;: j • : i. 1 : r; ~ t ,t!:', ! t, " ';;
JI~i j,1.1~iM , I 1m llillr' u1l!jf' llIil'll .11Ij] , " jFi~W~itn~!~, ,',i: iHl 11!11!l11:f!tl!n'lt tlJlcrlHffj\
..... I II .,~ .q ::t1; H· !' .. JHt " 'II' .1 t ,LltH-
. ') ;'~ :11 !!It! Ill; !I'I d "11 :!ll !i:: : IlliJ I W" I" I 0. ii· r ':, HJ+!Il1;llifH~' I,,: 111 l;U !/llli~) iI:,!!1
II ~: ;;11 II' I' lilli!: Wi lilliil np 'j , n;l tfii I· .1;" 'j~ i'lI! I!, '!!l}"i:I'ii{fhi,lltttl,' . "llJ !, 1m 1 I ! 11 'i r " '-e:l lit:
• i I; I" I'll!. f lI;! :f" HI " J .t· .!' It lJ.l , 'II jI "" , . 1 'Ii '~'i ~ I,ll ,I 1 II
(fl I 'tt • 'J II' 'g ," I, II' 1} l' 1.11 : g !". !:' , 1 j if ,., I!':I:[I, Ill! I ~I :;0 0 Iii I I !'t lj;1 iI III Itlll !i g, III L 1 I liL' tj I 'liH1111rr liw !ijt
l,l4' 'HI· i ! Iii I ' ,I[[j,q .tIl ifj'l iW rIll I l I I f r I Iii 1111 IlIhr Jli l ~'I
'~ltl~;jl j ! 1 ~I Illl!!d II 111 iii! ir~ , ,h it, 'I', itllllFf 1.'lf H',', 1'11r f@l. I ' I' I I t 4 I I ~ I t
Il'i! I ,11' j '11 Ell Ilil IH ilL it: Huh,. i fijt ,I lI[11l,litt 1:1: 11'1 llii ~ l' . , l1u ' I I· ,II "I 8+ffi
~ ! ijll' iUli I' ii!l!!fl: Ili'ii'j: '1' ' ',' I F'I jl, l' II iii I mil' rl~q I . ", . liir:rirrli~'r !i;! 'M
fIIlj '[I ill Wti1 J II i1llrn~H jill IllIW!
-IfIIHI1" !!i" l:t, II!.!l 'I· II lli'lllt~I"I.l!1 H!! I.ili tv .r II, dd I \'1 \, , I j.IN I lrf,
~ iUtlli ! '1: 1 I !]1I1 Iii IIJI Ull ,il! j'l Ijll ill: tHit' i 'Hlf if! t ·f 'It1~ .. , 1 "t i 111'~i " !I. ,IIi Ii till I 'iiil I' ttl I ii' h;' l~"'j mi '111 'I 1!H' t', W: ,,~ lift
mfi :11 1 ij 11m'!!) '1m! 11 "I • ' 1 md, '1 1, :.II I' r. III. Jl'], IH I '! ~ JtillI;1 ,j I I I Hil. i, II i!:Jllj
00
IV . IV
00
''..J
~. (fl
'0 (1) () M
Fuel Weight Take-off Weight
c.> o
f- -li-H+P"jf'+r:+.+++t II'-
c.> 01
N ~ 0 -4f!J+1+t+tl
N .... ' o
IV
IV
IV
""-
IV
0.
IV -all.l.U.>.C-.L
00
. ~ o
;.-en
16 g.
~ M o·
O'l o o
J
Take-off Weight (1000 Ib)
-l o o
00 o o
to o o
.... o o o
.... .... o o
... r.: <:: c
'lin ifml{li'l'fTl ,'1, q ft" ""j"t" l"llf iUi j Iii": i'i; lii ti! j':11 til II! I; I ,." .. I.. _+1t-+ _.'" .~.J.~ ~_ ...... ++.. 1-+., :;" ,;" fC':,,' "I"'; .,t! 'If! 11'1)1:.: 'F
.... ~:;I:" I:::;: !i::, 'i;;:IIII,;I::l::I[::,,::
~ C,::t:l' ;n.' ::iird" ni;~i~j' :j;!' ::~~ l!il~j,t;i,~jt: :',.,. ,,,., .," , .. , ,ttr Ii: , '" .L ··~1· t't···· ,I,; "~,I 1'1: 1 If!' liti :11; ';',;q 1',II'i' t:I;:h is h t ifi : t.;: ~H Wi :ftiffS t~ihl
~ ~;:~. !~lj: lj~]i~~ :;,1 ~I~!:i~.ij" w' ::1 iii! I! H!: ill; t'; :i'l "i: ;'If I,'!I' ' ,,"~ 111 '11 ' HI 'H, "U" I"'" ." ". !Ii Ii,,, 11', 11; !!, I'!" \'J'I Iii' Ij 11 j i I ti'l!i : iii MIT IT" ltt 11· 14 ' , jl +,~, H .' "r' Ir;!t~itt; :1" I,' :1, :,!: 11:<) Ir!!i;~11 I'" jJ If·: ...... ;jllflH il' tl, it t 'ii·1 'jlf !,!ditl t it t- t I II
~ ll!!PliL 1l1l1!ilj'[JI 0\ Il\iilfH,! I J 1\11 thi ;111 Hillin llirf tn . Jl iLl l11-TI ~r TnH[llt u fila 'fill P!~ I~I! lIlf ':1 Vlf,Hlt'i' ij~.m 'I ,t)l n"lj'Il~' jlHI, ,m .' "ii, '[ I
t.:> In; :1:1:l,! ;,,1, tHllllil .11 if 'jl! . I o 1i!1 :''''!i j::' ,[ ill II 1'lIt1 11, 'j ttll,·1t Hit illl ii.i, J Ii P p!:..(1 +H lil ltLi jill
i!!, PQlflhi:f: ;~~l~!!:n!i !\li !Jlj 1 'f!H1 illl iii; lih ::jij'ii l.Jl:~' ,tilt! nli Ilill jill • ire ,nt H-ti,l:h, Itl! j "rm t T, ~t'l H1' .'-11'!:; if!1 [ill I!:: t:,,' 'rl'l ,11' I ,tit: ITill'!.; L..... II " " I ' ~J, -t I ~ ,- t t, • , I ~ _
• tIl ')' I " tt. "'t' ,', ". \ " , ~ I" ji!,!,;" .. 1: :II! p:!l' 1'1' Hilt';' "j 1/', I'::: Hr: i-lfl' 1ft n:t +F:8~~ I !.~:' im'·, f:~~ .... ,', "tl !'" 'I . I'" liIl }JIllli I, , i'" rr+i'HHH.mlJjll~''; :::! , HIli H I; liU ,l!ll id\W/ [l:\ :i\!!l:~\ i Pi{ u!l fill Hll ~W
M iLi i1l!I,H ;II: Pi; :1,', 'I'! ;1 :;;, !1!! ~!!: ;If: ... '" -'11 I, III, ,I,·.i..!..g -L..::" I ,. t'" l- • ~, ~
~ ',ll'![i lii! fil' ;::':': It li~I i:l' i :t: ::" '",ltll 't .. 1 .+ .. ,11j ,~' 't'l jl'l + " .. 1.11 ,,, 'I "I"" ,,11 , .. 'I I\: ~I I 1'111 ,.' 11'1 jll' il! 1\1' Iii' ,1: ' ,I 11!,1 ,: , " 11:1 ~HttH+-H ri- r :~ I ;",! ":, I, *h j,l; I I, I. : ,,1' ,Pt,l' ':: lilt ii, 111'Ij']::' " f:!~ nJ: ;t~·I~Jl. ;~~, in! ;Jl hH H ili .. ~t: HI'
t.:> li~", " tiL I,,·till: :'11 l~' t':! iii: h ;,., Trii:I"I"'; .: ji:'; ,·",mc t;li Itji irfi il II;:! 7:"':"
1,11 ,!;!!liil H11 LH!1 jiil :d 'Hi lw +I[jl!i': ::" :Itt 1:" Ii ',,j ,rlt :11· :'Ii :':: ,bEF,r; " ! : '! l! J: ; i ! ; ~ !.! ~! ~::: ; 1., d.; 1.1. ' ;; i ... ; jI' :,'i i,i \: ,t :ill :11) 'I;:::ntl' i;:' I,: :)~ gi; 1i ~Hl L:~~ ~i~l J.!! ! dF .Z~·tl ~ JUP ::i f!;!;!;;; ::n iii, r ILi r ff lJif !rU
00
'-----'
:0 -<
0 0 0
-<-' ..::: bll ...... Q)
~ ,,-, 4-J 0 ..... I
....... Q) W ..':<!
('j E-;
~)
~, Take-off Speed 160 knots
I00°r;iFiffiTPiTC:-f:J;; ~"::, ::~r~~: FL jTv--:-~nB'Tifl~F 71J;ITr-:: :::: :;,;1:::; :;:~ ;;: /::;: IT?:::: Itlil !lllii l , ;/'1 ' If "I, • ,~,. -Ifj~f,,'I: .I';~t !., '" ,.j' ~ I, !I'~ ,1. 1 !" 'I-"j "':r:'I; ';1 ." 1":1"" l'lf I!., tit·",·· '"'-rl ••• '" ,I" "11 "'J.'" ""I" ""I""~"I"'" , ..• j,," "'llil'" .," .,,' "'1'''' I· .. h-· "" I'" ... ~ .. , ...• 1 ....... , ",' ·t .... ,. ;Jf j!.: ill: j;~: ;JI [; ';.: ;:'::: ::!llP:'· 'I :~'; ;~"I:::: :::; :;'; :::i .::: :;::·1;;: dt;i,r:' ::tl It: :;:11;;:: t;l; ;~!: ;;i Illl :1:: ::~! I' f ':I~ .t:; :;:: ::.: :::: ::~; '.1 :':' ~~, ·:-:P.::t:::: ++r ;:~~n";: :.~' :::: ~~: ~I; :,.:',' '~:: ~::: :~~~ :~;. :::; :~~ ;':; :::1 :·~tt ,jtl ,,', "" ,'" ,,1'1 "" .. ,:i:, , ,,, :\"" ,", .. " ...... ""j' ". I " .. " ,", I'" j, I "" :1 d" .. ,:1 "1 ',,, '" ", /1'1 ·t:1 J'" : ; : 1 ;.: i :: t I . : If; !;;! : ~ l' .. :!:: j : ~ l' ;)" ::; " . : . : ; ! .; , i: ;;. t I!:; 'i; ~ . ~:; !.!: ; ! ~ ;'; i I ~; ::;; :; t ' L : :. ;.: i :;:: ; i; ,;t : , : ~ .::; "I'"'' P"tijJ., "+'1"'" '1"'" , .... 1. ....... I ... ·',·t·· .. ·'··1' .. ·· .... I, .. , ' " 'Tt 11,,/").., 'r.~" •• ''''j'''t ........ ', .. 'HI 'r"I"" itf~: :\~1 ~tH'H~ :~!! i::: i:::!;!:: ::::I>i1 :~::'I:;':~":::\ :ii: i:;l :i:; !!i: i~!: ;it; !d~ Hi! :~;~ !it: i~;: Hit ::U H!: i:: II~~' ;tj'l i;'j!
000 ":1 l~I'!J.l·fH·'·'" "" "'ii"" """';L .• " ., , " !""ftG'" "i' '''t''· ." ,,,I ~fEt!S'" ,' .• ,,, ,J,1 ''I, ',,1 ",,t, ,', ,;j ,I ,t 1 iJ ----':-Ij. itt r-;-r!" I • Ii ',. :'I,!","-:,;rl--~;--:; .~:-r--:~7 !:T .. : "'! , •. j 1'-:--;-, ·;i .1,11. ~ ~ , •... :-r;, i~·11 ';~l 1/' I' ,!l ." .. ,!, Ill' 1"1 I.! I ,li. ;1" I,. " " .• '.,- •. j ····1 ., •• ' ,,! ,JI, .,.: ",; .,1'1."; .t .. ,.l,! ...• 'f-j '!i' .;;1 'I t 1'1: ... , tllJ l'I' :::1 i:ll[)f~'!!':i :;~ .:::::,:,' -;::1:::::.:(::, :::,;.:. ~i:E::: ";:1::;: :.:,;1\: ["Y!: !!i' 1'1:: :i;: Jill I;" i:i; :~::~j:: d:[!:l~.
iiil ii~l j~!: ~:-: ;::d;U~~j:~~' ~~~i-~i :"~_:~ii:: GL~ ;i::'~;~i <;;~L2lI: :lU~:i;: :~~1~~ :;;~ 'ff{ IT Ul1 q:: i;:! ~~~~ ii!~ ;~;;p~~; !l-~i!~lii ;;i~j~~~_; ~;~~ ';.~ ~i::L~i; t~:'!::;~ ;i~~!~:~~ ~~~~ll!~1 ;~~~jin} ~::: :iF ~:~;Ljl ~~~: .~!E tiL! !lft fJllHH !iIl l~L ;iH.Eti !~;!!!i~~
J
H-:, , .... , ...... rtR"····f7':~,,, ,1'''7 'j' . '" ~ftH4·: .. ':-:-~~;,'I" '" 11'1 ",Ym 'm"" '''' "/' ""1:'" :;:il';~: :'::1:" ;:::(, N.J':: G~ :::,\::: ;:::1:,:, ·;mii : .:i;, ' ;:::;:;:: :i:11ij IHl hi lit:~i';ll::i ti': ,:,diU!l; il::':":
80~-~~::; :~;~ ~t;t+8H'f;~~~ ;:';':~': :~;;*H--•. ;::'~!:f8f+ :;;:): ~;:;I;::t?:.:::; :::~ :::; rrtn'mt"fH+ 'ttl1t~: J~.: ~:~?H-il: ~:~! ~~~~ l,I,iH,,, .. "',, .. ""I ,,' , ,~ .t", "I '''' It, ", ",' " .. I ....... ,r,.I, ... j~ ... ,:1 '1 f '''1 ,j.,.,j "" 1 .. 1 '/'1 ittl ,'" .. " 1+1~1.;~~: :::;',:::; :.:;.!,~~!r :: .. ~ ;_:! 1 ..... ;:t -I;.!J -.. :; d-' '::! ;;:~ i~;: ::::t:..;~: ":-~:!~;.:~ .. -+ U~t + 1'1 '!' ~:: .. '::L:i ~;"t:- t"l ~~!. ~t :::.: ~~t.;
. "all" " "I·t
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<1 I "~" t· ", .• , , •• , •. ~ J .. , •• '1 "~11 , •. "r' "T' ,,1,1·· •• ". ,.~ .' .• , • .; •• ,. ! • .,.. .- t ~ .. {, >.t1- 'f •••• .,/ .... , • ... ·1· .. • "I.,.! ... , ...... " ,', ........... ,} . ." H," •• ., , .• " 'I'" - ... ".+ .•. , ':T!'''ITffo!t1tt .J. '1"-:-::111.·fl1"1 + 2.61 i-j::t;:;~ :;:: :~:; I:': ;~'; ;::t .:~ !;;;tr;;: :;:1(::;; ;;,:1:;:: :::; ;.!; tiU ;;~: ::;;t:: •. 1 ~ I, ~ lit' :n trIo +"!! , I I I' ," r"; f • , j , I ; ; ; : ' • . '. • 1 'I f.· . '. !: I 1 ' " r 1 •.• i 1 • " • It·· t" "1' • r T r -,.. "t t . r ' - ~ • , 1 • t , I + t t. 1 f '"t ., + ,J •• H+TH-~~-;+:-ii;y'~j-:+H" ;-HiH+ *t++P: .::;';':: ':': :.~, -:::; ::', :;-:.i:SF~ti:': tit ":J t t: t·" +1 ':::I:~;~ 'II',I .... I!: .. ! ... , "/11 "'1' '".",. )"/ "., , ... 1"" Id.l,t" ,q, , .. , ,.+.j:,., ''''1'''- !!t~l:l~! . 'l'I'mt "-L'I' f'jt'l'l'I'~ 'tm ... ~.t'"
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H'11l1IUlI~il HllJ lllj~ i!jl Ujl IW1HU\lU l!Hi\~il tHhllH ilit:i. n Uti ali!l [11 1 1U! ~i; . t 500 IttllmltfHl lin Inu ' itll 1m llHtlWlilllH!li i!!1iHu lPillfi:lfl!1 :iil!UH nnn mHl t1 ii! t til
~ I
70 80 o 100 110 120 130 140 150 Take-off Wing Loading (pst)
Figu~e 68. Take-off Weight vs Wing Loading
TIME ALTITUDE VELOCITY WEIGHT RANGE THRUST DRAG MACH Q GAMMA ALPHA CLAVAL COACT CDCOR COO
COL CDF ALD PR SFC PfC WA WF WOX AE
FUEl OXID I IER RIC DVDT VI PTl CMl
* • * * * * * • •. * * * * * * * * * * * * * * * * * * * * * 30. 50. 368. 599741. 1.
346751. 175654. 0.33 161. 6.79 15.00 0.2734 0.1815 001815 0.0354
0.1331 0.0130 1.51 1.000 0.983 1.00 3153. 95. -0. 3960.00
2743. -0.00 39.44 2.34 1072. 15.844 1.03
11. 5000. 659. 595884. 4. 350659. 77335. 0.60 444. 12.15
4.86 0.2584 0.0362 0.0362 0.0137 0.0106 0.0119 7.14 1.000 0.942
1.00 3062. 92. -0. 3960.00 6600. -0.00 138.58 1.99 1072.
15.599 1.03
lla. 15000. 846. 591966. 10. 314921. 96492. 0.80 535. 14.54
4.32 0.2128 0.0374 0.0374 0.0176 0.C080 0.0118 5.69 1.000 0.874
I.e 0 2542. 76. -0. 3960.00 10517 • -0.00 212.45 3.81 1072. 12.649 1.03
) 159. 25000. 915. 589135. 16. 257360. 92155. 0.90 446. 13.76
5.05 0.2556 0.0429 0.0429 0.0186 0.0121 0.0122 5.96 1.000 0.828
1.00 1947. 59. -0. 3960.00 1 '1'1L.Q • ..,J-rv. ~o.oo 21i.61 1.39 1072 •
9.236 1.03
238. 37500. 988. 585214. 28. 186923. 102484. 1.02 323. 6.46
6.71 0.3593 0.0659 0.0659 0.0286 0.0244 0.0129 5.45 1.000 0.785
1.00 1320. 41. -0. 3960.00 17269. -D.CO 111.19 1.03 963.
5.964 0.99
277. 39500. 1085. 583620. 34. 186040. 122438. 1.12 353. 2.38
6.40 0.3295 0.0719 0.0719 0.0381 0.0210 0.0127 4.58 1.000 0.779
1.03 1316. 40. -0. 3960.00 18863. -O.lO 44.97 2.18 933. 6.115 0.93
321. 41700. 1162. 581907. 42. 180147. 126334. 1.20 365. 2.34
6.47 0.3181 0.0718 0.0718 0.0381 0.0210 0.0127 4.43 1.000 0.774
1.04 1276. 39. -0. 3960.00
) 20576. -0.00 47.38 1.67 955.
6.085 . G.94 . . "
Table 12. Final Variable Sweep Configuration Data
114
)
)
TIME THRUST ALPHA
COL PFC
FUEL PH
ALTITUDE DRAG
CLAVAL COF WA
OX 101 lER eMl
VELOCITY MACH
COACT ALD
wF RIC
WEIGHT Q
eDCOR PR
iiOX OVOT
RANGE GAMMA
COO SFC
AE VI
* • • • * • * * * * * * * * * * * * * * * * * * * * * * * 417.
184536. 6.34
0.0174 1.07
24321. 6.768
49.7. 1'19817.
5.86 0.0140
1.12 27606.
8.281
560. 224250.
5.11 0.0108
1.13 30435. 10.678
611. 251 943.
4.52 0.C084
1. 14 33019. 14.066
707. 328381.
3.46 0.0045
1.14 39 1 93. 29.575
714. 411233.
2.84 O'()026
1. 15 45201. 62.089
832. 455881.
2.48 0.0017
1.15 51321.
122.936
45000. 133902. 0.2718 0.0124
1312. -0.00
1.03
47000. 144638. 0.2274 0.0120
1431. -0.00
1.40
48000. 158548. 0.1875 0.0114
1623. -O.LO
1.62
48700. 174112. 0.1562 0.0108
1855. -0.00
1.84
49500. 223212. 0.1024 0.0095 2644. -O.CO
2.36
50000. 255441.
O. C716 0.t083
3847. -0. (0
2.86
50400. 285632.
0.0535 O. t073
4010. -C.GO
3.33
1356. 1.40
0.0654 4.15
39. 34.23
1550. 1.60
0.0596 3.82
42. 26.28
1743. 1.80
0.0541 3.47
47. 17.43
1937. 2.CO
0.0498 3.14
54. 15.00
2421. 2.50
0.0424 2.41
75. 9.71
29C5. 3.00
0.0345
107. 9.18
3370. 3.48
0.0293 1.83 115. 8.48
578162. 424.
0.0654 0.999
-0. 2.01
514877. 504.
0.0596 0.978
-0. 2.55
572048. 6C8.
0.0541 0.960
-0. 3.38
569464. 726.
C.0498 0.941
-0. 4.15
563290. 1091.
0.U424 0.887
-0. 5.88
557283. 1534.
0.0345 0.827
-0. 8.89
551162. 2C26.
0.0293 0.827
-0. 9.86
Table 12 (cont'd). Final Variable Sweep Configuration Data
115
62. 1.45
0.0356 0.765
. 3960.00 1073.
82. 0.97
0.0336 0.761
3960.00 1412.
99. 0.57
0.0319 0.762
3960.0(' 1634.
114. 0.44
0.0306 0.769
3960.0C 1846.
149. 0.23
0.0285 0.818
396V.OO 2356.
118. 0.18
0.0236 0.934
3960.00 2854.
2G8. 0.14
0.0202 0.9(.6
396U.CO 3327.
TIME ALTITUDE VELOCITY WEIGHT RANGE THRUST DRAG MACH Q GAMMA ALPHA CLAVAL COACT CDCOR COO
COL CDF ALD PR SFC PFC WA WF WOX AE
FuEL OXIDIZER RIC OVDr VI PTl CM1
* * * * * * * * * * * • • • • * • • • * • * • • • • * * * 834. 50500. 3390. 550929. 209.
457940. 285872. 3.50 2039. 0.84 2.48 0.0531 0.0291 0.0291 0.0201
0.0017 0.0073 1~83 0.827 0.905 1.15 4026. 115. -0. 3960.00
51554. -0.00 49.45 9.58 3346. 125.877 3.35
886. 56400. 3874. 545140. 240. ~30358. 250942. 4.00 2009. 1.73
2.57 0.0530 0.0259 0.0259 0.0173 0.0018 0.0068 2.05 0.827 0.900
1.18 3770. 108. -0. 3960.00 57343. -0.00 11 7.23 9.62 3829.
188.879 3.82
950. 67500. 4363. 539181. 283. 314240. 180544. 4.50 1496. 1.94
3.19 0.0702 0.0250 0.0250 0.0151 0.0031 0.0069. 2.80 0.827 0.912
1.27 2793. 80. -0. 3960.00 63302. -0.00 147.63 6.90 4307.
211.325 4.23
) 1031. 76400. 4818. 533332. 345. 249552. 144211. 5.00 1212. 1.15
3.76 0.0852 0.0247 0.0247 0.0132 0.0047 0.0068 3.45 0.827 0.939
1.36 2210. 65. -Oe 3960.00 69151. -O.CO 98.20 5.11 4811.
?r;?~r;? 1.. ~'2
1132. 84700. 5396. 527267. 430. 202447. 120'760. 5.50 995. 0.77
4.37 0.1019 0.0252 0.0252 0.0117 0.0061 0.0068 4.04 0.827 0.977
1.47 1899. 55. -0. 3960.00 75216. -O.CiO 12.58 4.56 5316.
2'38.458 4.99
1255. ,91300. 5913. 520865. 545. 172450. 107073. 6.00 873. 0.47
4.80 0.1139 0.0255 0.0255 0.0104 0.0085 0.0065 4.48 0.827 1.021
1.58 1671. 49. -0. 3960.00 81618. -0.00 48.06 3.78 5823.
369.023 5.36
1278. 98570. 5942. 519873. 567. 137567. 102453. 6.00 626. 2.25
6.21 0.1582 - 0.0340 0.0340 0.0104 0.0157 0.0079 4.66 0.827 1.021
1.78 1332. 39. -0. 3960.00
} 82610. -0.00 233.08 0.92 5819.
/ 256.479 5.11
Table 12 (cont'd). Final Variable' Sweep Configuration Data
116
TIME AL Tl TUDE VElOClTY WEIGHT RANGE THRUST DRAG MACH Q GAMMA ALPHA CLAVAL COACT C l)COR COO
\ CDL CI)F ALD PR SFC J PFC WA wF wax AE
FUEl OXIDI ZER RIC DVDT VI PTl C Ml
* * * * * * • • • * • • * * * * * * * * * • * * * * * • '" 5199. 103192. 5942. • 415014. 4399.
103299. • 103299. • 6.00 626. • 6.25 • 0.1596 • 0.0343· 0.0343· 0.0104 * ·Start of
0.0160 • 0.<,019· 4.66· 1.000 • 0.963 • Cruise 1.78 • 1335. • 28 •• -0. 3960.00 *
187409. -0.00 -0.10 • 5818. * 256.188· 5.16 •
5318. 94000. 5430. 414216. 4511. 21782. 91879. 5.50 648. -0.83
5.15 0.1259 0.0294 0.0294 0.0111 0.0100 0.0071 4.28 1.000 1.C16
1.57 1295. 8. -0. 3960.00 1'88261. -0.00 -18.75 -4.52 5333. 192.275 4.90
5416. 93000. 4933. 413489. 4594. 26040. 94779. 5.00 561. -0.12
5.70 0.1464 0.0351 0.0351 0.0132 0.'Jl29 0.0t.89 4.17 1.000 0.971
1.57 1111. 7. -0. 3960.00 188994. -0.00 -10.63 -5.28 4825. 114.364 4.43
) 5506. 92000. 4437. 412883. 4664. 23855. 91412. 4.50 475. -0 .15
6.33 0.1736 0.0425 (,.0425 0.0151 0.0173 0.0102 4.G8 1.000 0.949
1.56 1043. 6. -0. 3960.00 189600. -O.tO -11.40 -5.65 4315.
e5.461 3.96
5591. 91000. 3941. 412364. 4722. 21170. 100231. 4.00 3 93. -0.16
1.09 0.2110 0.0529 0.0529 0.0173 0.0240 O. e1l5 3.99 1.000 0.936
1.55 911. 6. -0. 3960.00 190C99. -0.00 -12.25 -6.01 3804.
35.«;55 3.50
5610. 90000. 3446. 411983. 4710. 18022. 103580. 3.50 315. -0.22
6.01 0.2648 0.0682 0.0662 0.0201 0.0352 0.0129 3.88 1.000 0.941
1.53 779. 5. -0. 3960.00 190500. -0.00 -13.25 -6.55 3289.
18.914 3.02
5741. 86600. 2941. 411634. 4811. 15881. 105858. 3.00 271. -0.87
8.51 0.3<.·93 0.0810 0.0810 G.0236 0.0437 0.C131 3.82 1.000 0.968
1.46 7G9. 4~ -0. 3960.00 19C 849. -0. (0 -44.67 -6·54 2778.
) 10.121 2.58
Table 12 (cont'd). Final Variable Sweep Configuration Data
117
TIME ALTITUDE VElOCl TV WE IGHT RANGE THRUST DRAG MACH Q GAMMA ALPHA CLAVAL COACT C l)eOR COO -" COL COF AlD PR SFC )
PFC WA WF wax AE FUEL OXIDIZER RIC OVDT VI
PTl CM1
lit * • • • • • • * * • * lit lit * * * * • * • • * * * * * * * 5833. 78500. 2442. Itll050. 484q.
15773. 105083. 2.50 275. -2.19 8.02 u.3060 0.0794 0.0794 0.0285
0.0374 0.0135 3.85 0.907 1.363 1.34 425. 6. -0. 3960.00
191433. -0.00 -93.13 -5.17 2282. 7.352 2.17
5928. 69000. 1941. 410526. 4883. 14899. 99225. 2.00 275. -2.83
7.70 0.3059 0.0748 0.0748 0.0306 0.0308 0.0134 4.09 0.951 1.238
1.23 390. 5. -0. 3960.00 191957. -I).DO -95.92 -5.03 1779.
5.302 1.73
6039. 57500. 1453. 4C9949. 4914. 14089. 93778. 1.50 268. -3.85
7.58 0.3142 0.0726 0.0726 0.0345 0.0246 0.0135 4.33 0.990 1.350
1.13 412. 5. -0. 3960.GO 1<12534. -C.VO -97.52 -4.11 1254.
4.316 1.23
) 6187. 45000. 968. 409091. 4943. 10972. 72988. 1.00 211 ~ -4.02
7.26 0.3891 0.0699 0.0699 0.0268 0.0295 0.013 7 5.56 1.000 2.058
1.00 491. 6. -0. 3 960,,00 193392. _() . GO -67.91 -2.63 968.
• r;L 0 ~~
Jo. vv
6247. 40000. 872. 4U8689. 4952. 9442. 63035. 0.90 223. -4.94 6.91 0.3780 0.0587 0.0587 0.0186
0.0266 0.0135 6.44 1.000 2.741 1.00 573. 7. -0. 3960.00
193794. -0.00 -75.08 -1.45 968. 4.618 1.00
7247. 4GGOO. 872. 397088. 4952. 52947. 52947. 0·90 223. -J ... 9 4
.t; '" (~4· O.37'io6 0.0493 0.049:) O~C135 (;.C2('9 (:.C1'+9 7.60 1.0:';0 0.789
}"OO ~3 8 .. 12. -C .. 3960.0C 2'J,'j,Q5 & -0. C. C -75.08 -1.45 96F..
L\>.~6l8 1 ~ t 0
lS'4.:.r ~ 40GOQ. 872. 389124. 4952. 51£:759 51275. 0.<:;0 223. -4.9"
6&24- O~3603 0.0477 0.0477 0.0135 C~0193 0.0149 7.55 l.,.OOC 0.789
1. C 0 4334: 11 ~ -0. 3960.00 \ 2133594- -O.lO -75.08 -1. 4 5 968.
) 4,.618 l.00
Table 12 (c01:t'd), Final \T2.ria.b It, S\v~;ep Config-<.lration Data
: l~
~)
TIME AL TI TUDE VELOCITY WEIGHT RANGE THRUST DRAG MACH Q GAMMA ALPHA CLAVAL COACT COCOR COO
COL COF ALO PR SFC PFC WA WF WOX AE
FUEL OXIDIZER RIC OVOT VI PTl CMI
* * * * * * * * >0< * * * * * * * * * * * >0< * * * * * * * .. * 8090. 25000. 813. 387621. 4972. 9987. 66297. 0.80 352. -7.70 4.51 0.2256 0.0391 0.0391 0.0176
0.u090 0.0124 5.78 1.000 4.882 1.00 1003. 14. -0. 3960.00
214862. -0.(0 -109.00 -0.37 968. 8.325 1.00
8310. 50eO. 659. 382876. 4999. 9838. 65419. 0.60 444. -7.09
3.71 0.1773 0.0306 0.0306 0.0137 '0.0051 0.0118 5.79 1.000 10.811
1.00 1867. 30. -0. 3960.00 ) 219607. -0.(0 -81.28 -0.7C 968.
15.599 1.<'0
8498. 50. 244. 380549. 4999. t:!~"'" n ~,.." .,., 0.24 486. 4.00 ;)'00. :;;Joo::>.
3.71 0.1773 0.0306 0.0306 0.0137 0.0051 0.0118 5.79 1.000 1.668
1.00 586. 3. -0. 3960.00 221934. -0.00 o. -0.70 968. 15.274 1.(0
Table 12 (cont'd). Final Variable Sweep Configuration Data
)
119
)
AERODVNAMIC DEVICES WING + WING MOUNTED CONTROL SURFACES VERTICAL SURFACES HORIlONTAL SURFACES FAIRINGS, SHROUDS AND ASSOCIATED STRUCTURE
BODY STRUCTURE
64985. 9230. 9515. 1321.
STRUCTURAL FUELtOR COMMON BASIC ENCLOSiNG STRUCTURE PRESSURIZED COMPARTMENTS SECONDARY STRUCTURE
PROPELLANT! CONTAINER O. 71342.
5193.
INDUCED ENVIRONMENTAL PROTECTION COVER PANELS,NON-STRUCTURAL INSUL AT ION
LANDING GEAR
MAIN PROPULSION ENGINES AND ACCESSORIES A I R IN DUC T ION NACELLES,PODS,PYLONS,$UPPORTS FUELlOR COMMONlCONTAINERS AND SUPPORTS OXIDIZER CONTAINERS AND SUPPORTS pROPELLANT INSULATION FUEL SYSTEM OXIDIZER SYSTEM PRESSURIZATION SYSTEMS LUBRICATING SYSTEM
AEROOYNAMIC CONTROLS
PRIME POWER SOURCES ENGINE OR GAS GENERATOR UNITS POWER SOURCE TANKAGE AND SYSTEMS
POWER CONVER.SION AND DISTRIBUTION ELECTR ICAL HYDRAULIC/PNEUMATIC
GUIDANCE AND NAVIGATION
INSTRUMENTATION
COMMUNICATION
ENVIRONMENTAL CONTROLS
o.
19189. 12680.
40716. 16708.
3649. -0. -0. -0.
1 524. O.
4215. 160.
2305. 1205.
4037. 1223.
EQUIPMENT 201. PERSONNEL 2150. COOLANT SYSTEM O.
85117.
82535.
31868.
18192.
67031.
5812.
3510.
5259.
800.
404.
2025.
7625.
COMPARTMENT INSULAT ION 5274.
) ~------------------------------------------------------~ Table 12 (cont 'd). Final Variable &Weep Configuration Data
120
)
)
PERSONNEL PROVISIONS ACCOMMODATIONS FOR PERSONNEL FiXED LIFE SUPPORT FURNISHINGS AND CARGO HANDLING EMERGENCY EQUIPMENT
CREW STATION CONTROLS AND PANELS
DRY STRUCTURE
DESIGN RESERVE
PERSONNEL CREW,GEAR AND ACCESSORIES CREW LIFE SUPPORT
PAYLOAD CARGO PASSENGERS
USEFUL LOAD
RESIDUAL PROPELLANT AND SERVICE ITEMS TANK PRESSURIZATION GASES TRAPPED FUEL TRAPPED OXIDIZER SERVICE ITEMS RESIDUALS
RESERVE PROPELLANT AND SERVICE ITEMS FUEL-MAIN PROPULSION OXIDIZER-MAIN PROPULSION POWER'SOURCE PROPELLANTS LUBRICANTS
WET STRUCTURE
IN-FLIGHT LOSSES FUEL VENT OXIDIZER VENT POWER SOURCE PROPELLANTS LUBRICANTS
MAIN PROPELLANTS FUEL
TAKEOFF,CLIHB,ACCELERATE CRUISE DESCENT LO ITER LAND
OXIDIZER TAKEOFF,CLIMB,ACCELERATE CRUISE DESCENT LOITER LAND
TAKEOFF ~EIGHT
82610. 104198.
12633. 19565. 2328.
-0. -0. -0. -0. -0.
4700. 308.
8116. 305.
1250. 25.
13000. 35000.
230. 1205.
o. 120.
o. -0.
176. 83.
(
1122. -0.
3525. 330.
221911.
-0.
(
Table 12 (cont'd). Final Variable Sweep Configuration Data
121
13428.
300.
323906.)
o.
1215.
48000.
49275.
1556.
259.
374996. ,
5511.
221911.
602483.'
)
VOLUMES BODY STRUCTURE CREW AND PASSENGER COMPARTMENTS CARGO COMPAR TMENT LAND ING GE AR BAY S PROPULSION BAY WITHIN BODY WING CENTER SECT ION FUELlOR COMMON PROPELLANT) CONTAINER OXIDIZER CONTAINER FUEL(OR COMMON PROPELLANT) INSULATION OTHER BODY VOLUME -
TOTAL BODY VOLUME
WETTED AREAS GROSS BODY LOWER SURFACESfTHERMAL PROTECTION) UPPER SURFACES(THERMAL PROTECTION) PERSONNEL COMPARTMENTS CARGO COMPARTMENTS
PLAN AREAS WINGIOR LIFTING SURFACE.eGROSS. EXPOSED WING AREA BODY MAXIMUM CROSS SECTION BASE VERT I CAL SUR FACE S HORIZONTAL SURFACES AIR INLET CAPTURE AREA
UNl' Wt:lbMI;)
WING VERTICAL SURFACES HORIZONTAL SURFACES BODY S TRUCTUR E e BAS IC ) LIFTING SURFACE MAXIMUM LOADING
DIMENSIONAL DATA WING
STRUCTURAL SPAN ROOT CHORD LENGTH THICKNESS RATIO
BOOY LENGTH WIDTH HEIGHT
ENGINE SCALE fACTOR
CU FT 10207.
8645. 903. 270.
o. 1475.
52377. -0. -0.
13037.
86913.
SQ. FT. 20413.75
5115.85 14691.90 4831.66
442.80
SQ. FT. 4819.81 3161.51
511.59 587.92
1108.64 1214.61
111,,39
l..B/SQFT 13.48 8.33 7.88 3.19
125.00
fEET
180.51 86.18 0.05
318.81 23.41 31.35
Tab Ie 12 (cont I d) . Final Variable Sweep Configuration Data
122
CU M 289. 245. 26.
8. o.
42. 1482.
-0. -0.
369.
2460.
SQ. M. 1896.44 531.00
1365.43 448.86
41.14
SQ. M. 441.71 293.11
53.10 54.62
102.99 112.84
10 .. 35
KG/SGM 0.61 0.38 0.36 0.17 5.61
METERS
55.02 26.45
91.11 7.15 9.56
8.0 BLENDED BODY CONFIGURATION
8.1 BACKGROUND
The basic feature of this configuration is self descriptive in that the wing is not discrete from the body. This blending of the wing and the body is motivated by both structural and aerodynamic considerations. Blending the wing/body intersection provides a better ratio of volume/lifting planfonn which in turn reduces the wetted areas which results in lower friction drag and '-lower structure weight. In addition, this blending can benefit the flow field for the engine inlets and minimize the wing/body interference drag. The basis for a configuration of this type was provided in the detailed Aerospaceplane studies conducted by Convair in 1964.
8.2 GROUND RULES
) In order to make a design study of the blended body concept and make it comparable with the delta wing and variable sweep configurations the ground rules shown in Table 13 were est,ablished. These are quite similar to the ground rules shown in Table 3 for the delta wing configuration.
)
Range = 5,000 n.mi.
Cruise Mach = 6.0
200 Passengers
5,000 Lb. Cargo + 40 Lb/Passenger
Selected Trajectory (Figure 20)
Fuel Rich Turbofan, STFRJ -230A
Subsonic Loiter = 1, 000 Seconds
Su~g,nie Dn", ~ .........., 100 n. mi.
~.~ ~ Integral Tanks. -
TABLE 1l. GROUND RULES FOR BLENDED BODY CONFIGURATION STUDY
123
, ) 8. 3 CONFIGURATION CONCEPTS
A review of the previous Aerospaceplane studies showed that with a 80 0 swept leading edge, an aspect ratio of .46 and a take-off wing loading of 64 lb/ft2 the low speed handling qualities were marginal. Both the longitudinal short period dynamics and the Dutch roll characteristics fell within the "acceptable for emergency operation" limits as established in Report No. TC-1332-F-1 "Handling Qualities Requirements for Hypervelocity Aircraft". These handling qualities are unacceptable for commercial operations.
Two alternate approaches were therefore adopted that would provide more acceptable handling characteristics: (a) The wing planform was modified to a double delta, or (b) var~able sweep wing panels were extended for landing and take-off. (a) vs. (b) is the classically different approach of providing either fixed or moving aerodynamic surfaces in order to get adequate stability.
A series of brief designs were therefore made in order to study these two approaches. The most significant of these are included in Figure 69 which shows different approaches to blended body configurations. The three configurations in the top row show different locations of the passenger compartment. The top location shown in the left hand column was selected as having the best compromise between cabin structure weight, cabin insulation, volume utilization, engine noise, passenger appeal (seating layout) and c. g. travel.
The left hand column in Figure 69 shows different vehicle shapes. The upper three show a low, mid and high "wing" layout. The low wing layout was eliminated because the flat lowel surface does not provide the aero/propulsIOn advantage mherent in the blended body concept. Compared with the high wing layout, the mid wing layout results in a better distribution of engines around the aft body, a wider tread and shorter length for the main landing gear and better structural blending. The lower configurations of Figure 69 show variable wing sweep approaches. The upper of these tW? shows a 80
0 delta wing with overlapping wing panels. The lower configuration shows an attempt at eliminating the duplicated structure of the overlapping wing panels, however, this resulted in a configuration that required more total wing area than the overlapping wing. In addition, the moving wing and the aft control surface almost violate the principles of a blended body without giving any obvious advantages.
It was therefore decided that two blended body configurations would be evaluated, the mid wing, double delta and the blended body with overlapping, variable sweep wing panels.
8.4 DOUBLE DELTA CONFIGURATION
) This design is shown in Figure 70. As shown in the cross ,sections, most of the blending of the body and the wing is on the lower surface. The upper surface has less blending
124
Double Delta Low Wing
Double Delta Mid Wing
Double Delta High Wing
Variable Sweep Overlapping Wing Panels
Variable Sweep Aft Wing
Panels
TOP
\, '-
-----.l
PASSENGER COMPARMENT LOCATION
FOR\\'ARD CENTER
"
Figure 69. Blended Body Configuration Concepts
125
)
< j t
! --.1 A -~- .... ! ~-I - i
i - _ --tIl-:__ : - - !
~-~!
Figure 70. Blended Body - Double Delta Configuration
126 -~
/ N'4TANIC
!
I ! I
L..--__________ : __ ~
~ ;
-c--------------oS'" \
~C;V
~ ',280:
Total volume
Fuel n~lumt.' . Landing gear main
Eng:mes type
Thrust
Passengers @. 6 abre3...'it
@ 4 abreast
Seat spacing
I L_~ __
1..-,
...Ind f, ,
0&
_0
28() it.
24.:) ft
lOG. 8:)f) ('U ft
61. lOt; ('u ft
37 fL
Pratt &: V.:litnc)" STFnS.;!,;{t.\{;.,
;5,00.0 tb SL.S each
24 rows
14 ro'\\"S
.36 in.
9t}Ocu. ft.
130' _______ .. ______ _
I
with the passenger compartment mounted on top of the blended upper surface so that the passenger floor provides the spanwise structural continuity. Five ~on-circular fuel tanks dominate the body volume. The structural concept iS,a nickel :-llloy hot structure with a nickel alloy tank liners insulated from the outer structure. The four turboramjets are mounted around the lower body and exhaust along the aft body. The main landing gear is mounted outboard of the engines and retracts forward. The droop nose provides adequate visibility during landing.
8.5 VARIABLE SWEEP, OVERLAPPING WING CONFIGURATION
This is shown in Figure 71. This configuration allows the vehicle planform to be dictated by transonic, supersonic and hypersonic drag considerations and to be independent of the take-off and landing requirements. The variable sweep wing panels provide the desired low speed performance and handling qualities. The wing area of the variable sweep configuration (which is less than the double delta) is dictated largely by the span of the variable sweep panels, the location of the pivot and the area of the elevons.
8. 6 AERODYNAMICS
The aerodynamic characteristics of the two blended body configurations were anfllyzed at several Mach numbers. Supersonic lift and drag ~1{~ estimated by breaking the total vehicle into components whose contributions were linearly superimposed; these components were:
1) upper surface - wing
2) lower surface forebody - blended wing-body
3) lower surface afterbody - boat tail
4) propulsion package and base
5) aft wing section - lower surface
6) vertical tail
Pressure distributions over each lift-contributing component were estimated using tangent-cone and shock-expansion methods, where applicable, over a range of anglesof-attack, with corrections to the integrated lifts and drarrs based on correlations with experimental data such as that of References 21 through 23. The subsonic characteristics of each configuration were analyzed using correlations with experimental da:a of low aspect ratio wing-body combinations, such as Reference 11, with corrections to account for the contributions oithe double-delta planform of the baselin~ configur'atio:: and the variable-s\'I"eep panel of the alternate configuration (with the panel in the forward
127
,
•
\ /
-7"-- __ _
/ ~O·
I"nt> OF V1"'j'
Figure 71.
I ~ ,- .. \
~/
I COHPT
------1
\
I
<?oo ""SS€NGE<? ... ---~ - ------- - -----t---
------- -"------------------
--------~----
Blended Body - Variable Sweep Configuration
128
/ /'
/ x
----I··~~ !
-NOSTANK
(§) ~' I ---r
--_._-
f)~'l't (m.L"-)
Fu\!J \'O!Wl1c
Thrust
Cargo 0;..1\ Yvhune
/'iIi, lo
72v sq. i1.
13,1.11 sq It .
. H
2lSt.i II
2.t.::; ft.
lO-l. OO(J CU. it
.39 •. ')00 t."U. ft.
{our 66:dS tire:,.; I!t~l' "leo
twin 1 ~x 12 u l'{,.~
3i ft.
Pratt & Whitne.\ STJ'lC::i-23UA (../.)
•. ).000 lb. SUi C'H:!,
2-1 rows
14 rows
36 in.
900 cu. ft.
l--S7' ___ __
--------7(!j'
-------------- - 148' SPAN AT 2'0·
~
I
)
)
, position). The resulting lift, drag and LID characteristics of each configuration are
shown in Figures 72 and 74. Figures 73 and 75 show the breakdown of the drag coefficient at a lift coefficient of .10 for each configuration.
The longitudinal stability characteristics of the biended-bodies were examined; ~~e configurations were found to ~e statically stable throughout the flight regime. The trim ~apability of the baseline configuration was determined at M = 4.5 and is presented in Figure .76. It is seen that the elevons have sufficient power to trim the vehicle at an angle-of-attack near (LID) m~ximum.
8.·7 -WEIGHTS ...
Based on data generated under contract NAS 2-3025, weights for the two selected blended body-...configurations were determined. These are shown in Tables 14 and 15 for the double delta and variable sweep versions respectively.
8.8 DoUBLE DELTA - GEOMETRY VARIA TIONS
....
The effects- of planform loadIng on the double delta-blended body were investigated using the synthesis program. Plan loading was increased by essentially increasing the thickness ratio of the configuration. The' resulting effects on take-off weight are shown in, Figure 77. This configuration differs from the delta and variable sweep configurations in that it has an optimum planform loading. This bucket results from the trade-
, off between structural weight and aerodynamic eff,iciency, lift-to-drag ratio. The effects of planform loading on cruise LID and body structural weight are also shown in Figure 77. As the planform loading is increased, the lift-to-drag r,atio decreases since the configuration effective thickness ratio is increasing. Also, as the planform loading is increased, the wetted area of the configuration decreases resulting in a decrease in body structural weight.
As with the ollier configurations, take'-off velocity considerations dictate an upper limit on planform loading. The subsonic lift characteristics and the adaptability of high lift devices is rather limited on the double delta blended body. For the take-off computations, a take-off angle of attack of 15 degrees was assumed and a take-off lift . coefficient of .46 was used. Incorporating the ten percent take-off velocity margin, a maximum allowable take-off wing loading of about 35 was obtained and is shown on , Figure 77.
8,.9 FINA L DOUBLE DE L TA CONFIGURA TION
A synthesis run was made for a final double delta blended body configuration having a planform loading of 35 psf. The performance· and weight/volume characteristics of this configuration are sho'WIl in Table 16.
129
)
)
8.10 VARIABLE SWEEP - WING GEOMETRY VARIATIONS
The effects of wing area changes for the variable sweep panels on the take-off weight of this blended body configuration were analyzed using the synthesis and performance program. The results in the form of take-off weight as a function of wing loading are shown in Figure 78. The wing loading as used on Figure 78 is defined as the take-off weight divided by the area of the variable sweep panels extended to the centerline. Increasing the wing loading decreases the take-off weight due to the reduction i:l wing area. The take-off weight characteristic,s of the blended body with variable sweep panels are much higher than the double delta blended body due to the fact that the vehicle is carrying a large block of inert weight that is only used a small portion of the time. The sensitivity of take-off weight to inert weight changes is rather large which causes the weight to spiral up rapidly with the addition of the extra wing pane Is.
Take-off velocity considerations dictate an upper limit of 130 psf on take-off wing loading. This limit line is shown in Figure 78.
8.11 FINAL VARIABLE SWEEP WING CONFIGURATION
A synthesis run was made for a final variable sweep wing configuration having a wing loading of 130 psf. The performance and weight/volume characteristics for this final configuration are shown in Table 17.
130
)
)
.25
.20 CL .15
.05 __ _
11.200
o 2 4 6 8
o 2
Figure 72.
10 12 o
4 6
.01
8
.02 .03 .ot .05
CD
'" IV
Double Delta Aerodynamic Characteristics
131
.06 .07
.03
) .01-. II
o 1 2 3 Mach Number
6C Dlower surface afterbody pressure
6CD upper surface + nose + vertical
_tail wave
~ IoIi 6CDlower surface forebody pressure
6CD l~-: friction
4 5 6
Figure 73. Double Delta - Drag Breakdown
132
.35 -')
.30
.25
.20 .20 C
L CL .15 .15
.10 .10
.05 .05
0 0 2 4 6 8 10 12 0 .01 .02 .03 .04 .05 .06 .07
(Y.o
) A = 20° .5 LE
6
5
LID 4.5 4 3.5 2.5
~E 80° 3 :=
2
1
o 2 4 6 8 10
) Figure 74. Variable Sweep - Aerodynamic Characteristics
133
.04
.03
.01 )
0" o
"- ) "~
~ - "
,~~~ ~:~=~~=l~~:-~"~i-~:~~,~L~~ ~~ __ I 2 3 4 5
::vlach Number
t,CDlower surface forebody + nose
,-+ t,C : :~t::1 Dfriction
6
Figure 75. Variable Sweep - Drag Breakdown
134
)
Figure 76. Double Delta - Longitudinal Stability
)
135
)
)
AERODYNAMIC DEVICES WING + WING MOUNTED CONTROL SURFACES VERTICAL SURFACES HORIZONTAL SURFACES FAIRINGS,SHROUDS AND ASSOCIATED STRUCTURE
BODY STRUCTURE
o. 7684.
o. 1245.
STRUCTURAL FUEL{OR COMMO~ BASIC ENCLOSING STRUCTURE PRESSURIZED COMPARTMENTS SECONDARY STRUCTURE
PROPELLANT. CONTAINER o. 133442.
5193. o.
INCUCED ENVIRONMENTAL PROTECTION COVER PANELS,NON-STRUCTURAL INSUL ATION
LAND ING GEAR
MAIN PROPULS [ON ENGINES AND ACCESSORIES AIR INDUCTION NACELLES,PODS,PYLONS,SUPPORTS FUEL(OR COMMON)CONTAINERS AND SUPPORTS OXIDIZER CONTAINERS AND SUPPORTS PROPELLANT INSULATION FUEL SYSTEM OXIDIZER SYSTEM PRESSURIZATION SYSTEMS LUBRICATING SYSTEM
AERODYNAMIC CONTROLS
nn .... r rr. 11:0 cnllorc::c '. ''-.~ ,~ ~ ... ~ --~ ENGINE OR GAS GENERATOR UNITS POWER SOURCE TANKAGE AND SYSTEMS
POWER CONVERSION AND DISTRIBUTION ELECTRICAL HYDRAULIC/PNEUMATIC
GUIDANCE AND NAVIGATION
INSTRUMENTATION
COMMUNICATION
ENVIRONMENTAL CONTROLS EQUIPMENT PERSONNEL COOLANT SYSTEM COMPARTMENT INSULATION
o. o.
53813. 22500.
4259. o. o.
11731. 1556.
O. 4820.
160.
2368. 1268.
424B. 1239.
210. 2150.
o. 5214.
8929.
138635.
o.
19063.
98844.
6093.
5487.
800.
511.
2025.
7635.
Table 14. Blended Body, Double Delta, Baseline Weight and Sizing Data
138
)
)
PERSONNEL PROVISIONS ACCOMMODATIONS FOR PERSONNEL FIXED LIFE SUPPORT FURNISHINGS AND CARGO HA~DlING EMERGENCY EQUIPMENT
CREW STATION CONTROLS AND PA~ElS
DRY STRUCTURE
DESIGN RESERVE
PERSONNEL CREW,GEAR AND ACCESSORIES CREW LIFE SUPPORT
PAYLOAD CARGO PASSENGERS
USEFUL LOAD
RESIOUA( PROPELLANT AND SERVICE ITEMS TANK PRESSURIZATION GASES TRAPPED FUEL TRAPPED OXIDIZER SERVICE ITEMS RESIDUALS
RESERVE PROPELLANT AND SERVICE ITEMS FUEL-MAIN PRUPULSION OXIDIZER-MAIN PROPULSION POWER SOURCE PROPELLANTS LUBR ICANTS
WET STRUCTURE
... ..... '" L""- ... ..,-=-::0-
FUEL VENT OXIDIZER VENT POWER SOURCE PROPELLANTS LUBRICANTS
MAIN PROPELLANTS FUEL
TAKEOFF,CLIMB,ACCELERATE CRUISE DESCENT LOITER LAND
OXIDIZER TAKEOFF,CLIMB,ACCELERATE CRUISE DESCENT LOITER LAND
TAKEOFF WEIGHT
-0. -0. -0. -0. -0.
-0. -0. -0. -0. -0.
4700. 308.
8116. 305.
1250. 25.
13000. 35000.
282. 1268.
o. 127.
o. o.
176. 83.
2036. o.
3512. 330.
271600.
o.
13428.
300.
305388 ••
o.
1275.
48000.
49275.
1677.
258.
356598. ,
r;;A77
271600.
634075.)
Table 14 (cont 'd). Blended Body, Double Delta, Baseline Weight and Sizing Data
137
)
)
VOLUMES BODY STRUCTURE CREW AND PASSENGER COMPARTMENTS CARGO COMPARTMEN T LANDING GEAR BAYS PROPULSION BAY WITHIN BODY WING CENTER SECTION FUEL(OR COMMON PROPELLANT) CONTAINER OXIDIZER CONTAINER FUEL(OR COMMON PROPELLANT) INSULATION OTHER BODY VOLUME
TOTAL BODY VOLUME
WETTED AREAS GROSS BODY LOWER SURFACES{THERMAL PROTECTION) UPPER SURFACES(THERMAl PROTECTION) PERSONNEL COMPARTMENTS CARGO COMPARTMENTS
PLAN AREAS WING(OR LIFTING SURFACE) (GROSS) EXPOSED wING AREA BODY MAXIMUM CROSS SECTION BASE VERTICAL SURFACES HORIZONTAL SURFACES AIR INLET CAPTURE AREA
UNIT WEIGHTS WING VERTICAL SURFACES HORIlONTAL SURFACES BODY STRUCTURE (BASIC) LIFTING SURFACE MAXIMUM LOADING
DIMENSIONAL DATA WING
STRUCTURAL SPAN ROOT CHORD LENGTH THICKNESS RATIO
BODY LENGTH WIDTH hEIGHT
ENGINE SCALE FACTOR
cu FT CU M 15560. 440.
8645. 245. 903. 26.
1902. 54. o. o. O. o.
64040. 1812. o. o.
3762. 106. 17925. 507.
112737.
SQ. FT. 31120.29
o. c.
4831.66 442.80
SQ. FT. o.
-0. 830.91 111.16
1318.59 o.
150.00
lB/SQFT o. 5.83 o. 4.29
56.60
fEET
o. o. 0.05
287.81 o.
26.95
7.00
3190.
SQ. M. 2891.07
o. o.
448.86 41.14
SQ. M. o.
-0. 77.19 15.90
122.50 o.
13.93
KG/SGM o. 0.26 o. 0.19 2.57
METERS
c. o. J
87.72 O. 8.21
Table 14 (cont 'd). Blended Body, Double Delta, Baseline Weight and Sizing Data
138
)
)
AERODYNAMIC DEVICES WING + ~ING MOUNTED CONTKOL SURFACES VERTICAL SURFACES HURIZCNTAl SURFACES FAIRINGS,SHROUDS AND ASSOCIATED STRUCTURE
BODY STRUCTURE
45468. 8663.
o. 1074.
STRUCTURAL FUEL(OR CGMMON BASIC ENCLUSING STRUCTUKE PRESSURIZED COMPARTMENTS SECONCARY STRUCTURE
PROPELLANT) CONTAINER o.
INCUCEO ENVIRONMENTAL PROTECTION C0VER PANELS,NON-STRUCTURAL INSULATICN
LANDING GEAR
MAIN PROPULSION ENGINES AND ACCESSORIES AIR INDUCTION NACELLES,PODS,PYLONS,SUPPORTS FUEL(OR COMMON)CGNTAINERS AND SUPPORTS OXIDIZER C~NTAINERS AND SUPPORTS PROPELLANT I~SULATION
FUEL SYS TEM OXIDIZER SYSTEM PRESSURIZATION SYSTEMS LUSRICATING SYSTEM
AERODYNAMIC ceNTROLS
PRIME POwER SOURCES ENGINE OR GAS GENERATOR UNITS POwER SOURCE TANKAGE AND SYSTEMS
POWER CONVERSION AND DISTRIBUTION E: L E C T RIC AL HYDRALliC/PNEUMATIC
GUIDANCE ANO NAVIGATION
INSTRUMENTATION
COMMUN I CA T lU N
EN~IRONMENTAL CC~TROLS
EQUIPMENT PERSONNEL COOLANT SYSTEM COMPARTMENT INSULATION
127093. 5193.
o.
o. O.
53813. 22500. 4259.
o. o.
12036. 1662.
O. 4943.
160.
2476. 1376.
4609. 1318.
226. 2150.
O. 5274.
Table 15. Blended Body. Variable Sweep Baseline Weight and Sizing Data
139
55205.
132286.
o.
20547.
99373.
6572.
5926.
800.
469.
2025.
7651.
)
)
PERSONNEL PROVISIONS ACCOMMOCATIO~$ fOR PERSONN~l fIXED LIfE SUPPORT FURNISHINGS AND CARGO HANDLING EMERGENCY EQUIPMENT
CREW STATION CONTROLS AND PANELS
DRY STRUCTURE
DESIGN RESeRVE
PERSOf~N EL CREW,GEAR ANC ACCESSORIES CRE~ lIFE SUPPOR1
PAYLOAD CARGu pAS S E NGERS ft.fl/U
USEFUL LOAD
RESIDUAL PROPELLANT AND SERVICE ITEMS TANK PRESSUR.IlATION GASES TRAPP f:C FUEL TRAPPED OXIDIZER SERvICE- ITEMS RESIUUAlS
RESERVE PROPELLA~T AND SERVICE ITEMS FUEL-MAIN PRCPULSION UXluIZcR-MAI~ PROPuLSION POWER SUuRCE PROPELLANTS LUBRICANTS
wET STRUCTUKE
IN-r-LlbHI LU~St:)
FUEL VENT OXIDIZER VENT POWER SOURCE PROPELLANTS LUBR.ILANTS
MAIN PROPELLANTS FUEL
TAKEOFF,CLIMB,ACCELERATE CRUISE CESCE:NT LOITER LANC
OXIDIZER TAKEOFF,CLIMB,ACCELERATE CRUISE GESCENT LOITER,-LANe
TAKEOFF WEIGHT
-0. -0. -0. -0. -0.
o. -0. -0. -0. -0. -0.
687841.)
Table 15 (cont'd). Blended Body, Variable Sweep Baseline Weight and Sizing Data
140
)
)
VOLUMES BODY S TKU C TUR E CREW ANC PASSENGER CCMPARTMENTS CARGO COMPAKTMENT LANDING GEAR dAYS PROPULSluN BAY WITHIN BODY WING CENTER SECTION FUEL(OR CCMMGN PROPELLANT} CONTAiNER OXIDIZER CONTAINER FUEL(OR CGMMON PRUPELLANT) INSULATION OTHER ~ODY VOLUME
TOTAL BOCY VOLUME
WETT EC AR EAS GROSS t30DY 00WER SURFACES(THERMAL PROTECTIDN) UPPER SURFACES(T~EkMAL PROTECTION) PERSONNEL COMPARTMENTS CARGU COMPARTMENTS
PLAN AREAS WING(UK LIFTING SURFACE)(GRUSS) EXPOSED ~ING AREA ~OOy MAXIMU~ CRUSS SECTION BASE VERTICAL SURFACES HORIlCNTAl SURFACES AIR INLET CAPTURE AREA
UNIT WEIGHTS WIN\;> VERTICAL SURFACES HORllUNTAL SURFACES BUD~ STRUCTURE teASIC) LIFTING SURFACE MAXIMUM LOAUING
DIMENSIONAL CATA kING
STRUCTURAL SPAN ROUT CHORD LENGTH THICKNESS kATIO
BUOY L ENGT H wIDTH HEIGHT
ENGINE SCALE FACTOR
CU FT 13429. 8645.
903. 2064.
o. 1079.
66506. O.
3858. 14290.
110774.
SQ. FT. 26d58.30
O. O.
4831.66 442.80
SQ. F T'. 3598.31 3598.31
795.01 147. 72
1303.23 O.
150.00
LB/SQFT . ~ 1.':'.0
6.65 o. 4.73
76.40
FEET
145.17 35.15
0.08
286.13 O.
26.19
7.00
CU M 380. 245. 26. 5 R. o.
31. 1882.
o. 109. 404.
3135.
SQ. M. 2495.14
o. o.
448.136 41.14
SQ. ~.
334.28 334.28 73.86 13.72
121.07 o.
13.93
KGI SGM ,.." """"' v • ..-
0.30 o. 0.21 3.47
METERS
44.25 10.71
'L~ •
87.21 O. 8.16
Table 15 (cont'd). Blended Body, Variable Sweep Baseline Weight and Sizing Data
141
mm Take-off Speed = 160 kts. 700
)
:0 -0 0
600 0 M
..., ..c b.O .-CJ.)
~ '-'-0 500 I CJ.)
~ E-!
Planform Loading (psf)
) Figure 77. Double Delta - Planform Loading Effects
142
)
)
AERODYNAMIC DeVICES WING + WING MOUNTEO CONTROL SURFACES VERTICAL SURFACES HORIZONTAL SURFACES FAIRINGS,SHROUDS AND ASSOCIATED STRUCTURE
BODY S TRue TUR E
, o. 5775.
o. 1389.
STRUCTURAL FUEL tOR COMMO~ PROPELLANT) CONTAINER o. BASIC ENCLOSING STRUCTURE PRESSURIZED CCMPARTMENTS SECONDARY STRUCTURE
INDUCED ENVIRUNMENTAl PRUTECTION COVER PANELS,NON-STRUCTURAl INSULATIUN
LANDING GEAR
MAIN PROPULSION ENGINES AND ACCESSORIES AIR INDUCTION NACEllES,PODS,PYLONS,SUPPORTS FUELIOR CCMMONlCONTAINERS AND SUPPORTS OXIDIZEK CONTAINERS ANO SUPPORTS PROPELLANT INSULATION FUEL SYSTEM OXIDIZER SYSTEM PRESSURIZATION SYSTEMS LUBRICATING SYSTEM
AERODYNAMIC ceNTROlS
nnT I: on ~.D Cr1IIDrl:c . £ - ~~, ~ ~ ~
ENGINE OR GAS GeNERATOR GNITS POWER SJURCE TANKAGE ANO SYSTEMS
POWER CONVERSION ANO DISTRIBUTION ELECTRICAL HYDRAULIC/PNEUMATIC
GUICANCE AND NAVIGATION
INSTRUMENTATILN
COMMUN lCAT ION
tNVIRONMENTAL CGNTRGLS EQUIP;'1ENT PERSONNEL COOLANT SYSTEM CCMPARTMEN{ INSUlATIGN
125213. 5193-r-
c.
U. C.
44374. 24200.
3817. n ' ~ ...
-(,. 12233. ~--1363.
o. 4001.
160.
2188. 1lH38.
3643. 1091.
183. 215(, •
o. 5274.
Table 16. Final Blended Body, Double Delta Configuration Data
143
o.
16j 5 72.
"----9C148.
5290.
3275 •
4735.
80D.
547.
2025.
76C>3.
PERSONNEL PROVISIONS ACCOMMODATIONS FOR PERSONNEL FIXED LIFE SUPPORT FURNISHINGS AND CARGO HA~DLING
) EMERGENCY EQUIPMENT
)
)
CREW STATION CONTROLS AND PANELS
DRY STRUCTURE
DESIGN RESERVE
PERSONNEL CREW,GEAR AND ACCESSORIES CREw LIFE SUPPORT
PAYLOAD CARGO ---..". PASSENGERS
USEFUL. LOAD
RESIDUAL. PROPELLANT AND SERVICE ITEMS TArt"r; PRESSURIZATION GASES TRAPPEC FUEL TRAPPED OXIDIZER SERVICE ITEMS RESIDUALS
RESERVE PROPELLANT AND StRVICE ITEMS FUEL-~AIN PROPULSION o X I D IZ E R - MA I N P f<. 0 P U LSI 0 N POWER SOURCE PROPELLANTS LUBRICANTS
WET S TRUCTUR. E
IN-FLIGHT LOSSlS FUEL VENT OXIDIZER VI:NT POwER SOURCE PROPELLANTS LUBRICANTS
MAIN PROPELLANTS FUEL
TAKEOFF,CLIMB,ACCELERATE CRUISE DESCENT LOIT ER LAND
OXIDIZER t~ TAKEOFF,CLIMB,ACCELERATE CRLJISE DESCENT LOITER LAND
TAKEOFF WEIGHT
'7,3 74167.
0U8 82) 10953. 2239(".
5852. :z-05't4-1
-0. -0. -0. -C. -0.
13428. 4700.
308. 811Q.~
3U5.
300.
282298. ) .... ;,
o.
1275. 1250.
25.
48000. 13000. 35000.
49275.
1410. 213.
1088. ' .... :'1>.
o. 109.
250. o.
-0. 167.
83.
::<::<:l.?:l.? _ 1 -,..J-I'--''''''_ ..
5224. 1550.
-u. 3345.
330.
2(; 5340. 205340.
-0.
543797.}
Table 16 (cont'd). Final Blended Body, Double I;>elta Configuration Data
144
)
)
VOLUMES BODY STRUCTUR E CREW AND PASSENGER COMPARTMENTS CARGO COMPARTMENT LANDING GEAR bAYS PROPULSION BAY WITHIN BODY WING CENTER SECTION FUEL(OR COMMON PROPELLANT) CONTAINER OXIDIZER CONTAINER FUEL(OR COMMON PRUPELLANT) INSULATION OTHER BODY VOLUME
TOTAL BODY VOLUME
WETTEC AREAS GROSS BODY LOI-lER SURFACES« THERMAL PROTECTION) UPPER SURFACESfTHERMAL PROTECTION) PERSONNEL COMPARTMENTS CARGO COMPARTMENTS
PLAN AREAS WING(OR LIFTING SURFACE) (GROSS) EXPOSED WING AREA BOOY MAXIMUM CROSS SECTION BASE VERTICAL SURFACES HORIZONTAL SURFACES AIR INLET CAPTURE AREA
UNIT WEIGHTS WING VERTICAL SURFACES HORIlONTAL SURFACES BODY STRUCTURE (BASIC) LIFTING SURFACE MAXIMUM
DIMENSIONAL DATA WING
S TRUC TURAl SPAN ROOT CHuRD LENGTH THICKNESS RA T 10
BODY LENGTH WIDTH hEIGHT
ENGINE SCALE FACTOR
LOADING
CU FT CU M 17359. 491'.
8645. 245. 903. 26.
1631. 46. o. O. O. o.
48449,., 1371. ":"0. -0.
3921. 111. 18244. 516.
99152.
SQ. FT. 34717.53
o. c.
4831.66 442.80
SQ. FT. 15423.46 15423.46
756.84 190.95
1210.41 o.
161.34
o. 4.77 o. 3.61
35.25
FEET
o. o. 0.05
275.75 o.
25.82
5.77
2806.
SQ. M. 3225.26
o. o.
448.86 41.14
SQ. M. 1432.84 1432.84
70.31 17.74
112.45 o.
14.99
o. 0.22 o. 0.16 1.60
METERS
o. o.
8 • 5 o. 7.87
Table 16 (oont'd). Final Blended Body, Double Delta Configuration Data
145
T U~E ALTITUOE VELOCITY WEIGHT RANGE THRUST CRAG MACH Q GAMMA ALPHA CLAVAL COACT. coeOR CDO
CDL COF AlD PR SFC
1 PFC wA i'lF wax AE J FUEL OXIDIZER RIC DVOT VI
PTl CMl
* * * * * * * * * * * * * * * * * * * * * * * * * * * * * * 19. 50. 3C5. 541916. 1.
372576. 71327. 0.27 110. 8.39 15.00 C.0793 0.0335 C.0335 0.0020
C.0227 0.t088 2.37 1.voe 0.980 1.00 3375. 101. -0. 3960.60
1881. -0. (0 47.BO 3.76 5944. 15.478 5.43
5 d. 5tOC. 659. 537930. 4. 3816Ul. 1(;0482. 0.60 444. 12.39
3.05 C.('738 O.Oi47 C.0147 0.u053 C.OOIC, 0.OC84 5.U 1.GOO 0.942
1.CO 33~2. leG. -0. 3960.0(; 5867. -G.00 141.29 9.91 5944.
15.599 5.43
101. 150li0. 846. 534049. 9. 342709. 133436. 0.80 535. 15.46
2.50 0.(604 C.0162 0.0162 C.(j(j69 0.0008 0.L085 3.74 '1.000 0.874
1.vu 2766. 63. -0. 3960.00 9748. - (j. U) 225.62 4.04 5944.
12.649 5.43
) 141. 2500u. 915. 531083. 14.
280069. 128628. 0.90 446. 14.0C 2.95 0.(,727 0.0187 0.0187 0.u085
(;.\.1014 0.(;(;88 3.89 1.000 0.828 1.( 0 2119. 64. -G. 3960.CC
12714. -l..'.0v 2<:: 1. 28 1.41 5944. 9.2~o 5.43
214. 37500. '::/tlts. ?L/V':JU. L?
203415. 117053. 1.02 323. 7.34 3.98 0.102(; ,(\.0235 0.0235 {J.OI08
0.0033 (;. vl;94 4.34 1.(;(.e 0.785 1.1.. U 1437. 44. -0. 3960.00
16707. -C.OO 126.27 1.17 956. 5.<;04 0.S8
248. 39500. 1085. 525602. 31. 2G265C. 134013. 1.12 353. 2.85
3.62 0.0937 O.U246 0.0246 0.0120 0.003.:> o.oe93 3.81 I.COG 0.778
1.02 1433. 44. -0. 3960.00 18195. -O.tO 53.88 2.61 986.
6.12U 1~ CO
287. 41700. 1162. 523949. 38. 196337. 145(;26. 1.20 365. 2.47
3.46 o.c906 0.0257 0.0257 0.0130 0.0035 a.GO'l2 3.52 0.999 0.774
1.03 139u. 42. -0. 3960.00 LS848. -0.00 50.18 1.77 1040.
1~
) 6.099 1.04
Table 16 (cont'd). Final Blended Body. Double Delta Configuration Data
146
TI ME .ALTITUDE VELGCITY WEIGHT RANGE THRUST DRAG MACH Q GAMMA ALPHA CLAVAL COACT CDCOR COO
COL COF ALD PR SFC
) PfC WA wf wax AE
FUEl OXIDIZER RIC DVDT VI PH CMl
* * '" * * * * "* * * * * * * * * * * * * * * "* * * * * * * 382. 45CO\,j. 1356. 519888. 58.
20(859. 157560. 1.40 424. 1.38 2.90 u.C776 C.('241 c .• 0 241 v.Ol17
G.CC34 v.tG89 3.22 0.991 0.765 1.04 142 b. 43. -0. 396Ci.OO
23909. -C.00 32.55 1.91 i220. 6.828 1.21
46'1. 470uv. 1551.;. 516057. 79. a 562tl. 17CJ'::>67. 1.60 504. 0.88
2.'tb 0.C64<; 'j. (.1219 0.0219 0.0106 0.0028 L. UL 85 2.96 0.969 0.762
1.G5 1546. 46. -0. 396l-.00 2774J. -0.(,0 23.91 2.32 1495.
8.3,,(' 1. '::>2 '-
530. 48LCli. 1743. 512797. 97. 241f76. 1851i9. 1.Bu 608. 0.55
2.18 0.C534 0.0197 O.C197 0.0('96 (j. (-02(, C.l-vBI .2..71 0.951 0.763
1. -j6 1753. 51. -0. 3960.00 3100G. -C.lIC 16.79 3.25 1699. 1D.696 1. 73
588. 4870u. 1937. 5G9991. 113.
) 27102.. 2(;0267. 2.lie 726. 0.45
1.93 0 ..... '+45 (J.OI7S C.0179 C .0087 0.0015 0. Cu 77 2.49 0.932 0.770
1.06 2li(:3. 58. -G. 3960.00 33 8Do. -c. (,0 15.35 4.25 1899. 14.C86 1.93
674. 495(;C. 2421. 5lAC 13. 143. 353539. 24(678. 2.50 1<;91. u.7lf
1.47 0.0291 0.0143 0.0143 0.0070 0.0e07 O.l.O66 2.04 0.877 0.819
1.06 2849. b(;. -0. 3960.0C 3<;7 64. -O.lO 11.e5 7.05 2395. 29.605 2.44
7 3iJ. 50GOO. 2905. 498579. 168. 441661. 276989. 3.GO 1534. 0.21
1.16 O.GZC4 C. 011 7 0.0117 0.0(;56 C.O('J4 0.C057 1.74 0.816 0.935
1.Ub 4130. 115. -0. 3960.00 45218. -v.lO 10.85 10.51 2886. 62.145 2.94
757. 5040li. 33 7C. 494645. 182. 612723. 30624 '7. 3.48 2026. 0.29
U.96 C.<"'152 0.0(,98 C.()U98 0.0046 O. Cu02 O. CO 50 1.55 0.816 0.906
1.06 5390. 154. -0. 396U.OO 49151. -0.(0 17.01 19.77 3355.
123.046 3.43
Table 16 (cont'd). Final Blended Body, Double Delta Configuration Data
147
)
TIME )" HR US T ALPHA
C0l PFC
FUEL PTl
AL TI TUDE DRAG
CLt~VAL COF WA
OXIDIZER CMl
VELOCITY MACH
COACT AU)
WF RIC
WEIGHT Q
COCOR PR
wax OVOT
RANGE GAMMA
COO SFC
AE Vl
* * * * * * * * * * * * * * * * * * * * * * * * * * * * * 75 8.~
615203. 1.1.95
G. COQ2 1.06
493U8. 125.991
785. 568134.
0.94 C.00u2
1.J6 53279.
189.152
920. 400216.
1.15 0.0004
1.09 57487.
212.134
86b. ~utJ21.
1.36 t..O:J..Jo
1.12 61681.
254.445
';)28. 237151.
1.59 C.ODl8
1.16 66149.
303.448
1015. 188771.
1.76 0.0011
1. 2v 7128J.
379.324
tUill 108621.
3.6a C.0045
1.43 74167.~
179.812
;){., 5(.·J.
3(,62.79. r,• D151 V. 01,., 5C
54C9. -('.(,0
3.45
5640(;. 263 6l; 3.
('. C151 (.(.045
4977 • -L..CC
3.94
675tJL. Ib4vo5.
(,.\.;<:01 u.C044
3557. -t .l·O
4.41
7640u. 144574.
0.0245 0.(0-+5
2786. - U. (. 0
4.87
847liv. 121632. 0.0293 D. (T48
2225. -t..(;{;
5.32
913Uu. 113352.
<:.0328 0.(053
1836. -C.C0
5.77
'-U7454.] 89531. (;.(;678 0.0073,
lC43. -C.OC
5.51
3390. 3.50
C.vv97 1.55 155.
98.9;!
3674. 4.(,0
G.ULb5 1.76 142.
221.13
4363. 4.5(;
O.OC80 / 2.52 101.
264.62
4878. 5.UO
'j-;-OU7 3.16 so.
1b6.81
5396. 5.50
C.(;{'79 3.7C
64. 113.29
5913. 6.l.0
0.I..lL84 3.90
54. 61.11
5991.
L.0139 ~ \ 31. ) 8~
494488. 2039.
C.V097 0.816
-0. 19.16
49C518. 2009.
0.0085 0.816
-0. 18.15
48631J. 1496.
0.0080 0.816
-0. 12.36
482115. 1212.
C.UC77 0.816
-0. 9.7U
477648. 995.
('.(\(;79 0.816
-0. 7.11
472517:] 873.
L.CC84 0.816
-0. 4.81
46963C!.· ~-4r9~
c .,..'139 0.816
-0. 0.88
183. 1.67
C.0046 0.905
3960.00 3374.
199. 3.27
0.0038 0.900
396C1.CQ 3859.
222. 3.48
0.0032 0.912
3960.(.( 4345.
257. 1.96
0.0027 a .939
3961j.OO 4856.
310. 1.20
0.0023 0.977
396(1.0C 5370.
391. 0.59
0.0020 1.021
3960.00 5884.
4bO. 0.71-
0.0020 1.021
3960.00 5924.
Table 16 (cont'd). Final Blended Body. Double Delta Configuration Data
148
TlME THR US T ALPHA
COL PFC
FUEL PH
ALTITUDE: DRAG
CLAVAL CDF VIA
OX ID I lER C 1-11
VELGCITY MACH
COACT ALD
\"IF RIC
WEIGHT Q
eDeOR PR
wax DVDT
RANG-E GAMMA
COt) SF!:
.0'1' A'E uJ. ~," VI ~
* * * * * * * * * * * * * * * * * * * * * * * * * * * * * ,~-,
~ as 706.·
3.&9* 0.tO .. 5*
1.44* 166051. 179.796*
5149. 22240.
2.n 0.0(,29
1.30 166829. 130.5;'5
526L. 21178.
3.40 0.004':'
1.32 167517 •
75.857
5372 • 19231.
4.(;0 C.0058
1.34 168111.
42.3l:!6
1853 ... 4.27
0.V07L) 1.32
168b50. 25.5J5
559u. 19258.
4.1l7 O.0U68
1.26 16;188.
16.d99
56<;4. 19521.
3.91 0.0007
1.2(; 169722.
10.951
\ 112231~ 897(,6.~
C.0679* C.C074* 1044.· -v.tO
5.51 *
1030U(;. 74982. 0.0541 0.(,C62
1058. -O.tO
5.16
102500. 73174. C.('644 0.0060:.
940. -0.( 0
4.66
102(,00. 73850. 0.0762 0.0069
835. -D.Ot!
00 .'1 '
76649. (J.0868 c,. 0(, 74
794. -o.oc
3.7C
930(JO. 80920. O.C8b5 U .0078
831. -0.00
3.26
86600. 85863. 0.C880 O.()G82
871. -O.lO
2.81
5991 ... 6. 0.0
lJ.0139* 4.8_-1'"
,."- --.., 24. *:
5463. 5.50
C.0113 4.79
6. -70.36
4965. 5.00
0.0131 4.94 ,
6. -4.43
4467. 4.50
G.0159 4.92
S. -4.66
4. (;C' U.0182
4.76 5.
-28.35
3453. 3.50
0.(;191 4.53
5. -56.48
2947. 3.00
O.02es 4.28
5. -63.70
371745,. 419. •
0.0139 • 1.000 •
-0. -1.15· .!-~
376967. 430.
0.0113 1.000
-0.--4.04 .
376280. 363.
0.0131 1.GOO
-0. -4.42
375686. 301.
0.0159 1.000
-0. -4.64
273. 0.0182
1.0CO -,J.
-4.76
374608. 275.
0.0191 1.000
-0. -4.77
374075. 271.
0.0205 1.000
-0. -5.02
-433fw
0.0020* 0.975·
3960 .00 *~ '" 5923'. *
4458. -0.74
0.0023 1.016
3960.00 5412,.
4555. -0.05
0.0027 0.917
3960.00 4904.
4640. -0.06
0.0032 0.949
3960.00 4394.
-0.41 0.0038
0.936 3960.00
3884.
4783. -0.94
0.0046 0.941
396(;.00 3378.
4837. -i.24
0.0056 0.968
3960.00 2875.
Table 16 (cont'd). Final Blended Body, Double Delta Configuration Data
149
?~tI *i511# of Cruise
(
)
)
TIME THRUST ALPHA
CuL PFC
FUEL PTl
Al TI TUDE DRAG
CLAVAL C[)F 'riA
OX ID I ZER CM1
VElUCITY MACH
COACT ALD
WF RIC
WEIGHT Q
CXOR PR
W[)X DVOT
RANGE GAMMA
COO SFC
AE VI
* * * * * * * * * * * * * * * * * * * * * * * * * * * * * ,5786. 13851.
3.63 0.0062
1.15 17C339.
7.43d
'j 877 • 14<;43.
3.36 0.0057
1.10 170855.
5.337
5<;72 • 15<;77.
3.18 0.0052
1.06 171386.
4.338
6e 92. 1216:;.
4.3v U.C038
1.u0 172143.
4.068
&~U~ ~'t(.
4.24 0.003\..
1.00 172502.
4.618
7140. 60G51.
4.37 C.0019
1.tO 185671.
4.618
783 d. 59441.
4.17 0.0018
1.(;0 194892 •
4.618
7850~.
92299. 0.u871 0.(.086
447. -O.lO ~.35
69(:00. 99540. 0. \.;870 G. v{;90
417. -",.I)IJ
I.!:: 8
57500. 106354.
O.L893 0.(;1:.94
450. -\.i. 00
1.39
45(;00. 80967. 0.11 04 C.lJu98
5j5. -li. t l
1.l C
40UOG. TY(!vo.
u. Ie 71 0.01.97
620. -v.00
1. L (;
401.06. 6l 092. 0.1066 0.0103
517. -c. lIO
1.C.O
40(J(j\,;. 59441. 0.1015 0.0103
516. -o.ou
1.00
244~.
2.50 v.0218
4.00 6.
-9('.04
1941. 2.00..,
0.0235 3.71
5. -105.93
1453. 1.50
U.0Z57 3.47
6. -121.75
968. 1.CO
0.v2.42 4.56
7. -!:l2.94
872. V • .::tU~
(;.1.;212 5.04
B. -95.70
87L.. ",.90
0.<.,175 6.10
13. -95.70
872. (.;.9-)
0.0173 5.87
13. -95.7u
373453. 275.
0.0218 0.888
-D. -5.58
372941. 275.
0.0235 0.937
-0. -5.55
372411. 268.
e.0257 0.979
-0. -5.13
371649. 217.
0.0242 1.000
-0. -3.21
--/\ ( 3712:5. )
....... L-L~
0.0212 1.000
-0. -1.85
358125. 223.
0.0175 1.000
-0. -1.85
-(348705. "-. 223
O.1iT73 1.000
-0. -1.85
"'
4iH8. -2 •. 11
0.007C 1.541
3960.00 2374.
4911. -3.13
0.0(;87 1.296
3960.00 1874.
4937. -4.81
0.0112 1.316
3960.00 1378.
4961. -4.91
0.0106 2.030
3960.00 968.
4968;' --cT.--;;r<;T
0.0(,85 2.6C 1
3960.00 968.
4968.~
-6.30 0.0(152
0.789 3960.00
968.
4968. ~ -6.30
0.0052 0.790
3960.00 968.
\ \ Table 16 (cont'd). Final Blended Body. DoUble Delta ,Configuration Data
150
TIME AL TI TUlJE VELOCITY WEIGHT RANGE THR UST DRAG MACH Q GAMMA ALPHA CLAVAL COACT COCOR COO
COL CDF ALO PR SFC PFC WA WF wax '.' AE
FUEL .. OXIDIZER RIC Dvor VI PTI CM1
• * * * * * • * • * * * * • * * * * * * • * * * • * * * * • 7939. 25000. 813. 347733. 4982.
13637. 90686. 0.80 352. -11.80
2.58 0.C625 0.0167 0.0167 0.0069
0.0009 0.(v89 3.74 1.000 3.988
1.00 1104. 15. -0. 3960.00
1960t3. -O.L 0 -166.25 -0.57 968.
8.325 1. (0
8C79. 5(UO. 659. 344403. 499Q ..
14499. 96518. v.60 444. -11.68
2.08 o. (;492 0.0141 0.0141 0.0053
0.0004 c. LO 83 3.49 . 1.000 8.116
'1 ~ 1.(,0 2047. 33. -0. 3960.GO
, / 199394. -v.ClO -133.34 -1.15 968.
15.5S9 l.lO
'" 8270. 50. 202. 338550. 4999. "'-
3185. 1.., o'/.n 0.20 486~ 4.00 .L' v..,.ve
2.v8 0.0492 O.U141 0.0141 0.UU53
0.OU04 O.b083 3.49 I.Ot,;o 1. I~/j
1.00 340. 2. -0. 3960.00
205246. -o.tO o. -1.15 968. ~~ 15.082 1.lO '!;;,,~
II )
Table 16 (cont'd). Final Blended Body, Double Delta Configuration Data
151
17717 J TAKE-OFF SPEED = 160 KTS.
900
880
-;9 0 860 0 0 ~
~ be .... ~
840 It:: 0 I
Q)
'@ E-t
) 820 ' .
"
800 100 120 140 160 180 200
Wing Loading (psi)
Figure 78. Variable Sweep - Wing Loading Effects
)
152
)
)
AERODY~AMIC DEVICES W I·o,JG + wING MOUNTEiJ CONTROL SURFACES vEkTICAL SURFACES HURILCNTAL SuRfACES FAIKINGS,SHROUJS ANO ASSOCIATED STRUCTURE
BOey STRUCTURE
73132. 10382.
o. 1254.
STRUCTuRAL FUfL(OR CC~MCN BASIC tiKLOSING STKUCTURE PKESSUKIlEC CCMPARTMENTS SECONUARY STRUCTURE
PROPELLANT) ~ONTAINE~ o. 154141.
5193. o.
INCUCED ~NVIRC~MENTAL PR~TECTION COVER PANElS,NCN-STRUCTURAL INSULATICN
LANDING GEAR
MA IN PROP uL S IJJN ENGINES AND ACCESSORIES AIR IN[)ULTION NACELL~S,PUJS,PYLONS,SUPPURTS FUEL(LR CGMMCN)CONTAINERS AND SUPPORTS OXIDIZER CONTAINERS ANO SUPPORTS PkOPELLANT INSULATIUN FUEL SYSTtM OXI0IZE~~ SYSTEM PRE~SU~IlATICN SYSTEMS LUdRICATING SYSTEM
AEROUYNAMIC CONTROLS
& ~ "'... .~ v ....
ENGINe OR GAS GENERATOR UNITS PUWER SOUkCE TANKAGE AND SYSTEMS
POWER CCNVERSION AND DISTRIBUTION ELECTRICAL HYDRAULIC/PNEU~ATIC
GUIDANCE AND NAVIGATION
INSTRUMENTAT ION
CCMMUN ICA TlON
ENVIRONMEhTAL CC~TROLS EQUIPMENT PERSCNNEL COOLANT SYSTEM COMPARTME~T INSULATION
o. o.
72820. 29838.
5149. o.
-0. 14421.
2057. o.
5923. 160.
2848. 1748.
5856. 1620.
282. 2150.
o. 5274.
Table 17. Final Blended BodY,Variable Sweep Configuration Data
153
84769.
159334.
o.
25687.
130367.
8229.
4"lq,;
7477.
800.
514.
2025.
7707.
)
)
PERSONNEL P~JVISIONS ACCCMMODATIO~S FOR PERSONNEL FIXEJ LIFE SuPPORT FURNISHINGS AND CA~GU HANDLING EMERGENCY EQLIPMENT
CREW STATIUN CONTROLS AND PANELS
DRY STRUCTURE
DESIGN RESEKVE
PERSO,~NEL
CREW,GEAR ANC ACCESSURIES CREW LIFE SJPPURT
PAYLOAD CAKGO, PASSENGERS
USEFUL LGAU
RESIDUAL PkJPELLANT A~O SERVICE ITEMS TAN~ PRESSURIZATIO~ GASES TRAPPED FUlL TRAPPED OXIDIZER SERVICE ITEi"\S Rt:SIDUALS
RESERVE PR0PELL~NT AND SERVICE ITEMS FUll-MAIN PROPULSIDN OXIDIZER-MAIN PROPULSIUN POW ER SUURCE fiR GPElL AIHS LUoR I CAiH S
WET STKUCTURE
I N- F L I G H T L D'S S E S FUEL VENT OXIDIZER VENT PUWER SOURCE PROPELLANTS LUBRICANTS
MAIN PROPELLANTS FUEL
TAKEUFF,CLIMB,ACCElERATE CRUISE C E SC EN T LJITER LMW
OXIDIZER TAKEUFF,CLIMB,ACCELERATE CRUISE DESCENT LO ITER LANfl
TAKEOFF wEIGHT
100563. 222640.
19372. 23033. 4499.
-0. -0. -0. -0. -0.
4700. 308.
8116. 305.
1250. 25.
13000. 35000.
384. 1748.
o. 175.
o. -0.
199. 83.
2769. -0.
3989. 330.
369889.
-0.
«
13428.
300.
445232.)
o.
1275.
48000.
49275.
2307.
282.
497096.)
7088.
369889.
874073.)
Table 17 (cont'd). Final Blended Body, Variable Sweep Configuration DaUB.
154
)
J
VOLUMES BODY STRUCTURE CREW AN~ PASSENGER CCMPARTMENTS CARGO CCMPARTMENT LANDING GEtR BAYS PROPULSIUN BAY WITHIN BuOY ~ING CE~TER SECTION fUEL(OR CuMMON PROPELLANT) COI'HAINER OXIDIZER CONTAINER fUEL{OR COMMeN PROPELLANT) INSULATION OTHER BJOY V0LUME
TOTAL BODY VOLUME
WETTED AREAS GROSS bODY LOWE~ SURFACES(THERMAL PRUTtCTION) UPPER SURFACES{THERMAL PROTECTION) PERSCNNEL COMPARTMENTS CARGO CUMPAKTMENTS
PLAN AREAS WING(OR LIFTING SURFACE)(GROSS) EXPJSEU wING AREA tiUOY MAXIM0M CRUSS ~ECTIUN
B,t\S E VERTICAL SURfACES HURIZUNTAL SURFACES AIR INLET CAPTURE AREA
UNIT WEIGhTS T .~ ., ~
VERTICAL SURFACES RORIlCNTAL SURFACES BOCY STRUCTURE (gASIC) LIFflNG SURFACE MAXIMUM LUADING
DIMENSIQNAl CATA WING
STRUCTURAL SPAN RUUT CHURD LENGTH THICKNeSS RATIl]
80DY LENuTH WIDTH HEIGHT
ENGINE SCALE FACTOR
CU fT 15681. 8645.
903. 2622.
o. 2048.
7 87/ 20 • ,r/ '-0.
; 4622. ;
~-\ 18031. .l!>
\
""-13,)771.
SQ. FT. 31362.00
o. o.
4831.66 442.80
SQ. FT. '6827.70 6827.70
928.32 17 2.49
1521.76 o.
198.92
LB/SQFT ,r. -'1 ~~- . 6.82 o. 4.91
83.21
FEET
199.96 48.42
0.08
309.18 O.
28.95
9.47
CU M 444. 245. 26. 74. o.
58. 2468.
-0. 131. 510.
3956.
SQ. M. 2913.53
o. o.
448.86 41.14
SQ. M. 634.29 634.29
86.24 16.02
141.37 o.
18.48
KG! SGM n /.0 ~
0.31 O. 0.22 3.77
METERS
60.95 14.76
2
94.24 o. 8.82
Table 17 (cont 'd) • Final Blended Body. Variable Sweep Configuration Data
155
TPIF. ALTITULlE VtL,JCI TV WEI;HT RANGE THRUST iJK.:~G t-14Ct-t i,J GAMMA IILPHA Cl:'VAL COACT COCUR COO
) COL COl-' AL) PR SFC , PFC (iA flF wax AE
fUEL, LJXI U ILEf{ :<../C DVOT , VI PTl GMI
* * * * * * * * * -* ::~ * * * * * * * * * * * * * * * * * * * 11. 50. 290. 371233. O.
.)IGC36. 731:>C8 • 0.20 100. 8.84 1:;.00 0.17<;8 0.05:'.2 0.0:>b2 0.0166
0.0277 0.011 <; 3.20 1.000 0.979 1. J J 5519. b 6. -0. 3960.00
2E40. -O.CO 30. ,;1 2.62 1042. 15.':':)2 1.00
'-55. SOC O. 65.;1. 3650Sli. 3.
f>Zb22). 100934. 0.60 444. 14.08 :; .11 0.1::'>30 0.021 7 0.0217 0.0068
O.CO">:> 0.0114 7.76 1.000 0.942 1.00 :),+:;8. 164. -0. 3960.00
dS82. -O.JO 160.1'~ 11. 71 1042. 15.5~9 1.00
90. 15000. 846. d59846. 7. 5t2405. 135367. O.dO 535. 19.75
4.27 0.13(:5 O. 0241 0.0241 0.0088 o.con 0.0116 5.67 1.000 0.874
1.uO 4540. 13 7. -0. 3960.00 142.26. -0.00 285. ':14 5.12 1042. 12.b49 1.00
) 121. 2:;00,). 91S. d56060. 11. 459bO~. 14ld26. O.~O 446. 18.35
5.22 0.1~44 0.0303 0.0303 0.0112 :J.C070 J.Ol.21 5.43 1.000 0.828
1.00 3477. 106. -0. 3960.00 1 Q (11 ".2 -,\ Ilr) .., O:..} ""\~ , 0 I 1 "I . ..., .LI.,.:VLJe v.vu £.. U u • ..,;u ,l..U'"T l.v..,.c...
9.23b 1.00
173. 37500. 988. 850917. 20. 333819. 1bC710. 1.02 323. 9.13
7.61 J.2345 0.0474 0.0474 0.0146 J.Cl"7 0.:)132 4.94 1.000 0.785
C.':I9 235>3 • 73 • -0. 3960.00 2315..,. -0.00 156.72 1.46 97-0.
5.':164 1.00
205. 39500. 10dS. 848974. z:;. 332141. 1903E5. 1.12 353. 3.64
7.27 0.2165 0.0513 0.0513 0.0165 0.0217 O.OUI 4.22 1.000 0.779
i.03 2549. 72. -0. 3960.00 25C':I9. -O.vO 68.90 3.34 922. 6.114 0.92
235. 41100. 1162. 346865. 30. 321416. 214419. 1.20 365. 3.19
7.28 0.2093 0.0559 0.0559 0.0182 O.024b 0.0131 3.75 1.000 0.774
1.0 4 2276. 69. -0. 3960.00 27208. -0.00 64.74 2.28 936. 6.081 0.92
)
Table 17 (cont'd). Final Blended Body, Variable Sweep Configuration Data
156
)
J
TIM£:. THRUST ALPHA
l-iJL PfC
FUt:l PTl
ALTITUDE DRAG
CLAVAL C'JF ·r/A
OXliJIlEK CMl
VElOC lTV MACH
COACT ALD
wF K/C
WEIGHT Q
CDCOR PR
wax OVOT
RANGE GAMMA
COO SFC
AE Vl
* * * * * * * * * * * * * * * * * * * * * * * * * * * * * * bO~ •
801317 • 2.06
0.C015 1..13
t3820. LiS.n,;,
bYte 74'71'70.
2.11 0.0015
1.15 09699.
189.008
(;74. 551C61.
2.90 0.0028
1.24 76IBl.
i 211.54J
728. 442:43.
3.65 0.0042
1.35 e3022.
2:2.531
BuO. 36568 7.
4.53 O.(j062
1.49 90th.l8.
2'n.e tlO
900. 319234.
0.C080 ~.64
10C298. 36t.G91
903. 3013 24.2 •
5.57 0.0090
1.69 100563. 340.906
50500. 34E620. 0.0358 0.OC72
7045. -J • ( ()
-3.38
5':1400. 311JJI.
0.0.359 0.OOc6 6563. -0.00
3.85
67500. 2 11353<;. 0.0475 O. CC70
"t8 <; 8. -0.00
4.26
16400. 214910.
0.0577 0.;)C75
4025. -0.00 4.64
d4700. 2C4767.
0.0089 0.OC87
3430. -0.00
4.97
91300. 20d752.
0.0768 0.0105
31 C:i. -0.CO
5.30
':12689. 208630. 0.0815 0.0110
29<;5. -o.UO
5.25
3390. 3.50
0.0163 2.20 201.
813.48
3874. 4.00
0.0148 2.43 187.
193.62
4363. 4.50
0.0155 3.06 140.
229.46
4878. 5.00
0.0169 3.42 115.
143.08
5396. 5.50
0.0196 3.52 99.
96.24
5913. 6.00
0.0228 3.37
91. 54.67
5919. 6.00
0.0243 3.36 87.
442.25
810253. 2039.
0.0163 0.825
-0. 17 .14
804374. 2009.
0.0148 0.825
-0. 15.89
797891. 149b.
0.0155 0.825
-0. 10.72
791051. 1212.
0.0169 0.825
-0. 8.32
783265. 995.
0.0196 0.825
-0. 6.04
773774. 873.
0.0228 0.825
-0. 4.30
773509. 819.
0.0243 0.825
-0. 1.76
142. 1.50
0.0076 .0.905
3960.00 3354.
160. 2.86
0.0066 0.900
3960.00 3838.
187. 3.01
0.0058 0.912
3960.00 4313.
228. 1.68
0.0051 0.939
3960.00 4813~
289. 1.02
0.0046 0.977
3960.00 5313.
382. 0.53
0.0042 1.021
3960.00 5814.
385. 4.29
0.0042 1.021
3960.00 5811.
Table 17 (cont'd). Final Blended Body, Variable Sweep Configuration Data
158
()
,r ')
/
T IMc THRUST ALPHA
CDL PFC
FUF.:L PTl
ALTITUDe l)RAG
CLAVAL CDF WA
OXIDIZER GMl
VElOC I TY MACH
COACT AlD
wF RIC
WEIGHT Q
CDeOR PR
wax DVDT
RANGE GAMMA
COO SFC
AE VI
* * * * * * * * * * * * * * * * * * * * * * * * * * * * * * 514' •
21l92d. • 5.66 •
0.L09j· 1.70·
323203. 340.20d •
5232. 47S57.
5 • .3 :) 0.((;84
1. SC) 324360. 183.132
5314. 44679.
(:.20 0.0115
1.6.2 32541<3. IDS .S4 2
53n. Lt07't2.
7.27 0.0164
326.313 • ;)0. ·'vv
:5 463. 36(5).
8.64 0.02'1-5
1.68 :3 27.C27.
; 1. 362
5525. 30t54.
10.50 0.0392
1.70 3275(:5.
15.945
5590. Z" 738.
lC.04 0.0395
1.54 328(84.
10.574
<..9:>92.. 2111328. * o .OdZ8 * 0.0111·
3014. • -G.ro
5.24 •
"5)00. lSU,j20. 0.0801 J.U1UO
2234. -0.00
4.88
945CO. 14"188.
0.')953 0.0105
20C7 • -0.01)
4.3>3
940()0. 1533'77. 0.1156 0 .. (J1l5
1778. -O.CO - ,,-. .• v
93500. 1£::2725. 0.1430 0.0130
1550. -0.(;0
3 • .39
93JCO. 178937.
0.1843 0.0152
1323. -0.00
2.88
86600. 180468.
0.1875 0.0150
1327. -0.00
2.51
5919. • b.OO
0.0246 • 3.36 • 57. •
5434. 5.S0
0.0231 3.47
14. -52.04
4938. 5.00
0.0271 3.51 12.
-6.14
4443. 4.50
0.033 7 3.43 11.
-6.63
394':3. 4.00
0.0442 3.25
9. -7.47
3453. 3.50
O. )620 2.97
8. -8.74
2947. 3.00
0.0634 2.96
8. -99.07
:;50869. 819. •
0.0246* 0.825 •
-0. 1.76 *
549712. 619.
0.0231 0.825
-0. -5.67
548655. 523.
0.0271 0.825
-0. -6.09
547760. 434.
0.0337 0.825
-0. -6.57
547046. 351.
0.0442 0.825
-0. -7.39
546508. 275.
0.0620 0.825
-0. -8.64
545989. 271.
0.0634 0.825
-0. -7.80
4':>14.
0.0042" 0.968"
3960.00 " 5809. •
4595. -0.55
0.0046 1.016
3960.00 5333.
4665. -0.07
0.0051 0.977
3960.00 4819.
4726. -0.09
0.0058 0.949
3960.00 4300.
4775. -0.11
0.0066 0.-936
3960.00 3774.
4813. -0.14
0.0076 0.941
3960.00 3236.
4847. -1.93
0.0090 0.968
3960.00 2742.
Table 17 (conttd). Final Blended Body. Variable Sweep Configuration Data
159
·Start of
Cruise
)'
)
lIME ThRuST pLPHi-I
COL PFC
fUEL PTl
ALTITUDE uKAG
CLAVAL COF
OXIJIIER CMl
VELOCITY MACH
COACT ALl)
,oiF R/L.
WEIGHT Q
CDCOR PR
wax OVDT
RANGE GAMMA
COO SFC
AE VI
* * * * * * * * * * * * * * * * * * * * * * * * * * * * * * ~\..:5o.
272C? S.27
O.G377 1.39
32Bt7o. 7.30"t
572(;. 27644.
8.60 0.036<;
~.26
329508. 5.2Bd
5tOl. .:::896"t.
" ... ' Ce,-'-
0.0.383 1.13
33C2<;7. 4.30';
5S 13. 16C82.
7.63 0.G190
1. UO 331457.
4.068
5<;68. 1348b.
7.24 0.0136
1.00 332092.
4.618
oS68. 61846.
6.87 0.0057
1.00 3456t:3.
4.618
7666. 60S18.
c.64 0.0053
1.00 355126.
4.618
7E5CO. 181416.
0.18Su 0.0145
755. -0.(,0 ~.12
69000. 185030.
O.LE54 D.(H42
702 • -v.LO
1 • ()s·
:>7:>LiO. 19lCt:5.
0.1900 0.0140
752. -o.uo
1.20
45000. 100945.
0.2365 0.0138
861. -0.00 1.00
40000. 8963L. 0.2295 0.0135
10C8. -0.00
1.00
40000. 61846. 0.2281 0.0141
480. -0. (.0
1.00
40000. 609ltl. 0.2204 0.0140
478. -0.00
1.00
2442. 2.50
0.0629 2.95 11.
-121.25
1941. 2.00
0.064G 2.90
9. -134. d4
1453. 1.50
0.0682 2.79 10.
-150.2Q
968. 1.00
0.0470 5.03 11.
-75.02
872. 0.90
0.0383 5.99 12.
-80.51
872. 0.90
0.0264 8.64 14.
- 80.51
872. 0.90
0.0260 8.47 13.
-80.51
545197. 275.
0.0629 0.913
-0. -7.52
544504. 275.
0.0640 0.954
-0. -7.07
543776. 268 •
0.0682 0.992
-0. -6.33
542616. 217.
0.0470 1.000
-0. -2.91
541980. 223.
0.0383 1.000
-0. -1.56
528410. 223.
0.0264 1.000
-0. -1.56
518947. 223.
0.0260 1.000
-0. -1.56
4876. -2.85
0.0107 1.388
3960.00 2253.
4901. -3.98
0.0129 1.207
3960eOO 1758.
4922. -5.94
0.0159 1. 239
3960.00 1232.
4944. -4.44
0.0142 2.405
3960.00 9-68.
4953. -5.30
0.0112 3.311
3960.00 968.
4953. -5.30
0.0066 0.790
3960.00 968.
4953. -5.30
0.0066 0.790
3960.00 968.
Table 17 (cont'd). Final Blended Body. Variable Sweep Configuration Data
160
-)
TIME ALTITUDE VELOCITY WEIGHT RANGE THRUST DRAG MACH Q GAMMA ALPHA CLAVAL COACT COCOR COO
CDL COF ALD PR SFC PFC WA WF WOX AE
FUEL OXIOIlER RIC ovaT VI PTl CMl
* * * * * * * * * * * * * * * * * * * * * * * * * * * * * * 7tl03. 25000. 813. 516476. 4972.
13E03. 91727. 0.80 352. -8.00
4.31 0~1378 0.0248 0.0248 0.0088 0.(;037 0.0122 5.56 1.000 6.170
1.00 1774. 24. -0. 3960.00 357596. -0.(;0 -113.21 -0.39 968.
13.325 1.00
8014. 5000. 659. 508464. (,997.
13657. 90927. 0.60 444. -7.42 3.31 0.1079 o. 0195 0.0195 0.0068
0.0014 0.0113 5.53 1.000 13.763
) 1.00 3319. 52. -0. 3960.00
365b09. -0.00 -85.09 -0.73 968. 15.5'::19 1.00
8203. 50. 261. 503964. 4997. 5127. 14534. 0.26 486. 4.00
3.31 0.1079 0.0195 0.0195 0.0068 . ~~~ , 1 1 0
U.UU1't U.Ul1~ ??~ ".vvv ~."'.LV
1.00 263. 2. -0. 3960.00
370108. -0.00 o. -0.13 968. 15.359 1.00
:', J Table 17 (cont'd). Final Blended B'ody, Variable SWeep Configuration Data
161
9.0 PROPULSION EVALUATION CRUISE MACH> 8.0
See Report GD/C-DCB-66-004/2A.
)
)
162
)
)
10.0 COMPARISON OF CONFIGURATIONS
From the data shown in Sections 6.0, 7.0, 8.0 and 9.0, a comparison of the delta wing, variable sweep,blended body and scramjet configurations can be made. Each of these sections is consistent within itself; however, continuous updating was done throughout the study so that much of the data shown in one section cannot be compared directly with that of another section. Near the end of the Phase I studies, all configurations were checked for relative consistency and revised computer runs were made. These are the full computer printouts shown in Tables 10, 12, 16, 17 and 2l. Data from these tables are summarized in Tables 22 and 23.
Table 22 shows the final data for the four candidate configurations for the cruise Mach number regime from 3 to 8. The take-off weight of the blended body/variable sweep configuration is considerably higher than that of the other three and this configuration is therefore eliminated from further consideration.
Table 23 shows the four remaining competitive configurations of delta wing, variable sweep wing, double delta-blended body and scramjet. Included in Table 23 is the sensitivity to take-off weight of various key parameters. For the delta wing configuration, it can be seen that the addition of a passenger plus baggage (weight = 215 lb) wlll Increase the take-off weIght by 1800 tb ( 8. 5:1). Also, It can be seen that take-off weight is 5.5 times as sensitive to an increase in subsonic range compared with an increase in hypersonic range (830 compared with 150).
Before selecting which of these configurations would be best for the Phase II studies, it was necessary to consider sonic boom, mission analysis and cost effects. Selection was therefore held in abeyance pending the outcome of the results from these other study areas. (Final selection is made in Section 14.0.)
199
l\:) o o
'----'
, V ~
BLENDED BODY
DELTA VARIABLE DOUBLE VARIABLE WING SWEEP WIN,G DELTA SWEEP
~ ~8- ~ -c:-> WING/ PLAN WADING tH/ - 1::::::>/- -/35 130/82
ASPECT RATIO 1. 45 1. 47/6. 5 1.5 .67/6.1 THICKNESS .06 .05 - -SWEEP 70° 70° 80°/65° 80°
BODY FINENESS 12 12 - -
PROPULSION M=0-3 TURBOF ANRAMJET
CRUISE MACH NO. 6-7
TRAJECTORY TRANSONIC L1P ~ 3 PSF (M = 1. 4 @ 45,000 FT) q 2000 PSF
INLET PRESS. 130 PSI DESCEND ::::: L/D MAX LOITER M = • 9 <g 40, 000 FT
100 N. M. + 1,000 SECS.
TAKE OFF WEIGHT (LB) 537 040 602,483 543,797 874--,-0.-73
Table II . Com parison of Configurations, Cruise Mach 3-8
t-.:> o ......
L
WING LOADING
CRUISE MACH NO.
TAKE OFF WEIGHT ~----
SENSITIVITY TO WTO:
RANGE (LB/N.M.)
PASSENGER (LBI PASS)
SUBSONIC RANGE (LBI N.M)
SUBSONIC LOITER (LB ISE:C)
Table III.
\~
DELTA VARIABLE WING SWEEP
~ -<Bt 81 125
6-7 6-7
537,040 602,483
140 158
1,700 1,900
830 660 90 72
Final Configuration Data
J
BLENDED
BODY SCRAMJET
~-::}- ~ 35 35
6-7 8
543,797 846,927
145 200
1,750 2,400
920 1,300
98 138
)
)
11.0 MISSION ANALYSIS
The final objective of the Phase I studies is to select configurations for detailed analysis in Phase II. Number of passengers, cruise Mach number, and design range are important parameters to be selected for these Phase II configurations. A mission analysis was conducted to:
(a) Consider the implications of a hydrogen-fueled transport.
(b) To define, or at least to bracket, the design range, the cruise Mach number, and the passenger capacity of the vehicle.
11.1 SELECTION OF DESIGN RANGE
The range for which the vehicle should be designed is principally determined by the geographic/passenger densities of the future.
11.1.1 PASSENGER TRAFFIC 1965-2000. Lockheed, Reference 12, as part of their study of a suborbital global transport system, completed some extensive projections of future world population, gross national product, and trade for each country from 1965 to 2000. Lockheed's assumptions and projections form the basis for many of the conclusions of this study, since their data appeared to be the most consistent and extensive set available.
Figure 96 presents world total annual two-way air traffic from 1980 to 2000. The low and hfgh estimates are 95% confidence limits about the mean. As can be seen, long range projections are subject to an order of magnitude variation. The total traffic, P ,between countries is assumed to be proportional to the products of their per caPitaTproducts (GNP divided by population) and the sum of the imports and exports (E A + E
B) between the two countries.
i. e., P T "" (PCP 1) (PCP 2) (E A + EB
)
By studying actual data from 1953 - 62, Lockheed derived a correlation coefficient of 7.3 x 10-5 , thus
PT
= 7.3 (PCP1
) (PCP2) (E
A + E
B) 10-
5
giving the annual two-way traffic in millions of passengers.
202
200 11--1-.--. -·I~·t:-;-III--II-IIII ~:- ...... t-.1
• ~ ~-~=~-~~ ---+-----
~~_I--_L .. f-.-- . ~ - - ... ~- - -,: _ L_~_._ _ .. _. ___ _ ,
p:--.:=-"-~c , " ._~~:. C;T-:-: ==-j~_~ ~--~-~ , .. _--- ---.... ---
, -------;.-~~-I--- - -'--S~- ~ I - _~_ .• _!~._L ......... _. _~-~:.-.~ , I ,: ': i . -=-=-~-~ ~:~.=-:~ .... ~---.... -~.-~ 100 + 20 HIGH
.:""'" ...... • •. --'- -··:;':'·i-cc. ::' t..~ .. ~: .. ,i.E·. : .. c.~ C 0
.. , ::_z- .~-...• _._ .. . -~;.i'::.:'-: ~; .:,. c.
• - _ _ _ *_. , .• _._ _ _ : -4 .~: ~7.. -: .' ~ ... 'n
~ ~
80 .'= '."'"'. . .... _ .i-~ === -:-"".-~ ~:'~ =.:~L···
E-~ -=-::~.E:!: C:-~.:=~..:..": .. -:':-~: .-:: ~ ;==_'~:i.·: _
u 'n CH CH cd ~
E-i MEAN
) ~ :?:
I 0
~ r-!
~ ~ "" r-! cd .p
~ 'D r-! ~ 0 ~.~ --
, , , :7 .,-
V I
! , If I ,¥ ! , '.-I', , !
I I , , , I I
i I I...,. I I , ! !
10 1960 1970 1980 1990 2000
YEAR
) Figure 96. World total annual two way air traffic
203
It is interesting to note that due to a lack of data, Lockheed excluded from ') its estimates the Soviet Union, the European Communist Block, and Communist China
which represent about one-third of the world's present population. It is possible that by the latter part of this century there will be significant traffic between the Free World and parts of the excluded areas. Table 24 shows the countries included in the various world areas by Lockheed and the selected principal terminal in each area.
)
At this time in the study, it was not known what the economics of a hydrogenfueled, commercial transport might be. It was peSSimistically assumed that, at least in the early phases of operation, this type of a commercial transport would mainly benefit businessmen and government o(ficials who would be attracted by short trip times. (This basic assumption has been questioned and will be reviewed in Phase n.)
In Reference 12, Lockheed estimated the percentage of the total traffic between each world area pair that would be for business and government purposes. Values from 30% t? 75% were estimated for each area pair. The lower values were assigned to travel between mature, well-developed areas and the higher values were assigned to the emerging, under-developed area pairs.
Figure 97 shows the business plus government traffic as determined by Lockheed. The curves show the business and government traffic between the major world areas during the years 1980-2000. The most Significant feature of Figure 97 is that North America to Europe will have by far the heaviest traffic.
Figure 98 summarizes the passenger traffic for the years 1980 to 2000 and shows that:
1. Of the total inter-area passengers, 65% will travel between North America and Europe.
2. Of the total inter-area passengers, 37% will be on private or government business.
3. Of the total business and government traffiC, 53% is between North America and Europe. The right-hand block in Figure 98 also shows where this traffic originates.
The data shown in Figure 98 are valid for all years between 1980 and 2000 and are essentially independent of the level of the estimate; i.e., low, mean or high, as shown in Figure 96.
11.1. 2 GEOGRAPHIC DISTRffiUTION OF PASSENGERS. It is clear from the foregoing discussion that the North America to Europe route dominates the flight frequency. Since the purpose of this analysis was to determine the design range, Figure 99 was constructed. This shows how the 1980 mean level of two-way business and govern-
) ment traffic is distributed with range and shows that New York to Paris provides 42%
204
WORLD AREA COUNTRIES INCLUDED AIR TERMINAL AT
North America U.S.A. New York City, Canada Los Angeles Mexico
Europe Austria Paris Belgium Denmark France Germany Iceland ireland
) Italy Luxemburg Netherlands Norway Portugal ~n~in -,..
Sweden Switzerland United Kingdom Finland West Berlin Yugoslavia
South America Argentina, Bolivia, Brazil Rio or Sao Paulo Chile, Columbia, Equador, Paraguay, Peru, Uruguay, Venezuela
Japan Japan Tokyo
Oceania Australia, New Zeland Sydney
) Table 24. Global Transport Air Terminals
205
.. r" i/
10 9
8 h~itdjj4~t$!l!$~~4~8~~~j NEW YORK - EUROPE (PARIS)
7
6
5
4
r.I)
3 Z 0 ::3 ~ ..... ~ I
u 2 ..... r... r... <: 0:: E-< E-< Z ~ I'I~~IIIIIII~II'I~II LOS ANGELES - EURO\P)~.,:p.,::s) NEW YORK - JAPAN -
~ Z 0:: ~ 1.0
·····-,-~+-,--l-~NEW YORK - SOUTH AMERICA
;> 0 .9 0 r.I)
.8 0 EUROPE - OCEANIA
~
) 0.. .7 EUROPE - JAPAN Y:;;:'
EUROPE - SOUTH AMERICA .6
.5 NEW YORK - OCEANIA d:P), ~'-;-J
LOS ANGELES - JAPAN
.1
1960 1970 1980 1990 2000 YEAR
Figure 97. Passenger Traffic Projection
206
L
NOTE
100
U < ..... 0 !:xi ..... !:xi P:; < IJ:.l IJ:.l p:; ~ Pot E-; <00 80 P:; ~ t:r::E-;~ < E-; IJ:.l ~ P:; I 0
0 Z ~ E-; ....:I 60 < ;:J ~ N z 0
-.J < ....:I < E-< 0 40 E-; 0 ....:I P:; 0 ~ ~ 20 0 E-< z IJ:.l u P:; IJ:.l 0..
0
."-.-/ o
THESE DATA A~E ESSENTIALLY VALID FOR EACH PROJECTION LEVEL
(LOW, MEAN, ~IIGH) AND FOR ALL YEARS 1980 TO 2000.
E-; Z IJ:.l ~ Z P:; IJ:.l :> ~~ 0 l') 01J:.l U) p:;~
....:I ~ ~Z < 1J:.lp:; E-; 01J:.l 0
E-;6 E-; <0 0 UU) ....:I ..... ;:J P:; P:;....:I 0 ~Pot ~ <U) U) t:r::1J:.l E-<Z p:; ..... ou) Z~ f PARIS TO LOS ANGELES TO I'AllIS _4U71 LOS ANGELES TO PAIUS TO LOS ANGELES
PARIS TO NEW YOHK TO PAHlS
I NEW YOHK TO PA HIS TO NEW YORK
Figure ~)~. DistriiJutiun of passenger traffic
0
....
>::j N ..... ~ '1 (D
:.0 ;-c
~
~ ,/ I7.l
5' (D I7.l I7.l
'C -~ ..,. 0 c <: (D
8 S (D
= ~ <:J1
~ '1
~ ~ > ..... Z n
<: Q I7.l tx:1
( ~ ..... ....,
~ .... 0)
= 0
'\% 0 0
Z :~
~ ~ ;-' ..... = ~ -.J
0 =
\ ft 00
\ ......
CIJ
to
.... 0
o
80(;
PER CENT OF 1980 MEAN ANNUAL TWO-WAY
BUSINESS PLUS GOVERNMENT TRAFFIC
o
I I
N o
~ o
I
..,. o
<:J1 o
NEW YORK - EUROPE (PARIS)
I i I I ~ NEW YORK - SOUTH AMERICA (SAO PAULO)
i I I I ! I
I
I
~EUROPE (PARIS) - AFRICA (JOHANNESBURG) I I LOS ANGELES - EUROPE (PARIS) I - .
........ EUROPE (PARIS) - SOUTH AMERICA (SAO PAULO) l EUROPE (PARIS) - JAPAN (TOKYO) I
I I I , i I
NEW YORK - JAPAN (TOKYO) i !
/
I I
I
- NEW YORK - OCEANIA (SYDNEY)
. I I I I EUROPE (PARIS) - OCEANIA (SYDNEY)
(
/
,/
~.
C
J
)
) /
of the total traffic at 3,150 nautical miles. The data of Figure 99 are then accumu1.<1ted and shown in Figure 100 for the twenty-eight most important pairs of terminals in the transportation system. Figure 100 provides a fairly rational basis for selecting the design range; e.g., 5,000 nal!JjJ:!Jtlrniles will permit nearly30o/(LQLthe .. .totaLbus.iness /-------.----- ------- __ " __ ___ . __ . _____ ~._~ __ -._.; .. _._·,f~· ,,-
and government traffic to be carried non:stQI).._.5 .• 5.0.D.-nauJical mlies~l1 include 8010., 6,000 nautical miles will include 86%-:· ;;~d7, 800 nautical miles will in~i~~90%:-'---A range of 9,200 nautical miJes is required to carryall of the business plus government traffic.
Figure 100 includes a recommendation that 5,500 nautical miles should be the design range. This is based on four things:
1. It is at a knee of the curve shown in Figure 100.
2. As shown in Figure 99.5,500 nautical miles will provide non-stop flight between all the most heavily traveled routes except New York to Japan and flights to Sydney.
3. Table 25 shows great circle distances between New York or Los Angeles and major European cities. (It is most likely that in addition to the principal terminals listed in Table 25, more will be required for Europe and other areas. However,
the ground facilities that will be required to support a hydrogen-fueled aircraft may limit the number of terminals. )
Table 25 shows that a range of 4.870 nautical miles is required to reach all European cities and Cairo from New York. However, a range of 5,500 nautical miles will permit non-stop flights from Los Angeles to all cities in the first category except Cairo. The other heavily traveled routes are North America to South America and North America to Japan. It is seen that a range of 5,500 nautical miles covers the principal cities ill South AllIer ica fr om both New YOl k and Los Angeles. Also, non-stop flights to Tokyo are possible from Los Angeles. It is not possible to fly non-stop from either North America or Europe to Oceania with a range of 5.500 nautical miles. In addition, some service between Europe and South America and between Europe and Japan can be accomplished without stops.
4. Table 26 shows some other possible non-stop routes within a 5,500 nautical mile range. However, the traffic flow on these routes will be rather small compart:;d to routes in Table 25.
A possible route structure for the hydrogen-fueled transport is shown in Figure 101.
11.1.3 RECOMMENDED DESIGN RANGE. In view of the foregoing discussion, a design range of 5,500 nautical miles is recommended. This will cover about 80% of all business and government traffic with non-stop service. This recommended design range does not reflect additional constraints such as take-off gross weight restrictions due to runway limitations., allowable sonic boom restrictions on take-off weight, passenger load factors, and economics.
209
00: 00' ~i
Z' 90 ta
:;:J': ~;
, ;>4 80 < ~ I
RECOMMENDED RANGE 5500 N. Ml. 0 70 ~ , ,
~
~ < 60 :;:J Z Z < Z 50
) < ~ /
~ 0 40 00 ~ .-t
~ " ,
0 30
~ Z , , ~ 20 0 ~ ~ .,. , , O-t 10 ~
> .... ~ 0 < ~
~ 3 4 5 6 7 8 9 :;:J ~ i:l
RANGE (1000 N.M.) 0
Figure 100 .. Cumulative traffic vs range
)
210
~
tv ~
~
New York Los Angeles
New York Los Angeles
Los Angeles ~.
Paris
Paris
(Ran
North Americ
Madrid 3110 5030
North Amerio
Santiagc 4460 4860
North Americ
Tokyo 4760
Europe to Sou
Rio de J 4950
Europe to Jap
Tokyo 5240
Table:
---------~--~~--~
~
" '--..j
.~
(
ked by Traffic)
:I. to Europe
London Paris Berlin Rome Leningrad Moscow Cairo ~1000 3150 3440 3310 3720 4050 4870 N. Mi. 4730 4900 5020 5500 4960 5270 -- N.Mi.
I
i to South America •
Buenos Aires Rio de Janeiro 4600 4180 N. Mi. 5320 5470 N.Mi.
-to Japan I
N". Mi.
h America
aneiro
N.Mi.
n
N.Mi. .
5. Possible Houtes With a 5500 N. M i. Hang-e
\,-----,
tv I-' t-.J
Los Angeles
Honolulu
Paris
Tokyo
Paris
Table 26.
,
\.J ,J
Seou Peking 518 5430 N. Mi.
Los Ang;eles New York Sydney Manila Hong Kong 2240 4130 4420 4600 4822 N.Mi.
Joha p.nesburg 4710 N.Mi.
Sydn 'Jy 442 ) N.ML
Kara chi Delhi Peking Seoul Bangkok Hong Kong 33 0 3552 4430 4820 5097 5201 N.Mi.
Other Pc ssible Non-Stop Houtes with a 5500 N. Mi. Hange
)
LONGITUDE
PEKING •
)
30"S. ----------r-------1:~~~~~~S~Y~D~N;.E;Y;=~----====------
6 hrs 6 hrs
90"W I .
I
I
SANTIAGO
6 hrs
Figure 101. Route St ructure
11
6 hrs Time Difference
213
") ) 11.2 CRUISE lVIACH NUMBER
Since trip time is dependent on the cruise Mach numb~r, the question naturally arises, "What cruise Mach number will make the hydrogen-fueled transport most attractive? II To answer this question, several interrelated areas were studied:
(a) Block Time (b) Local times of arrival and departure (c) Airplane scheduling (d) Utilization
,.-/
This analysis did not consider any technical or economic constraints. These combined effects are discussed in Section 14. O.
11. 2. 1 B LOCK TIME. A primary factor in determining the best cruise Mach number is the time to accomplish the trip. Block time (take-off to landing) shows how much time is saved by cruiSing at higher speeds. Because of short trip times, the ground time to get to and from the airport becomes relatively significant. Table 27 was therefore prepared to show typical ground times and to define three Significant increments of time for non-stop and for refueled trips:
Block Time
Air Travel Time
= Time from start of taxi until wheels stop at the destination.
= Time from arriving at the departure air terminal until leaving the destination air terminal.
Total Travel Time = Time from leaving horne until arriving at the destination hotel or business a,wointment.
Figure 102 shows unrefueled range vs block time, air travel time and total travel time for various cruise Mach numbers.
From Figure 102 it can be seen that for a 5,500 nautical mile range, tbe significant times are:
It is evident that increasing the cruise Mach number brings diminishing returns in reduced trip time. Beyond Mach 6.0, the absolute gains in trip time are small.
214
t-=>
01
•
'..,/ .
!
Home to Airport
30 Min.
...
Home to Airport
30
.......:
- .
"
Check-in, Buuru
20 Min,
il ... -
Check-in Taxi Board
20 5
1--........--.
- ----
,
V I .. V .
,
TIME r'OR UNRJ..:l<'Ul!:I1<.:D 'l~UP
I,,:> "-Taxi Aec¥erate CC'uL;c IJ.:t-liuwn liulu Tu.xl r;xit aircraft, Airport
& Clilllu C 1111 In Uaf',F,I1~e, to . r (M) r (ralle.'e) - r (M) AI' nill[:'.e ~:.rowld Hutel
\ 'rnlfl'; lJu T'Lti t ion j J Min V:iri"3 ... ) Hill. ) Mtll. 1') Hili. 30 Min •
" liI1XK TIME .... 1
I -Am THAVEL TIME ....
_ '£CYl'AL 'l'.MVEL 'l'IME - .... -
'rIME FOR m.:r'UEU:D THIP (OVER ~500 N. MI.) -Accelerate r.ruise li~t. Hull 'l'u.xi I(et',.;_el 'l'axl Ac:ce lerti te CruiHe ~t- Huld Taxi Exit \Airport & Climb DUll! I & Climb Down Aircraft, to
Claim Hotel Baggage
r(M) f(raIlGd f( ) r(H) f(range) f(M) Arrange
____ I __ ~_ I Gro\ll1d
--- --- Transp. • Varies ___ ,.
.... 5 20 5 ... Varies-- :> 5 15 30
,~~
... 1 __ - 1:1UXK 'rIME --.
A]J{ THAVEL TIME .... ._---- 'lUlU 'l'Hli.V EL TIME --------_._--- .....
__ it .
T:[ lle27. DefinitilHI of T.'r1Jicai ..
..
-- - ------- ----- - ----
tv ~ cr.
\~../ 14
12
III
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10 'I' ali Ltil I" I "1' " ,pi ;1"1" 'ikJ' ' 'l~ 'jj' rnn .i.1. lUI il"llill r):" ltl~ i,," j"t! " , ' fi i'{' j ," ~,." j'l 'I:;' "" , "\' '" I. tllil \[ .i', '"'I'' ii''' I' I"·' ~~ 111' . 1~"'" 1'1' rt"t .... it" 11' '[';;'" 'ii' '-"il' " I ' r "., '11 '~ltt· ftj'l .\" I '." ", t:.:J' .. I", , , ""., ' , I 1;, ""'" ' , . t f't ' ,~ ". , H r~tTT.~ ,- .'. . - t t r) -t. r. . - i- !+T t .' ~,+:"'"" 1'"j: > I ' t t ' - t +t . llli' lin, j:tll11:] Hi' HI' It:;tHj Iq, 'iH l'lt:. lh! dHIl}l ~L!~ ~dt t~tf 1141 jjrr ~hm ;t:il H.I·tti .. J" il,1 'qj ",,~In [It :Itt ,,, ~HI I,. t .... t.t • ..f, 111,j~tf· !It·dl, ,jm
J"~' ,'. 'ill '!'I 1"li*1 I 1":' I'~ IMJ nit ,. lilli' 'ltW, till' !1 I I' t'li liP I" '''I' f+ Ij" '·Iil' tH 'l't' I' 'I! lti! i .~' 5500 n mi ' 'Ii' 'II 'jii d 1 tt:l qu lUi +1.:' il ~B:.l ~,.t ~-~.' r I ttH ~ .. 4-8.- L++r:i . . ,1: llit 11.t~-·~~tt ,,~ ,...l,' . ! • :;-r ~1~; ,,' .. -l~ -·-t'l ~. It"' -~+~ .. , 11" +,,1 "'I' Itt::' ~>- I - ,. +. j - • :'~;~m
[,II "II iI.l, 'I .. il:' ;. I .... "".IJrl- 1 /;111 It:TlTil!i. lttl lit\· ",1 ,." j-jt' lIt!. ,W dl, !I"'I" !~ .• ',-' '." , " iii! :: "i;! :.t·rr.. :. _In ttt "11, iI! t~,~ tt-~: -id'! r " 11:-: :;~ •. ilr! :.;\
ij ." CrUlse ,,, ."., "I'M' It"ltll1" ""'1 'ill I'!"'" " .. 'II [ID"" .,!, ii" "I' tH'~'I"'" ii-: • -l.:;: ~.';! 'i:1 1: ./ dtl tV t'.L - !tt ~t·tt V-;-t ~{t-i ~I- r '~.~-'-' ;.+~~ t"·.t dJ~ .. ;:.1 Mach No . .:.tl1-8:-: Irtr I" ',i, •• 1,:1+ 1 ltID1"llb JJi,fl;.;.:,:1 • I,ll F ,.it H,I. i<
II' ) 9 It;1 111; ;1:1 .11ltHlliJlr t'll IH f!~il Rf" I'll j", I ~t'ii t·p tlU'tX\ :trr ~4 ( . i '" "', '1 " ,," • , "l" 111 h" --" .•. It , " ,t:, 1.+ I l;;t; ;-r~ I'f[l.l,.. "" 11 "1/"tt lPI I • j.", +.;. t ", -+.--,.t ~'I'- '-1'''' tjt ~ .. -f"!- t tifT It' tr!
,,' "" .. , ... ,., TI1+' 'i, llW' '" l' :w1'j-t II' Ill.' I, "in I-t '1") "'11m ", Jm .!~. ;"i ,q '!'i j r +ti~rt ~.~, I 1+t' l~ +iI'~' .".--.. H-I!lt i ".- y: ,Jit· '1Q t ~ .• -t~1i'\'" PI! ~1:1 ';11 ;',; +4 ti-w.t !:-~Jl .~~. -!:r-r.l:_t -'~1: :~~nt : -i .iji; i'-~' '-~1'L:- ~I tt- tim t
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()_ "\ ~ . ':i I • "" t~, 'II.·, •.• I , f·t···, I -,-t )-+1 *- .. j ~-; ., . .!..
o 2 4 6 8 ,10
Unrefueled !'UllGe -- 1000 n, mL
0'
MACH 3 CRUISE WITH A I-HOUR STCP AT 3500 n. mi.
Figure lOZ, Trip Time vs linrcfueled Hange
,
)
)
11.2.2 EFFECTS OF LOCAL TIME. Referring to Figure 101 on page 213, it is seen
that the air terminals for the hydrogen fueled transport are located in three clusters around the world.
From Figure 97, it can be seen that most passengers are traveling East/West rather than North/South. Figure 101 on page 213 shows that these clusters are widely spaced in local time. Tables 28 and 29 were therefore constructed to indicate these local time differences. Table 28 shows which cities are on the same time and the time between groups of cities. Table 29 selects nine cities which are on the West edge, near the middle, and on the East edge of each of the three clusters. When it is noon in the city in the left-hand column, the local time in the other cities may be read .
•
To further study the effects of different cruise Mach numbers, Tables 30 and 31 were constructed. These tables show typical schedules for Los Angeles/New York/ Rome and for Los Angeles/Tokyo, respectively.
D~artures are timed to bracket the business day in the ,?ity of ori~in. Air travel time is used (the middle ordinate in Figure 102). The local arrival times are shown, together with an indication of the day of arrival. Both schedules show reductions in air travel time (and a consequent reduction in passenger fatigue) with increasing cruise Mach number. For all speeds, the associated departure and arrival times result in some inconvenience for the passenger.
By studying Tables 28 - 31, it can be seen that local time is a sjgnificant factor in scheduling and that the higher cruise :Mach numbers offer only limited advantages in arrivatdepartnre times Over the complete route structnre, the higher cruise Mach numbers are expected to provide scheduling advantages and give reduced passenger fatigue by having shorter trip times. However, for the densely traveled / East/West routes, the ad'yantageof higher speeds is less Significant because of local tItimeeff~cts-. This will be investigated further in Phase II.
--~-----
11. 2.3 UTILIZATION. Cruise Mach number has a strong influence on the daily utilization and, therefore, the revenue generating capacity. To determine the time required to complete a typical trip, Figure 103 was prepared. Three cruise Mach numbers were chosen for comparison. The air travel time is plotted for the passenger/range data of 1980 business plus government traffic. The curves begin at the left-hand side with the New York to Europe (PariS) air travel time. Compared with the subsonic jets, Mach 3 reduces the air travel time from 7.2 to 2.8 hours while Mach 6 further reduces the air travel time to 1. 9 hours.
From 45 to 75 percent of the traffic, the Mach 3 and 6 time lines rise slowly. Beyond 75 percent, the curves rise rather steeply due to the longer distances involved, as well as the ti~e penalty associated with refueling which is assumed to occur after 5,500 nautical miles. The bulk of the unrefueled Mach 6 trips can be completed
217
I:-.? >-' IX'
< . ..--.
Cities
lDndDn Madrid, Paris, Rome, Berlin Johannesburg, Cairo Leningrad, Hoscow
Karachi Delhi Bangkok DJakar1~ Hong Kong, Peking, ~il.a. Seoul Tokyo Sydney
Ikmolulu
IDs Angeles
Mexico City, Chicago New York, :~ntreal Santiago Buenos Aires, Rio de Janeiro
~ 'J
IDeal Time (Rel.a.ti ve to lDndDn)
l200 1300 1400 1500
1700 1730 1900 1930 2000 2030 2100 I
2200
0200
0400
0600 0700 0800 OC)OO
Table 28. Local Times
-
I~ ..... <t::
'~
W11"1l it in 1 Ilul,l, in Ulili ,:1 Lj
K.!\HACiII
!!orm KONG
0lDllt:Y •
1.;\
NY
HIO
WNDON
Hot-IE
MOfX;OW .,
K.i\l{j\,:lIl
NOON
O'J()O
0'( )0
0100
;~200
2000
1'(00
l()OO
11100
.
\--..J
-Locul il'illll~
HONe; KOIJG ~;'[DNKi U\ .
.l~OO l' (OC) ~' SO()
NOON 140(; <!OUll
lUUU Noon 18(11)
Oil 00 O(JOO rJOO]J
0100 0300 u;;OO
:2300 ()JOO (!'(('U
2000 2200 oil U()
1~100 2100 0500
1'(00 1)00 0100
Table 29. , Local Tilw's
,
. J
NY HI',.! LOIJlI0N HOt>1f<; MO~JCOW
0, ,h' (Ii, ,J,' 'i( JC! cXlO(1 woo
;'jCJO (11)0 ()1100 0;'00 0700
2100 230() u;~oo 0300 0)00
l;,UO l'(oO ~:uoo 2100 2300
NOON 14()O l{OO .Woo 2000
1000 NOON l:,OU I(JOu 11300
U(UO Ui}O(1 NOON 1300 15(X)
o(,uu O(!(Ju 1100 NOON 1400
().i,O() (J()UO 0')00 1000 NOON
.i
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CITY OF ORIGIN
lAX
ROME
NYC
R<l4E
Departure Data
DEPAR'lURE TIME (ARRIVAL AT OOIGIN AlRPOOT)
0800 1200 1700 0800 l'{OO
.0800 1700 0800 1200 1700 0800 1200 1700 0800 1'rOO 0800 1200 1700 0800 1700 0800 1'{00 0809 1200 1,{00 0800 1200 1'(00 0800 J200 1'700
.
\ ,'-"
Arrival Data
CRUISE MACH DESTINATION CITY DEPAR'ruRE TIME Am '!RAVEL NO. (AT DESTINATICfi TIME, HOORS
... CITY)
0·9 Ra.m 1700 ll·75
t (5500 N .Ml. ) 2100 ~ .
j 0200
3 1700 4.1 3 0200 4.1 6 1700 '2·5 6 0200 2.5 0.9 lAX 230Q 11·75
~ (5500 H .Mi.) 0300
~ 0800 3 2300 4.1 3 0300 4.1 3 0800 4.1 6 .2300 2·5· 6 0800 2·5 0·9 ROO! 1400 1·50
+ {3310 D.Ni.} 1800
+
j 2300
3 1400 2·9 3 2300 2·9 6 1400 1.9 6 2300 1.9 0·9 NYC 0200 3·50
t (3310 N .M1.) 0600 ~ liOO
3 ,I 0200 2·9 :3 oGoo 2·9 3 1100 2.9 6 0200 1.9 6 0600 1·9 6 1100 1.9
.,
'J'abl ~ 30. Typical Schedule - Los Angeles/New York/Rome
'--.-/1
ARRrI AJ... TIME
0445 X 0845 x 1345 X 2110 X 0610 X 1930 x 0430 X 1045 x 1445 x 1945 x 0310 X 0710 X 1210 X 0130 x 1030 x 2130 x 0130 X 0630 X ],655 X 0155 X 1555 x 0055 X 0930 x 1330 X 1830 X 0455 x 0855 x 1355 x 0355 X 0755 x 1255 x
Same Jext • De.y Day
I:\:) I:\:) .--'
~
CITY' OF' ORIGIN
LAX
TYO
Departure Data
DEPAR'ruRE TIME (ARRIVAL AT ORIUIN AIRPORT)
0800 1200 l'{OO 0800 1~!00
1'700 Woo l;~()O
1'(00 0800 1;'00 1'(00 OaOO 1'{00 01300 1'(00
CRUIS!<; MAC I NO.
0·9 . ~ 3 3 3 6 6 6 0·9
~ j
3 6 6
Tat
',,-,,' J
,
Arrival Data ·1 DgSTINATION CI'l'Y DEPARTURE TIME AIR TRAVEL ARRIV AL TIME
(A'l' DESTINA'l'ION TIME, HOURS CITY)
TY'O 0100- 10·3 1120. 1 (4'760 N .M!. ) 0500
~ 1520 X 1000 2020 X 0100 3.8 0450 X 0500 3.8 0850 X 1000 3.8 1350 X 0100 2·3 0320 X 0500 2.3 0720 - X 1000 2·3 1220 X
LAX 1500 10·3 0120 X (4'(bO N .M1. ) lYOO
J8
052Q X
j
21100 1020 X 1500 1850 X 2ltOO 3.8 '0350 X 1500 2·3 1'(20 X 21100 2·3 0220 X
Day Same lIext Before Day Day
Ie 31. Typical Schedule - Los Angeles/Tokyo
~~
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CIl
~J 0 ~ , .~ 8
rl
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ti t: . t .j~l. LlJ~ttl:Hl~r::;_tt/.T-: li:tl-~~"~ -d':'- 1- - - .~ ttt t .,-~. ~-, .t.:", :t- -m . _ . t- .f ~
I t iii, ,Ll! Ii ··1 ,.,., , . jill , ,,,''', '·"H"",jI14 ,II, ;11"' 'l!:i' .".,.,., I·· til - '., ,+ ,Tt ' ill j.. . "
tit ~ lit FH I iH~cfm~ln Ii lJJ. -tlmtii1JtltJHl:Hdt +1:Hll cHirlfr::tji, tlLrr' ,Hl~!P It -111 ~l fUli -fd!trt 'n 'If , H 1) ItFl fn.t·t· ·Ift it . tlllH#'~"·. CrUl~e Mach::: 6.0 1~!t.1 'IJJnl~1 iji·'lfl·r·t I, ·1 ,lllltHl ·tt ·l·j: ·t1. '4.· j.llI11IrU- lit-1 .·t·:1
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10 20 3C 11.0 50 60
f±HttH±H:ttHtH:ffltHtH'tJll+UhlltH1Jtmtlr!rBt I .... UdUUfll1mnui ~ .( .. , H-
70 80 90
Cumulative Per Cent of 1980 Bean Annual TIm-Hay Business Plus Government Traffic
Figure 10~L Trip Time
100
within an air travel time of 2-1/2 hours. This corresponds to a block time of about 2 hours (c~f. Table 27).
Figure 104 shows a measure of daily aircraft utilization as a function of block time and turnaround time. The greatest unknown is the turnaround time required to r:efuel and service a liquid-hydrogen fueled aircraft. If this time is long (i. e. , 4-6 hours)! much potential revenue per aircraft will be lost, and a large fleet of aircraft ~ --
will be required to provide a given level of system service. If turnaround time is
independent of cruise Mach number; e. g., 1. 5 hours, the~Mach 12 cruise will give J' 30 percent more trips per day than Mach 6.0; i. e. , only 75 percent as many Mach 12 aircraft are required to give the same leyel of system service as a Mach 6.0 aircraft. Conversely, the Mach 12 vehicle can have 3;~2. 4 turnaround time, compared with 1. 5 h~~ the Mach 6_!1"Jl"ehic~d still fly the same revenue miles per day.;.:..
At this time, i1Js conjectured that turnaround time for a liquid hydroge~ fue led_
t:':~SP?E!~will?_e~-=--s-=--e-=--n:.::.t:..::i..:.:a:..::ll~y~in:..::d:::,e=-:p,=-e===n=-d=-e===n=-t~o=-f-=c-=-r-=-u=i=-se-=--=M~a:-=c~h=-n::::.u=m::.b=-:e,-,r,--,(beat flux time~tLIDe is approxir:r:tately the same, so !he heat to be removed from the cabin and fuel ta?k __ insulation will be about the same). Figure 104 therefore, shows that Mach 12 will give a Significant advantage in utilization.
It should be noted that a high cruise Mach number will im.prove aircraft utilization more for westbound flights than for eastbound flights. As an aircraft travels West, local time is gained. Thus, an aircraft traveling from East to West will arrive "earlier" in the day. This gives greater flexibility in scheduling flights beyond that point (either westbound or eastbound) as more potential passengers are available during the local hours of 0800 to 1700.
The conclusIOn on utllIzatIOn IS that the hIgher I\Iach numbers show significant gains, with possibly large effects on system economics.
11. 2.4 BEST CRUISE MACH NUMBER. From the above data, it seems that the J * choice of cruise Mach number will not be primarily decided by mission analyses. Technical feasibility, sonic boom and economics will have to be combined with these mission analyses in order to establish a firm basis from which to choose the cruise Mach number. This is discussed in Section 14. O.
11. 3 PASSENGER CAPACITY
An examination of the traffic estimateb from Reference 12 for all twenty-eight area pairs considered by Lockheed reveals that New York City will be the busiest terminal. Figure 105 presents the business plus government traffic leaving New York each day from 1980 to 2000. The dashed curves represent the mean, high, and low daily oneway traffic estimates from New York to Europe (Paris). The annual two-way New
'\ York-Paris-New York traffic figures are converted by dividing by 2 x 365... The solid .~ curves represent business and government daily one-way departures from New York
223
. III 1:\ ......
"' ~ .~ E-i
~ ()
) 0 rl r:q
6 65 4 3 2 1 0 Turn Armmd Time, Hrs.
5
4
~ ..J
2-
Range !II
5,SOOn.mi.
Typical for
Mach 3 cruise
~+ .,
Typicai for . t::!:;.~ . . J: • +. p.~i-
Mach 6 cruise :.;.~:-.,. +
'~-:-t
.-... + - "---:-t-. --: ... .,.
i,~- iti.. ~ , , ; '-LL •.
Typical for
1 - Mach 12 cruise ,,' , r . ,
o o 2
"f~ ~:;r.:t
::+f:j= +i- J-+--r
.ry -4-1--'-++
,~t± Typical turnaround time 1.5 hours • t" " •• , • , •
t
.. t-
~-;
4
utilizat.ioD) One-'viay T~'ips
Figure 104. Uiliz;:uion
224
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1-' \.J: o I-'
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to all eight world areas: Europe, South America, Japan, Oceania, Far East, Mid East, South ASia, and Africa. These solid curves show that the number of daily passengers could vary from a low of 2456 in 1980 to a high of 24,440 in 2000.
Figure 106reiates the number of passengers leaving New York each day to the total number of seats required e~ch day. This figure is valid for mean, high, or low levels of traffic estimates and is independent of any year.
Two levels of patronage for the hydrogen-fueled transport are shown: 25 per-J cent and 100 percent of all business and government passengers. Every departing airplane will not be completely full of passengers so that load factors of 50, 65, 80 and 100 percent are included.
The multiple abscissae show the numbers of aircraft leaving the New York area each day as a function of their size. The number of a given size of aircraft required is the number of seats required divided by the seating capacity of that type of airplane. It should be noted that many of the aircraft leaving New York each day will have arrived earlier in that day and been turned around for another trip.
It is concluded that a large transport is desired with about 300 seats because}" ? this will limit the number of flights leaving New York each day to a ;;asonable value. It is expected,· however, that passenger capacity will be dictated more by economics ~ take-off weight and sonic boom.
11. 4 CONCLUSIONS
... ~ mission analysis ·of a hydrogen-fueled transport operating in the 1980-2000 time period has revealed the following characteristics (these characteristics do not jnclllde
any constraints on take-;:off weight, sonic boom, cost, etc.)
1. Businessmen and government offcials comprise about 37 percent of the long range air traffic between the major population areas of the world. More than half of these passengers will travel between North America and Europe.
2. Based upon traffic estimates, a design range of 5,500 n. mi. is recommended. This will enable 80 percent of all business plus goverllil1ent travelers to be carried to their destinations non-stop. (Consideration of take-off weight, sonic boom, and economics may modify this choice.)
3. Incre~sing the cruise Mach number from 3 to 12' shows rapidly diminishing reductions in trip time.
4. Because of local time effects, no one particular cruise Mach number offers any Significant convenience for passengers.
5. If turnaround time for a liquid hydrogen fueled aircraft is essentially independent ) of cruise Mach number, the higher Mach numbers will enable Significantly greater
revenue miles to be flown per day.
)
)
6. Selection of the best cruise Mach number will be dictated by technology and economics.
227
N tv 00
~ v ,~
One-Fourth of Bus. Go Gov't. All Business and C'"overrunent. Passengers Pass. __ ----~ )l~ ______________________ ~
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Number of 2(;0 PassenGer lIST Ale Required Pel' Day
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Figure lOti, Passengers Leaving New York vs Number of Seats Required
I 125
I 150
)
)
SECTION 12.0 SONIC BOOM
12.1 INTRODU CTION
The sonic boom problem associated with the hypersonic cruise vehicle is recognized a s fundamental in any study of the feaSibility of such 8. concept. On the ba ,:;is of fairly recent work performed by Harry Carlson of the NASA/Langley Research Center, computer program methods have been developed for exploring the sonic boom intensity for various vehicle shapes, particularly from the standpoint of the "near field" effects. This program has been used in this study.
There exists substantial literature on the subject of sonic boom,e. g. , References 14 through 18. The problem is one of subjecting people and things on the ground to an essentially instantaneous overpressure caused by the passage of the shock wave over the ground. A cha:racteristic "N -wave" is developed, as shown below:
For this study, the sonic boom problem re:-;olves itself into the determination of the magnitude of the overpressure for various configurations, speeds, and altitudes, and also how this N-wave changes character during the early portions of the flight (near field effects).
229
)
)
12.2 METHOD OF ANALYSIS - FAR FIELD
In the far field, the various shocks originated by the body are given sufficient opportunity to coalesce into a single bow shock and a single tail shock. The classical theory, originated by G. B· Witham, (Reference 16) has been used to provide a convenient method to determine the intep.sity of the far field overpressure.
The following approximate expressions have been developed to obtain, rapidly, a first approximation of the sonic boom overpressure:
.6P = v
The total overpressure is
(M2 -1 )3/8 W1/ 2
h3/ 4 M Z 1/4 w
+~P 2 1
(1)
(2)
(3)
It should be noted that the K2 of the vplume expression and K3 of the lift expression are functions of the vehicle shape, but vary only through a faIrly limited range for practical vehicle shapes, with K2 = .62 and K3 = .59 being representative values.
The general expression for overpressure presented in Reference 16 has been non-dimensionalized in Reference 18 to yield the following expression:
~ p =
where I (T ) is'the maximum value of the integral of F (7) o
and F (1)
230
(4)
)
')
A machine program has been developed by NASA/Langiey Research Center to determine F(T) and the integral of F(T) using the vehicle shape as input. Using this program, the procedure used to determine overpressure is as follows:
1. The chordwise lift distribution is determined for the vehicle. For a simplifying approximation, it can usually be assumed that the pressure coefficient is constant over the entire wing area, resulting in a simple quadratic expression for a delta wing.
2. 'From J. plot of vehicle cross-sectional area (using normal cub for simplification rather than cuts along the Mach line which would be necessary for an entirely rigorous analysis) versus the nondimensionalized vehicle length t, entet the area at each station as program input.
3. From the lift distribution data of (1) above, enter the vehicle lift at each station as program input.
4. Determine the Mach number range and altitude range for which data is desired, and enter the corresponding C L parameter as program input.
5. The resulting maximum value of the integral of F (i) from the program output is used in Equation (4) to determine the sonic boom overpressure These values are plotted agc;tinst altitude according to the CL parameter used in the program.
12.3 METHOD OF ANALYSIS - NEAR FIELD
When the vehicle is flying supersonically relatively close to the ground, the individual shocks do not coalesce into a single bow and tail shock. The individual shocks which are felt at ground level are, therefore, of lower intensity than the single shock felt for the far field case. Again, the classical theory has been used to analyze this problem. Carlson has found that this near field effect can be measured with wind tunnel experimentation and he has further developed the classical theory to predict these near field effects. Correlation with wind tunnel results has been good. Also, measurements of sonic boom overpressure reSUlting from low flying aircraft have been in good agreement with the analytical results.
The analytical method consists of using the F( T) values generated in the far field calculation described earlier. These values are plotted against the non-dimen-
231
)
)
tionalized vehicle length parameter t, as shown below:
F(T) A
I
A line having the ~lope 1/k(h/1,)1/2 is passed through this curve in such a way that area I equals area II. The intercept "Aft above is then used in the following expression:
~= p
2 K y M F(T)
r (see Reference 18)
to determine the bow shock overpressure. The tail shock is computed in a similar
manner.
12.4 DISCUSSION AND RESULTS
Far field and near field sonic boom overpressures were determined for the following
configurations:
l. Delta wing
2. Variable sweep wing
3. Blended body
a. single delta wing
b. double delta wing
4. Scramjet
5. Supersonic transport
Far field overpressures were found for a Mach number range of 1. 2 to 6.0 and altitude range of 30, 000 to 95, 000 feet. Near field values were determined mainly at M = 1. 4. A vehicle weight of 600, 000 lbs was assumed for all hypersonic configurations. Due to the critical nature of the overpressure problem during the climb part of the trajectory, near field and far field vaiues were examined at a representative climb condition of M = 1. 4 and altitude of 40,000 feet. A cruise condition of M = 6. a at 95, 000 feet was also studied ...
232
)
12.4.1 DELtA WING CONFIGURATION. Sonic boom oyerpressures were determined
for the delta wing _configuration, shown in Figure 3. Overpressure values are shown
in Figure 107 as a function of altitude and Mach number. The dashed curve shows the near field effect at M = 1.4 to give an overpressure of 2.98 psf at 40,000 feet which
is a reduction of 0.27 psf over the far field value. The F (T) function used in determining the near field overpressure is shm\Tn in Figure 108 for M = 1. 4 and 29,250 feet.
12.4.2 VARIABLE SWEEP CONFIGURATION. Sonic boom overpressures for the variable sweep configuration of Figure 59 are shown in Figure 104 as a function of
altitude and Mach number. At Mach 1.4 and 40,000 feet a far field overpressure of 3.65 psf is obtained, as compared with a near field overpressure of 3.50 psf. This configuration differs from the delta configuration as far as sonic boom is concerned, primarily in the lift distributions, which are shown in the following sketch.
Unit Lift
Delta Configuration
Length of wing
Unit Lift
Variable Sweep Configuration
600,000 lbs
Length of wing
The F (T) function for the near field calculation is given in Figure UO.
12.4.3 BLENDED BODY CONFIGURATION. Two blended body configurations with
different wing planforms as shown in Figures 70 and 71 were analyzed for sonic boom overpressures. The results are shown in Figures 111 and 112. Overpressure values
for the far field and near field of 3.22 psf and 3.00 psf respectively were determined for the Single delta planform configuration, as compared to values of 3.35 psf and 3.0 psf for the double delta planform configuration. These values are similar to those
obtained for the delta configuration. Typical values of the F (T) function for the blended body vehicles are shown in Figures 113 and 114.
12.4.4 SCRAMJET CONFIGURA. TION. Sonic boom overpressures for the scramjet configuration shown in Figure 84 were determined and are presented in Figurel15.
233
)
)
o 1
7
6
~ 5 rn ..... 2
0.. 4 <l
2 Q) I-< 3 ;j CI) CI) Q) 1-<' 0- 2 I-< Q)
> 0
1
0
o 10 20 30
------ -------------.------_.-
• .. H-4- ... o-~ • J • ~ ?
+~ .... -t.,j, ..... ..~+-t·~ ...... .4--. !: :i. I'" ~~-+- --t.l-t" ;. . !J t.l. .. +
" +- +++ ~ .:. .;
... " ~ ~ .. ;. I .. fl ...
tt.ti ,a..+t- i .... +> .. t' r!i: -!!~ ~
2 3 4
Mach Number
40 50 60
Altitude ~ 1000 ft
~t~ ;l .. . ..... !1ti· j._ ••• l .... .l._.
~i~] :! t: "'t 1 .,
~#! .. , ;
5
70
1~+: n:t i tl.-:
..-i-• .J.
~ f-:!
6
70,000 ft 80,000 ft 90,000 ft
80 90 100
Figure 107. Sonic Boom Characteristics - Delta Wing Configuration
234
r
)
)
1 Slope = ---=----
K (h/2)1/2 = .01598
where K =
K = 6.7
{y + 1) ~
/2 (33/2
o .1 .2
?\ear Field Effect
~ p = ~ y M2 F (r)
p ff~ (h/e )1/2
~p
p .00378
P = 1173
~P = -1. 4:3 PSF
.3 .4 .5 .6
t - x/i,
Figure 108. F(r) Function - Delta Wing Configuration
235
.7 .8
)
)
)
7
6 +
~ rn 5 ~ l + ~
4 <3 +
I (1) ~ :l 3 til til (1) ~
eo 2 Q)
~ 1
ti .. ++1 •
0 o 1
6 +-
4 ~
3- Near @M
2
i
o o 10 20
.w-
~+tt .. :¢ ±tt •. , >~ .
t!+i- ++ ~~. f$l Hn :PH ill =p:p:
... ~n: ~:±p +
• .++. 't H+"" i++t+ +-> . ,. "-' ~ Alt . 40,00
~ • ++ f++-H
o H+
fHH t tint fH + ... +' +~it ~ tffil . . tht Ht
+ • +Hr
'+-Ht . HH ~
.+ . , ~+ .... ~ . +H ·I ~
., .H.+
f- t"; ii·i,
.I .+ Hj .: +H~ffi:
, t '±i±!;
f~ltt
2 3 4 5
Mach Number
30 40 50 60
Altitude - 1000 ft
,
-t+++ ::@.
l±= 4 ..... .4.-+ ....
tit ,-\i+
~~ ±t±:± -::tti
r& t!-
6
70
50,000 ft
60,000 ft
70,000 it 80.000 ft 90,000 ft
80 90
Figure 109. Sonic Boom Characteristics - Variable Sweep ConfigUration
236
100
)
Slope --: .0161
\1:= 1.4 [(129.250 [1
o .1
Near Fie ld Effect
6P
P .00·11:;)
r 117:1 [}P - ,).:2:! P~F
.3 .4. .5 .6
Figure llO. F(T) Function - Yariable S\\·cep Configuration
?:17
I ')
)
J
7
6
~ 5 0.-(
0.- 4 <l
Q.) ... 3 ~ Cf.l Cf.l C) ... 0.. 2 ;;.., C)
b 1
0
o
7
6
~ r.n 5 0.-?
0.- 4 <l
Q.) ;;.., ;:::: 3 Cf.l Cf.l C) ;;.., 0.. ;;.., 2 Q.)
:> 0
1
0
o 1 2 4 5 6
Mach Number
.. ~.~ ~"-i,- r:-c-tTf:: ~~~~~l~ -~~:J~~ ~~~]~-£~~::: ':J~f+?
70,000 ft 80,000 ft 90,000 ft
" .~h+ ~~-;-;-'i~~~~-::~:::~ ~~~J:;i;~~~~~~:~;-
10 20 30 40 50 60 70 80 90 Altitude ~ 1000 ft
100
Figure 111. Sonic Boom Characteristics - Blended Body Configuration Single Delta
238
f
012 4- G
:\Iach ~umhcr
)
o 10 20 30 ·10 50 GO 70 80 90 100
Altitude ~. 1000 fl
) Figure 112. Sonic Boom Characteristics - Blended Body Configuration Double Delta
239
(
)
)
K = 6.7
Slope = .0150
M- 1.4 @ h~ 29,250 ft
. 018
• 016
.014
.012
. 010
i=' -~ .008
.006
;;~
~ 0 .1 .2 .3 .4
t = x/Po
Near Field Effect
tiP --=.00396
P
P = 1173
tiP", 4.65 PSF
........
.5 .6 .7
Figure 113. F(r) Function - Blended Body Single Delta Configuration
240
(" ~-.- .---- .. -. .----~--------------------------------------------------------
K". 6.7
Slope: .0150
Near Field Effect
M -,1.4 @ 29,250 ft 6P
P .00392 -- ..
..... "'" .-.......... ... :, ............ . i4 • ~ .... !-
"fl' ;.""; • L • .L .... +-
' ••• r"" .- ... ~ .. ___ •• , •• 1 •
• " . .;...!J 1:--,.
.0150
)
o .1 • 2 .3 .4 .5 .6 .7
t ". x/t
J Figure 114.. F(T) Function - Blended Body Double Delta Configuration
241
{ I
)
)
7
6
~ en 5 ~ !
~ 4 <l
! (lJ 1-< ;::1 3 en en N (lJ 1-<
@ 0. 1-< 2 (lJ
> 0
1
o o
7
6
~ en 5 ~
!
~ 4 <l
(lJ 1-< ;::1 3 en en (lJ 1-<
fr 2 (lJ
> 0
1
0
0 1
.... " , .• , . I; ,. I I
. idt., - ~O, 000 ft
+
f+-'-4-<- ;::;:~1::r;:t -'-" r++++-
2 3 4
Mach Number 5 6
50,000 ft
60,000 ft
70,000 ft 80,000 ft 90,000 ft
2.0
i-i-+- ..;;; 1.4 ,-c-; I-iM- 1.2
, ,
ear Field Effect M- 1.4
I
10 20 30 40 50 60 70 80 90 Altitude - 1000 ft
Figure 115. Sonic Boom Characteristics - SCl'amjet Configuration
242
100
f
1
)
)
--_._---" ------------_._-
At the typical climb conditions of Mach 1.4 and altitude of 40, 000 feet far field and near field overpressures of 3.42 psf and 3.05 psf respectively were obtained. This configuration is similar to the blended body configurations and therefore exhibits similar overpressure values. The F(T) function for M = 1. 4 and 29,250 feet is shown in Figure 116.
12.4.5 SUPERSONIC TRANSPORT. For the purpose of comparison a typical supersonic transport design was used in the analysis at a weight of 450,000 lb. The resulting overpressures are shown in Figure 117 as a function of Mach number and altitude. At M = 1. 4 and 40,000 feet a far field overpressure of 2. 29 psf was obtained and a near field value of 2.25 psf. These values are in general agreement to those shown by Carlson in Reference 19 for a similar supersonic transport design where he indicates that for a 400,000 Ib vehicle/far field and near field values of 2.5 and 2.2 psf are obtained at the same climb conditions.
12.4.6 CO.MPARlSON OF CONFIGURATIONS. The far field and near field values of sonic boom overpressure at a representative climb condition of M = 1. 4 and 40,000 feet for the configurations investigated are shown in Figure 118. There is little difference in the sonic boom overpressures for the delta, blended body and scramjet configurations, however, the variable sweep configuration shows higher near field and far field overpressures. This is thought to be primarily attributed to the differences in lift distributi'On between the variable sweep and delta planform configurations. The supersonic transport is also shown. As can be seen, the hypersonic transport configurations show far field and near field overpressures that are, respectively, 1. 0 psf and 0. 8 psf greater than the supersonic transport configuration.
Far field sonic boom overpressures for a representative cruise condition (M:: 6.0 (a)
95,000 feet) are also shown on Figure 118. There is little difference in overpressure
among the configurations studied with b. P of 1.25 psf being a representative cruise value.
12.4.7 CONFIGURATION VARIATION. In order to determine the effect on sonic boom overpressures of variations in configuration, the delta"configuration was chosen as representative and the following parameters investigated:
1. fineness ratio 2. body shape 3. wing loading 4. wing aspect ratio 5. take-off weight
The effect of each of these parameters is discussed in the following paragraphs.
12.4.7 .1 Fineness Ratio. Body fineness ratio was varied from 7.25 to 15.5 _about
the baseline configuration fineness ratio of 10.9.
243
M = 1.4 @ 29,250 ft Near Field Effect
K= 6.7 llP .00396 -= p
.020 P = 1173
.01" llP 1173 (.00396 :w 4.65 PSF
.01l!
.014
.012
.010
.00 0
) .006
o .1 .2 .3 .4 .5 .6 .7
t :: xli
Figure 116. F(T) Function Scramjet Configuration
) t-• 244
I
! -)
(.J
p.. <l
7
G
5 ~ [fJ p..
I 4 p.. <l
') C,) v .... ;j {fJ (fJ C,)
2 .... 0. .... ill > 0 1
°
:::L:~'. ::: j"i __ ~_ ~r .. :[:: -+-: _:-J.:j_:~_::+._;:_: .... 11 : ....... oj..:.\ l~. ::' 3-01""-, (--)(--+)0-'--f~~t_.-. =:.=-:=:1:,'=::::
:.:.' :.' '.1'.' :.:. ;.:.::::.;V' .:.:::. i :. i J.. J t :.;;:~~ .. I. . .• _ .. , .. ' '. . .-
: I :: ;-~:::... ' ~ : 40, 000 ft \ . . ;.,L • • ~.--=50,OOOft--l::lfij~' __ .~_ ~ :60, 000 ft ~._
'11 ... : '.: ~. I : --: \::: 70,000 ft .. . :i· Ii i .. I . . : 1 :::.. . t
l, I .: I
I" I I I. , +--- . ·-·-t- -- __ ._1 ~-
.. j......... I .. ' ;: :.;: !
0 ') 6
... ... ':'::, ::, .... : il ,..:', , . I 7 _.--:.,:::: .:.(. . , ~ I I
L-~-4--+--+--+-~~~~: : : i I .- .. . . j. :-r-:---:-'" ---4·--- -: 1-' ---!-----I-_. , .. . -+. , -.. -. +--t---..t
6 -... :.::: ....... ". .::: I: Ii' I . -\
5 t-~-.~:+-: -: : :+:-; :-+ .•. -.' .-tOo -: . .....j:: 11-:-; ;+-' :.-' "+--V7-1 +-, +_ ...... :-. -t-. -!I-+---.-:-.: . -+. ---Li-': .~:.: ~~~i~~~ ~:~ :;~; ;.:i~ :::~ ··--1,· 11·\-:-:
Figure 11 7. Sonic Boom Characteristics -- Supersonic Transport
24~
------~--,
.~ v J
Overpressure,..,.. 6P ,..,.. PSF
o ..... 1:\:1 c,., tJ::.. c:.n 0) -l
~ ...... ~ Delta Configuration I-j C1> ..... l-' 00 . en 0 ::s ..... ()
to 0 0 s
I ~ ~ 1/ 1/ 0) ..... . . 0 tJ::..
Variable sweep8 @ @ Configuration ~ to Z "rj tJ::..
Jl) c:.n C1> Jl) 0 I-j
~ I» I-j ~
0 I-j "rj 0 ":tj 0
~ 0 .....
0 ......
0 C1> ..... C1> (") ::r'
~
..... ~ C1> ..... .....
0.. ..... 0.. ..... 0..
Blended Body Jl)
C\:) () .....
tJ::.. C1> Si ng le~~~! t_~_g~~!~!:ati on
m I-j ..... 00 ::1: () 00
I Blended Body (") 0 Double Delta Configuration ::s ....., .....
(Jq t: I-j Jl) ..... ..... 0 ::s (')
SCRAMJET Configuration 0
~ Jl) I-j ..... 00 0 ::s Supersonic Transport
If
( ) The total volume was held constant while the body length and cross section were varied to provide the fineness ratio variation. The far field and near field sonic boom overpressures are shown in Figure 119 as a function of fineness ratio. Increasing fineness ratio produces a slight decrease in both far field and near field overpressure.
)
12.4.7.2 Body Shape. The effect of body shape was investigated by holding the total volume constant and reducing the maximum cross sectional area and redistributing this volume in the front of the vehicle, producing an effective cross sectional area distribution as shown in the following sketch.
Effective Cross Sectional Area
reduced maximum
Vehic le Length
______ --"base line"
The maximum cross section was reduced 10 and 20 percent, which, when redistributed in the forward part of the vehicle effectively "blunts" the vehicle. This effect on sonic boom overpressure is shown in Figure 120. There is little effect on the' far field overpressure; however, bluntness is seen to produce a significant reduction in the near field overpressure, wIth most of the reductIOn occurring over the first 10 percent reduction in cross sectional area. This is a significant effect in that near field sonic pressures of less than 2 psf are realized. Carlson, in Reference 19, records a similar effect for the supersonic transport.
The resulting increase in take-off weight due to an increase in body bluntness and consequently the increase in drag is shown in Figure 121. For a 10 percent reduction in maximum cross sectional area, the take-off weight increases by 2.7 percent while the sonic boom overpressure is reduced by 41 percent.
12.4.7.3 Effect of Wing Loading. The effect of wing loading was determined by holding the weight fixed at 600,000 lbs and increasing wing area. Overpressures for wing loadings of 60 and 70 psf were determined in addition to the "baseline" vehicle with a wing loading of 80 psf. Figure 122 shQws there is negligible effect of wing loading on either the far field or near field overpressures. This is attributed to the small change in effective cross sectional area distribution associated with these wing loading variations.
12.4.7.4 Effect of Aspect Ratio. The effect of wing aspect ratio on sonic boom far field and near field overpressures are shown in Figure 123 to be small. Wing aspect ratio was varied from 1. 0 to 2.0 for a fixed wing area. The wing position relative to the body was also held constant.
247
I\:l
'''" 00
'---'
~ rn Il1 l
@j
Q)
6
5
'4
~ 3 rn Ul Q) H
e 2 Q)
6
".,..,../
Far Field . . ___ . Near Field
.& • Baseline Conf guration M :: 1. 4 @ 40, 000 t <
..
,J ... , + '. :j:'" ' .. . f' fftJ . . ,
... "1:
IlJ s: .... ,
rt- ,.
t. ,.
.,
, .
i :1' r
J i
It 1· ~ J. t
:rJ I ., ..
'~' r ilfH~ it' t j . t . Ht '+ rl 1r :t. . . +-' ' ·t'
. .. rtfl!f rTf" . Hi' , :t . . Itti .. ,j + . i .j., .. J ~r, . -1
·l :1' .• .. t.· ,;: +' t, ,.
5 6 7 8 9 10 11 12 13 14 15
Body Fineness Ratio
Figure 119. ~ffect of Body Fineness Ratio on Sonic Boom
J
o'
H .. In . ' ~'ntjf! tl-
.,
. ,
16 17 18 19 2 o
.. ,
..
)
)
700
600
TAKEOFF WEIGHT
(1000 LB)
500
o 10
Percent Reduction in :VIax. Cross Sectional Area
Figure 120. Effect of Body Shape on Sonic Boom
~ /l ~ .-
<: ~ " '\J 7
~2~:::=:::: -=--~r-~~~t:~;_ i~~~: ;~::~t:::: ;-::.- -,'1---:- -:- ,-r-:---'" - :~.r:-:: ::-:-: :~:: :~~: -=-:=-G:--:::
f----"- :-::::. ~:::- :-. -.. . •. : r:- : ..: :: -- t ...:.~: ::.: .. :: : -:: -:. . ... .. - -
~~~~~~;~~J2~~~ !i:~ _~1~~ ~~~ :::~ ;~:~ ~ ;~: J~ E : ~.-. ::: "~ ~~_:;;~: :::?~: ~~J:~::::: ~¥~~-:
:::: ;:: :::-~ :~~? :-~ ~~~:~:~K~ ~: :f:~: ~:::r:: ::t< :: :~r~j~'1 2. 9 ~~ ~ ~~:- -~ ~~J 1. 7 :::0 :.::>~:: 1 ::. : - 1. 6
~=-:l?:.~ ~~~:~:~-:::-:;-:::::;~t=~~f:== -~=::::: ~::ir=: :~~~J I
o 10 20
% REDUCT! ON IN MAX. CROSS SECTION
Figure 121. Effect of Body Shape on Take-off Weight
249
20
SONIC BOOM ( PSF)
~ rn 0.;
( 0.; <l
~ Il) ~ ;:l [J) [J) Il) ~
0-
) ~ Il) :> 0
J
Delta Configuration
Far Field
_ _ _ Near Field
7
6
5
4
3
2
1
W". 600,000 lb
£ • Baseline Configuration M = 1. 4 @ 40, 000 ft
, .
o ,llllll iii II ! i III II! II i ! ! IIIIII111 ! III i!lli i! I i I ii Iii i i! Iii! II! I! Ii! I iii Iii I! ! i ! II! ! ! I! i I ! Iffl#l 60 70 80
Wing Loading ~ W /S '" Ib/ft2
Figure 122. Effect of Wing Loading on Sonic Boom
250
)
)
Delta Configuration ___ Far Field
-Near Field
.& • Baseline Configuration M = 1.4 ((1.; 40,000 ft
7
6
5 ~ lfJ p... 1 4 p... <l
Q) 3 ... ~ CIl CIl Q)
2 ... 0.. ... Q) :> 0 1
° 1.0 1.2 1.4 1.6 1.8 2.0
Aspect Ratio
Figure 123. Effect of Aspect Ratio on Sonic Boom
251
)
)
12.4.7.5 Effect of Take-off Weight. The sonic boom of the delta wing configuration for take-off weights of 700,000 and 800,000 lbs was calculated. Figure 124 shows that increasing take-off weight increases both far field and near field overpressures, with the overpressure being roughly proportional to the square root of the weight.
12.5 MISSION SONIC BOOM CHARACTERISTICS
The sonic boom overpressures experienced along three typical cruise trajectories were determined and are shown in Figure 125. A 5,000 n. mi. cruise ra.."lge was assumed at cruise Mach numbers of 3.0, 6.0 and 12.0. A 600,000 lb take-off weight blended body was used as the vehicle. The figure shows far field overpressures of approximately 2.9 psf during the first 150 miles after take-off and overpressures of approximately 2.25 psf just prior to landing. For all three cruise Mach numbers, overpressure levels during cruise were less than 1. 5 psf. It should be noted that the values shown are far field values and, therefore, near field effects during climb and letdown will reduce the overpressures somewhat. For instance, as shown on Figure 112 at M = 1. 4 and 40,000 feet, a reduction of 0.35 psf is realized. Although values of overpressure of less than 1 psf are shown during cruise, it should be pointed out that the theory used in the analysis at these hypersonic Mach numbers has not been verified to date with experi mental data.
12.6 CONCLUSIONS
The following conclusions have been reached as a result of this study:
1. The far field sonic boom overpressures for hypersonic transports are approximately 1 psf greater than those obtained for the supersonic transport during the climb segment of the trajectory.
2. There is little difference in the overpressure level for the configurations considered. The variable sweep configuration had the highest overpressure and also the least near field effect.
3. Aspect ratio, wing loading and fineness ratio have little effect on overpressures. These parameters will therefore be selected for minimum take-off weight or to meet runway requirements for take-off and landing.
4. Bluntness of the nose of the body has a pronounced effect on the near field overpressures. Near field overpressure levels of less than 2.0 psf were obtained for the representative climb condition, with only a small penalty in take-off weight. This is an area for futher study during Phase II.
252
5
4
3
,) 2
o 600
:-+*
'TI-+
700
Delta Configuration
Far Field Overpressure
'++t r-"
r+-' I
r ,..;.
-j-1-
:;..::± --T+!- --1=+:;:!:
~. 0+:
i-+---t-i::!; . .;.::±::. ' +-i---i- ~-:-
. f+-7 ,-"-'-. -1-,. j:~ .:..::r:: ~,+- ++-:- ltt;: .-:J:::tt
800
Takeoff Weight ~ 1000 Ib
Figure 124. Effect of Takeoff Weight on Sonic Boom
253
\,--, \-/ J Blended Body Configuration
Mach 3.0 Cruise @ 80,000 ft
___ Mach 6.0 Cruise @ 110,000 ft 4~nn~~~~~~~~~~~~~~~~~rrh~'~nn~~~~~~~~~~~~~~~~~~~~~~~~~
~3 P-4
P-i <l
? Q) H ~2 Ul Ul Q)
H I'V P. 0. H ..,.. Q)
?
I 0
1
:tillllflHtlHlllliHllnmllUlHtlnHUUttHlHHlltHUHHtnnnJlllmtHlHltJUlfnalWnnmH1HlHtH1i
-
o~~ww~ww~~~~ww~~~~~~~~~~~~uw~~~~~~~ww~ww~ww~~~
o 1000 4000 5000
Range - n. mi.
Figure 2. rYrission Sonic Boom Characteristics
A e
A(t)
A " e
B(t)
d
F(T)
F' l
h
I(T 0)
) Kr
K2
T.T
l'..3
i
M
DP v
DP i
6P . t
6p
P
P a
p 0
q
)
12.7 NOMENCLATURE
A(t) + B(t)
2 Cros:-;-sectional area/ ~
Second derivative of A with respect to t e
2 q:Cx2 f; , dx o i
Diameter of equivalent body, ft.
__ 1_ j T Ae" d ~ 2rr 0 FT-~
Lift per unit vehicle length, lb/ft.
Altitude, ft.
Maximum value of integral of F(T)
Reflectivity factor (taken as 1. 9) = (y+ 1) M4//2 S3/2
Vehicle shape factor in volume expression (.62 is representative value)
Vehicle shape factor in lift expression (.59 is representative value)
Vehicle length, ft.
Mach number
Sonic boom overpressure due to vehicle volume, psf
Sonic boom overpressure due to vehicle lift, psf
Total overpressure, psf
Total overpressure from general solution, psf
Reference pressure = -rp-T, psf ""a -0 Ambient pressure at flight altitude, psf
Ambient pressure at ground, psf
Dynamic pressure, psf
255
t
w s y
Tand t..
NOMENCLA'IURE (CONTINUED)
xl t, non-dimensionalized distance aft of vehicle nose
Vehicle gross weight, lb.
JM2 - 1
Ratio of specific heat for air (1...: 4)
Dummy variables of integration, in same units as t
256
)
13.0 COST
Although for a hydrogen fueled commercial transport costs are a fundamental consideration, during Phase I, the extent of the cost analysis was to help in the selection of the configurations to be used during Phase II, rather than to be a detailed analysis of a hydrogen fueled transport. The analysis was, therefore, based on extrapolations of the supersonic transport, XB-70, and subsonic jet transports and not on engineering and factory estimates or specific route combinations for a particular design.
The main effort during Phase I was to establish the methodology and obtain typical values. A computer program was established to enable these comparative costs to be obtained rapidly.
13.1 COMPUTER PROGRAM
Reference 24 describes an approach to a cost program for multi -stage boosters that uses a weight/sizing routine to "drive" a cost program. Changes in a performance parameter which changes the weight and size of the vehicle is then reflected in changes in .cost. By using the weight/sizing "parts list" plus appropriate mathemati cal relationships, changes in a performance parameter can then be evaluated in terms of its cost. This same approach was used for the hydrogen fueled transport cost model.
Figure 126 shows the flow of the program which consists of seven submodels, as follows:
1. Weight/Sizing. The output of this sub-model is a self consistent list of weight and size data. This sub-model is the same as that used in the over-all vehicle synthesis program and which is described in Section 3.2. 1.
2. First Unit Cost. The output of the weight/sizing sub-model becomes the input to the first unit cost sub-model, and uses cost estimating relationships such as dollars per pound or dollars per square foot to determine the cost of each component. The aggregate of these costs results in the first unit cost. This estimate is re-evaluated as required to be consistent with changes and refinements in the vehicle's design.
3. Research, Development, Test and Eva Iuation Cost. The first unit cost is then input to the RDT and E sub-model according to the requirements of a deve lopment
257
)
)
WEIGHT /SIZING SUB-MODEL
FIRST UNIT COST SUB-MODEL
RESEARCH, DEVELOPMENT, TEST AND
EVALUATION COST SUB-MODEL
NON-RECURRING COST SUB-MODEL
RE CUR RING COST SUB-MODEL
OPERATING COST
SUBMODEL
BASIC PARAMETERS SUB-MODEL
Figure 126. Cost Program, Flow Schematic
258
)
j
schedule. This schedule is organized to reflect the manpower and materials requirements necessary to attain some predetermined operational goal; e. g. , 0.95 mission reliability, Mach 6 speed, etc. A typical development schedule would show the manning and material requirements for appropriate time intervals of development for each of the vehicle's components. The aggregate of the RDTE cost elements is the first major system cost component, and the completion of this phase of system development is a prerequisite to the operational phase.
4. Non-Recurring Cost. For the operational system, the RDTE facilities are expanded and new ones added to meet the needs of the operational system. This expenditure constitutes the second major cost component, or non-recurring cost, and is estimated as a function of vehicle size, complexity and the projected use rates of the vehicle.
5. Recurring Costs. The final component of system cost is the recurring, or operational costs. This cost is based on estimated schedules for use and production necessa.ry to attain the ;mission objectives. that is, the transportation of a quantit.y of cargo and passengers between points A and B within the operational constraints of the vehicle.
6. Operating Costs. The accumulation of the annual operating costs over the system lifetime is the total operational costs and the sum of the RDT and E noh-recurring and recurring costs is denoted as the total system cost. The operating cost is :si.mply 3 + 4 + 5, with the addition of certain administrative and operational expenditures associated with a commercial transport operation. The resulting total cost structure is divided into two parts--indirect and di rect costs. The indirect costs are the RDT and E and administrative costs associated with the operational sys-
irec cos s mc u e a e recurrmg an non recurring costs accumu lated in the previous models with the addition of insurance to account for losses of primary mission equipment.
7. Basis Parameters Model. This prints out the assumptions that are made and en-· abIes the results to be interpreted qUickly and clear ly.
13.2 ASSUMPTIONS
In order to establish approximate costs and so help in the selection of the Phase II configurations, several broad assumptions were made:
1. A successful supersonic transport is assumed as the technological baseline from which to deve lop the hydrogen fue led transport.
2. An independent turboramjet engine design and development program.
3. A 15-year system lifetime.
259
)\ / )
»)
))
4. Typical development and production schedules.
5. An operational fleet-size of 50 vehicles is accumulated over the system lifetime to the year 2000 A. D.
6. "Learning" effects are considered on vehicle use and production rates.
7. Scaling effects on vehicle weight and size (volume and wetted area) are used when appropriate.
8. Total flights for the system are based on a vehicle having an average block-toblock speed of 2500 n. mi. per hour and an 8-hour per day use.
9. Mission re liability is assumed to be unity.
10. Vehicle losses and replacements are accountable in the form of 100% replacement insurance for the primary vehicle and ground support equipment. ( Current jet transports use 12% initially, reducing to 5% at the end of 15 years. 17% to 8% were assumed for the hydrogen fueled transport to reflect higher state~f-the-art.
n. System costs reflect the operational and implementation development for one citypair only.
13.3 TYPICAL COST
USing the delta wing baseline configuration as a typical example, the print-out of the computer program is shown in Table 32. This table shows the print-out for each of
iscllssed'in Section 13 1 Included in the operating cost prjnt-ollt are three cost effectiveness parameters; i. e., cost per flight hour, cost per flight mile and cost per seat mile. These are shown for both the direct cost and for the total system cost. Direct operating cost as computed by the 1965 A TA formula is also shown on page 264. The DOC computed by either method is similar. The Convair method is about. 1% higher because of the inclusion of some ground fueling facilities.
As mentioned previous ly, the values shown in Table 32 were obtained from extrapolations of the cost of XB-70, SST and subsonic jet aircraft, solely for the purpose of helping in the selection of the Phase II configurations. A more rigorous
analysis will be made during Phase II.
13.4 COST SENSITIVITY
To help in the selection of the Phase II configurations, va riations in three ba sic parameters were investigated; i. e •• range. passenger capacity and fuel cost. The results are shown in Figure 127. This shows: (a) that DOC is very dependent on the cost of the hydrogen fuel and (b) that the knee of the DOC curve occurs at approximately
200 passengers and 5,000 nautical miles.
260
ROT+E COST $M 4670.36 /J tESIGN AND DEVELOPMENT 1600.94
/ AIRfRAME 495.43 PROPULSION 705.01 ASTJ~ {ONICS 26.44 MISCELLANEOUS SUBSYSTEMS 316.25 SUPPORT EQLIPHENT 12.50 SYSTEHS ENGINEERING 45.32
TOOLING AND SPECIAl EQUIPMENT 35.6Q AIRfRAME 35.60 PROPULSION o. ASTRIONICS -0. MISCELLANECUS SUBSYSTEMS -0.
SUBSYSTEMS TEST HARDWARE 2074.68 AIRfRAME 200.01 PROPULSION 1334.01 ASTRIONICS 6.58 MISCELLANEOUS SLBSYSTEMS 485.62 SUPPORT EQuIPMENT 48.47
SUBSYSTEM TEStiNG 65.74 AIRfRAME 25.90 P~OPULSIGN 16.40 ASTR IONICS 0.94 MISCELLANEOUS SlBSYSTEMS 11.25 SUPPORT EQ~IPMENT 11.25
STAGE STATIC TEST HAROWARE 177.72
) AH~FRAME 66.67 P~OPULS ION 22.23 ASTRIGNICS 3.29 MISCELLANEQUS SLBSYSTEHS 69.37 SUPPORT EQUIPMENT 16.16
STATiC TEST ANt ACCEPTANCE 41.51 OPFRATION 7 'iO
TRANSPORTATION 0.00 PROPELLANTS 2.51 TEST SUPPORT 31.50
FLIGHT TEST HAROWARE 511.01 AIRFRAME 200.01 PROPUL SIGN 66.10 ASTRICNICS 9.86 MISCELLANEOUS SlBSYSTEMS 208.12 SUPPORT EQLIPMENT 32.31
vEHICLE FLIGHT TEST 151.16 OPERATION 28.12 PROPELLANT S 117.20 TRANSPORTA T ION 0.00 RECOVERY o. RECONDITIONING 5.84 M ISCELLANEOU S o.
(J Table 32 (cont'd). Typical Cost - RTD & E Costs
262
)
)
OPERAT leNS CCST HARtIoORE
AIRFRAME PROPIJL S leN ASTRICNICS M1SCELLANECUS SUBSYSTEMS
ACCEPTA~CE OPE~ATICNS PERSCNNEl PROF ElLAN T S SliPPOR T
LAUNCH OPERATI(~S PERSCNNH TRAI\SPORTAIICN PROPELLANT ~ SIJPfOR T
RECOVERY OPERAIIChS IlECONCIlIONlNG
INSPEC TICN REPLACEMENT EXPENDABLES wEARGLT/OVERHAuL/REPlACE upon ING SFAR ES
SUSTAINING ENG[NEERING AIRFRAME PROPULSION ASTR ION Ie S MlseELLANECuS SIJBSYSTEMS
TOOLING AIRFRAME PROFI..L SICN ASTR leN ICS MISCELLA~ECUS SLBSYSTEMS
~ I SCElLANEOlJS
FAC IL IT lES ~ANUFACTURING fACILITIES
AIRFRAME PROPUL S leN ASH ION res M!SCELLANECUS SLB5YSTEMS
TEST FACILITIES . ~ ,~ ... ''-
PROPULSION ASH ICNICS MISCELlANECUS SLBSYSTEMS STAGE TEST FACILITIES
(EVELOP~ENT LAUNCh FACILITIES PAD RELATE£: ASSEMBLY BLILDING ANt CHEC LAUNCH CENTER RELATED AND G~O~ND SUPPORT EQUIPMENT
OPERATICNAl LALNCh FAC. A~D EQ PAC RELATEC ASSEMBLY BLIlDING A~D CHEC LAUNCh CENTER RELATED AND I GRCLN~ S~PPORI EQUIPMENT I
RECovERY ANI: RECGt\OITICNING FA~ RECCVERY FAC. I RECCNDITIOhING FAC. I
~ISCELlANEOUS FACILITIES TRAINING I SIT E AC TI ,.AT ION [
Table 32 (cont'd).
i I I I
I I
-I Typl
I I
I
,
)
OPERATING Cd>TS
MAINTENANCE PRIMARY VE~I.ClE
SLPPORT EQUIP~ENT
fACILITiES fLIGhT OPERATIGNS
fLIGHT PAY ANC ALLOWANCES PROPfllANT,Oll, LLBRICANTS
CAf ITOL ACCOlNT 5 INSURANCE (EPRECI~TI(N
TOTAL SYSTEM ~IRECT AVERAGE A~NUAl OIRECT·
IN_DIRECT COSTS SALES AND SERVICE
, ADMINISTRATION ReTE
TOTAL SYSTEM INDIRECT
TOTAL SYSTEM COST
COST EfffCTI~E~ESS COST PER fLIGHT tR(UCLlARSi COST PER FLIGHT ~IlE(D(LlARS.
COST PEK SEAT ~IlE (CENIS)
A TING COST AS CALCULA TED lULA (1965) ,
CPSM
trIONS 5.409 W Ca:lTS 0.027
3.997 r;a:lTS 1.385
iENANCE 0.159 IPMENT 0.t59
I
0.826 I 0.826 IPMENT I EASICPARAMfTERS SYSTE:M LlfE
MISSION RELIABILITY 1 [ING COST LESS (RDT&E) 6.394 I
RClTE TRIP TIME fLIGt-:T TIME:
GRGtNO M~NEuVER TIME LTILllATICN USEfUL LeAC CARGO NU~BER Of PASSE~fEKS
LC,:lC fACTCR CR E w I.
FLIGHT CREW
ATTENDANTS CESIGN PRCPEllA~1 FLIGHT PRGPELLANl GReSS TAKE-Off ~T.
e A SIC 11'4 V EN TOR Y
EX,RA vEhICLES REPLACEMENT VEHICLES
TOTAL NLMEER VEbICLES FLIGHTS PER YEAR
TOTAL FlIGrTS TOTAL PAYLOAD
1.(C0001 .679 I 1.(.CO(,OijrING COST INCLUDING RDT&E 7.073
5COO.Q 2.0~
.1.75 I 0.25
8.0 n
5001C.O 1 '1-' AI (8 I 13(.00.0 ut! lzatlOll per C hrs day) 16t; • 0 tion period
O. 8 abin crew 8.0 H2/flight @ $.20 per pound
4.00 ~perA/C
~le city-pair range 4. CO ~hic1e ~~-g-' -" ~LL 1~:~~r~_to_-_ru __ OC_k __ s~ __ e_d ________________ ~1 5'.1 9419.9,
1 .~)( 42.0l o. G. O. i
50.0j 1460.0
61320.0 6e7660.0~
34301.8
Table 32 (cont'd). TYP1i
264 I
,.
-Q) .... ..... e .... ro Q) [Jj
,.... Q) 0. [Jj .... c: Q) u -.... [Jj
0 U be c: ..... .... ro ,....
) Q)
~ .... u Q) ,....
Ci
12
10
8
6
2
o 4
100 o
TYPICAL FOR FINAL DELTA WING COr\lFIG.
+ H+H
5 200 20
Figure 127. Cost Sensitivity
265
6 300 40
Range (1000 n. mi. ) No. of Passengers Fuel Cost (cents/lb)
l
) J
)
)
14. a SELECTION OF PHASE II CONFIGURATIONS
From the data shown in the previous sections, an over-all system comparison can be made and the Phase II configuration selected.
14.1 PASSENGER CAPACITY
Table 33 shows the significant data which influences the selection of passenger capacity. 100, 200 and 300-passenger vehicles are shown. The first line shows the take-off weight (the delta wing configuration was used as being representative) and the large growth in take-off weight as the passenger capacity is increased from 100 to 300 passengers. The second line shows the number of flights per day leaving New York (the busiest terminal) in the year 2000. This is taken from Figure 106 (broken lines)
and assumes an 80% load factor. The third row shows the total operating costs 'and the
fourth line shows the approximate values of sonic boom during climb at Mach 1.4.
From the values shown in Table 33, a 100-passenger capacity is too costly and will impose significant air traffic congestion problems. A 300-passenger capacity is the best from these two points of view, but has higher values of sonic boom overpressures due to the heavy take-off weight. These data suggest that a 200-passenger capacity is a reasonable choice
14.2 DESIGN RANGE
Table 34 shows design ranges of 4, 000, 5,000 and 6, 000 nautical miles and the factors that influence the choice between these values. The first line shows that take-off weight changes significantly between 4,000 and 6,000 nautical miles. The second line shows what percent of the total business plus government passenger traffic can be carried non-stop at these design ranges (from Figure 100). The third row shows a large increase in operating cost at the longer range. The fourth line shows'increases in sonic boom because of the effects of take-off weight.
From the values shown in Table 34, a design range of 4,000 nautical miles is too short to carry an appreciable percentage of the passenger traffic non-stop while 6,000 nautical miles results in high costs per seat mile. A design range of 5,000
nautical miles is a fairly good compromise and will accommodate routes from Los Angeles to Europe and Tokyo, as well as from New York to South America (Figure 99).
266
NUMBER OF PASSENGERS
100 200 300
TAKE OFF WT. (LB) 402,000 571,000 732,000
FLTS/DAY (NEW YORK) (Fig. 106 144 ~')
1- 48
y/SEAT MILE 9.9 h,4 '), 8
SONIC BOOM (PSF) 2, 7 3 .) . ~ 3.4
Tah Ie 33. Se Iection of Passenger Capacity
) RANGE
4,000 5,000 6,000
TAKE: OFF WT (LB, . 11-:11 fififi C::7i 000 765,000 I , I
J
I I
PASSENGERS NON STOP 43% 69% 86)0 I 1 ~/ SEA T l\ULE 5.6 6.4 I 9. O~,
I I , I
f I
ONT BOO I k'\ 2. 9 3. 2 3.45 I S Ie
Table 34. Selection of Design Range
)
267
)
)
14.3 CONFIGURATION SELECTION
Table 35 shows the four candidate vehicles as finalized in Sections 6.0, 7.0,8.0 and 9.0 and compared in Section 10.0. These are the delta wing, variable sweep wing, double delta blended body and scramjet configurations. Table 35 shows the necessary data required to select between these configurations.
The delta wing and variable sweep wing configuration show comparative data for these two competitive wing shapes. For the subsonic cruise and loiter requirements established in Figure 44, the delta wing is chosen over the variable sweep configuration since it has better all around characteristics. This choice could be changed if more subsonic range is required (Table 23 showed that the variable sweep is less sensitive to increases in subsonic range requirements because of its better subsonic lift-drag ratio). The variable sweep configuration is, therefore, retained until air traffic control requirements and mission failure effects have,been defined in the Phase II studies.
The double delta configuration is shown in the two right-hand columns of Table 35 for turboramjet and turboramjet/scramjet propulsion systems. The higher cruise Mach number of the scramjet gives shorter trip times and better utilization; however, all other features weight heavily against the scramjet. The costs are nearly 50% higher because the high take-off weight requires more fuel. The scramjet configuration is therefore dropped.
The chosen configurations are, therefore, the delta wing and the blended body", with the variable sweep retained until the subsonic cruise requirements are better defined. These configurations are sufficiently different to give depth and meaning to the studies to be conducted during Phase II. Table 36 shows the areas to be studIed in Phase II and the configurations that will be used.
268
I~
~ O"l to
I
~
.,
CRUISE MACH NO.
TRIP TIME (5000 N. M.)
UTILIZATION (FLTS/DAy)
TAKEOFF WEIGHT
SONIC BOOM (M = l. 4)
APPROX DOC (~/ SM)
SELECT
RETAIN
Table 5.
0 ~
"
VARIABLE
DELTA SWEEP BLENDEr
WING WING BODY SCRAMJET
~ ~ -cC} --CE}-!
6 6 6 8
2.23 2.23 2.23 1. 97 I
7 7 7 8
537,040 602,483 543,793 846,927
3.2 3.7 3, 1 3.4
6.4 7.2 6.5 10.1
• • •
Selection of Phase II Configuration
-
) DELTA VARIABLE BLENDED WING SWEEP WING BODY
~ ~ ~ ,
AERODYNAMICS • • SONIC BOOM • • AERO HEATING • INLET/ENGINE MATCHING • INLET/ENGINE COOLING • STRUCTURAL CONCEPTS • • INTEG. /NON INTEG. TANKS • A TMOSPHERIC TEMPERATURE •
) ENGINE FAILURE • • FUEL RESERVES • • TAKE OFF & LANDING • • FUEL SYSTEM • GROUND FUEL FACILITY • . POST FLIGHT • AIR TRAFFIC CONTROL • • FINAL CONFIGURATIONS • • •
) Table 36. Phase IT Study Areas and Selected Configuration
270
)
1.
2.
3.
REFERENCES
Request for Proposal from Ames Research Center, ?vloffett Fie 1d, No PROC RFP ..\-10597 (WEB-32), dated 29 April 1965.
Anon. "A Proposal for Study of the Performance Potential of Hydrogen- Fue led, Airbreathing Cruise Aircraft, General Dynamics Convair, 26 May 1965.
Jarlett, F. E., Outline of Final Report and Schedule for Contract NAS2-3180. Letter No. 581-5-1836, dated 9 September 1965.
4. Jarlett, F. E., First l\Ionihly Progress Report, GD/C-DCB-65-040, dated 17 October 1965.
5. Jarlett, F. E., Second Monthly Progress Report, GD/'C-DCB-65-042, dated 17 November 1965.
6. Jariett, F. E., Third l\Ionthly Progress Report, GD/C-DCB-65-046. dated 17 December 1965.
7. Jarlett, F. E., First Quarterly Presentation Charts 1 Report (:;D/ C-DCB-66-002,
dated 5 January 1966.
8. "Theoretical Prediction of Prediction of Pressures in Hypersonic Flow with Special Reference to Configurations Having Attached Leading- Edge Shock, " Mead, H. R. and Koch, F., ASD TR 61-60, Parts II and III, May 1962.
9. "Pressure Distribution Tests of Several Sharp Leading Edge Wings, Bodies and Body-Wing Combinations at Mach 5 and 8," Randall, R. E., Bell. D. R., and Burk, J. L., AEDC TN 60-175, September 1960.
10. "The'Aerodynamic Forces on Slender Plane and Cruciform-Wing and Body Combinations," Spreiter, J., NACA TR 962, 1950.
11. "Low Speed Pitching Derivatives of Low-Aspect-Ratio Wings of Triangular Plan Forms," Goodman, A., and Soquist, B. 1\1., N,.-'\CA RM L50C02, 1950.
12. "Market and System Analysis for Potential Global Transport Application of a
Reusable Orbital Transport - Final Report, " Lockheed, California, Report LR 19044, 29 July 1965.
271
,
\ \
.~~ ...
-:) REFERENCES (Continued)
~)
13. "Study to Develop System Criteria for Reusable Launch Vehicle Concepts -
Final Report. "
North American Aviation, Space and Information Systems Division, Report SID 65-848-1, 30 June 1965, Vol. I Report SID 65-848-3, 30 June 1965, Vol. ITL
14. Carlson, Harry W., "An Investigation of Some Aspects of the Sonic Boom by Means of Wind Tunnel Measurements of Pressures about Several Bodies at a Mach Number of 2.01, " NASA TN D-161, 1959.
15. Carlson, Harry W., IIWind-Tunnel Measuremerts of the Sonic Boom Characteristics of a Supersonic Bomber Model and a Correlation with Flight- Test Ground Measurements, tl NASA TM X-700, 1962.
16. Witham, G. B., liThe Flow Pattern of a Supersonic Projectile." Communications of Pure Applied Math., Volume V, No.3, August 1952, pp. 301-348.
17. Hilton, D. A., Hickel, Vera, Steiner, R., Maglieri, D. J., "Sonic-Boom Exposures During FAA Community-Response Studies Over a 6-Month Period in the Oklahoma City Area, NASA TN D-2539, December 1964.
18. McLean, F. E., "Some Nonasymptotic Effects on the Sonic Boom of Large Airplanes, NASA TN D-2877, June 1965.
19. McLean, F. E. and Carlson, H. W., "The Influence of Airplane Configuration on the Shape and Magnitude of Sonic-Boom Pressure Signatures, " AIA...'\ Paper No. 65-803, November 1965.
20. Roland, H. L. and Neben, R. E., Aircraft Structural Weight Estimating Methods,
Report ERR-FW-241, 15 September 1964.
21. Applied Research and Advanced Technology for the Supersonic Combustion Ramjet for 1964, AFAPL-TR-65-15, The Marquardt Corporation, April 1965.
22. Comparison of Airbreathing PropulSion Systems for Scramjet Cruise Vehicles, PWA-2559, Pratt & Whitney Aircraft, April 1965.
23. Final Review of Analytical and Experimental Evaluation of the Supersonic Combustion Ramjet Engine, General Electric, 4 October 1965.
272
---) REFERENCES (Continued)
)
24. Bernick, C. R., "Economics of Advanced Launch Vehicles," Report GD/C-ERRA1\-793, dated 31 December 1965.
25. Fuller, Dennis E. and Campbell, James F., !lSupersonic Lateral Directional Stability Characteristics of a 45° Swept Wing-Body-Tail Model with Various Body Cross-Sectional Shapes," NASA TN D-2376, August 1964.
26. Goodmanson, J. T. and Swihart, J. M., "Variable Sweep and Its Application to The Supersonic Transport," SAE 858C, May 1964.
27. Spencer, B., Jr., !fA Simplified Method for Estimating Subsonic Lift-Curve Slope at Low Angles of Attack for Irregular Planform Wings," NASA TM X-525, May 1961.
28. Breuhaus, W. 0., Reynolds, P. A., and Kidd, E. A •• "Handling Qualities Requirements for Hypervelocity Airc raft." Cornell Aeronautical Lab., Inc. Report No. TC-1332-F-1, January 1960.
273