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NASA CR 182125 _
NASA
LARGE SCALE ADVANCED PROP-FAN
(LAP)HIGH SPEED WIND TUNNEL TEST REPORT
By: William A. Campbell
Peter J. Arseneaux
Harry S. Wainauski
(NAqA-CR-I_212_) LARG_-SCAL_ AnVANCED
PROP-FAN (LAP) NIGH SPEFO WIND TUNNEL TFST
REPORT (k_amilton Standard1) 199 p CSCL 2IF
NQO-iOO_
UnclJs
G3/O? 0237147
" ;-_% --HA MiLTON_DARD DWiSlON --
:= - _ - - _--:', :_:_= Prepared for
......... National Aeronautics: a==n_d_Splce Adminis-trat_fon --_--" _NA _SA_gWis Resea__rch centei .......
- __ .... -=_ __--__Contract NAS3--23051
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NASA CR 182125
whole or in part. Date for general release July, 1989.
NASA
LARGE SCALE ADVANCED PROP-FAN
(LAP)
HIGH SPEED WIND TUNNEL TEST REPORT
By: William A. CampbellPeter J. Arseneaux
Harry S, Wainauski
HAMILTON STANDARD DIVISIONUNITED TECHNOLOGIES CORPORATION
Prepared for
National Aeronautics and Space AdministrationNASA-Lewis Research Center
Contract NAS3-23051
AiJ
SECTION
l.O
2.0
3.0
4.0
5.0
6.0
7.0
CONTENTS
ABSTRACT
SUMMARY
INTRODUCTION
PROP-FAN DESCRIPTION
4.l General Description4.2 SR-7L Blade
4.3 Pitch Change Actuator4.4 Pltch Control
4.5 Hub and Blade Retention
4.6 Spinner
INSTRUMENTATION DESCRIPTION
5.1 General Description
5.1.1
5.1.2
Electronlc Data Acqulsltlon System
Prop-Fan Diagnostic MonitoringInstrumentation
5.25.3
Steady Pressure Measurement System
Unsteady Pressure Measurement System
TEST FACILITY DESCRIPTION
6.1 Wind Tunnel Description
6.2 Test Rig Description6.3 Test Rig Modifications
SPINNER AND CENTERBODY DRAG MEASUREMENT
7.1 Test Rig Vlbration Survey7.2 Wind Tunnel Corrections
7.3 Test Rig Corrections
7.3.1 Test Objectives?.3.2 Test Procedure7.3.3 D1scusslon and Results
(DAS)
PAGE
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3
5
11
11
11
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1818
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63
PRECEDING PAGE BLANK NOT VILMED
iii
SECTION
8.0
9.0
10.0
11.0
12.0
APPENDIX
CONTENTS(Continued)
BLADE STRUCTURAL DYNAMIC EVALUATION
8.1 Test Objectives8.2 Test Procedure8.3 D1scusslon and Results
8.3.1 General Discusslon
8.3.2 Data Reductlon
AERODYNAMIC PERFORMANCE EVALUATION
9.1 Test Objective9.2 Test Procedure
9.3 D1scusslon and Results
9.3.1 Data Reductlon Procedure
9.3.2 Data Presentation
BLADE SURFACE STEADY PRESSURE MEASUREMENT
10.1 Test Objective10.2 Test ProcedureI0.3 D1scusslon and Results
BLADE SURFACE UNSTEADY PRESSURE MEASUREMENT
11.1 Test ObJectlves11.2 Test Procedure
II.3 Discusslon and Results
CONCLUSIONS
12.1 Blade Structural Dynamic Evaluatlon
12.2 Aerodynamic Performance Evaluatlon
12.3 Blade Surface Steady Pressure Measurement
12.4 Blade Surface Unsteady Pressure Measurement
List of Symbols
References
A Chronologlcal History of Test
PAGE
71
71
7177
7784
I05
I05I05106
106
108
125
125125
136
157
157
157
169
175
175176176177
179
183
m_
iv
FIGURE
3-I3-2
3-3
4-I
4-2
4-3
4-4
4-54-6
4-7
4-8
5-I
5-2
5-35-4
6-I
6-2
6-36-4
6-5
6-66-76-86-96-106-116-127-1
7-2
7-3
7-4
7-5
7-6
7-77-87-9
ILLUST TIONS
Prop-Fan Concept
LAP Program Elements
Prop-Fan Test Assessment (PTA) Program Elements
Large-Scale Advanced Prop-FanDesign Requirements and Goals SummaryFeatures of the SR-7L Blade Construction
LAP Pitch Change ActuatorMotor Driven Pitch Control MechanismLAP Control Schematic
SR-7L Hub
SR-7L SpinnerH.S.D. and O.N.E.R.A. Instrumentation List
LAP Instrumentation System Schematic
LAP Blade Pitch Angle Measurement System
Steady Pressure Measurement System
Modane-Avrleux Aerothermodynamlc Test CenterSI-MA Large Atmospheric Wind TunnelSI-MA Wind Tunnel SchematicSI-MA Mind Tunnel Nozzle Inlet to Test Section
Tunnel Driving Power and Reynolds Number as aFunction of Math number
Prop-Fan Installation In Test Sectlon/Chariot No. 3
Prop-Fan Drive System Engines and Gearbox
Prop-Fan Test Rig Power Llmitations
Prop-Fan Drive SystemProp-Fan Test Rig Dr|ve Train "Balance" Schematic
Origlnal Test Rig Configuration
Modlfled Test Rig ConfigurationAxial Mach number Distribution for Freestream
Mach number .354Axial Mach number Distribution for Freestream
Mach number .615
Axial Mach number Dlstributlon for Freestream
Math number .721
Axial Mach number Distribution for FreestreamMach number .754
Axial Mach number Dlstrlbutlon for Freestream
Mach number .788
Ax|al Mach number Dlstrlbutlon for FreestreamMach number .829
Splnner/Body Tare Test
Test Rlg Support Structure for Splnner/Body Tare Test
Prop-Fan Blade Stubs
PAGE
68
1012
1314161719
2021242526
2932333435
3637
383941424344
48
49
50
51
52
5356
5758
V ¸
ILLUSTRATIONS(Continued)
FIGURE
7-10
7-11
7-127-137-147-158-1
8-2
8-38-48-58-68-78-88-9
8-10
8-11
8-12
8-13
8-14
8-15
8-16
8-178-188-199-19-29-39-49-59-69-7
Centerbody Surface Static Pressure Tap Locations(Forward Section)
Centerbody Surface Static Pressure Tap Locations(Mid and Aft Sections)
Spinner Bulkhead Static Pressure Tap Locations
Prop-Fan Test Rig Forces
Spinner Drag Coefficient vs. Corrected Mach number
Centerbody Drag Coefficient vs. Corrected Mach number
Test Rig Support Structure for Structural DynamicEvaluatlon (2 and 4 Blade Configurations)
Test Rig Support Structure for Structural Dynamic
Evaluation (8 Blade Configuration)
Two Blade Test ConfigurationFour Blade Test Configuration
Eight Blade Test Configuration
Strain Gage Arrangement for Two Blade Configuration
Strain Gage Arrangement for Four Blade Configuration
Strain Gage Arrangement for Eight Blade ConfigurationTest Points and Blade Shank Moments (2 Blade
Configuration)Test Points and Blade Shank Moments (4 Blade
Configuration)Test Points and Blade Shank Moments (8 Blade
Configuration)
Comparison of Signal to Noise for Flatwise Shank
Moment Gages (0° Inflow Angle)Trends of IP Response for 3° Angular Inflow
(Shank Moment)
Trends of IP Response for 3° Angular Inflow
(Radial Bending)
Radial Bending vs. Span for 3° Angular InflowComparison of SR-7L Natural Frequency Test
Results to Predictions
Test vs. Analysis (Radial Bending vs. Span)
Test vs. Analysis (Effect of Power)Test vs. Analysls (Effect of Mach No.)
LAP Blade Angle Potentlometer Output
Cp and CTNET VS, 3, 4 Blades, MN = .2, _ = 0=
Cp and CTNET VS, J, 4 Blades, MN = .5, _ = 0 °Cp and CTNeT VS. _, 4 Blades, MN - .8, _ - 0°C, and CTNET VS. J, 2 Blades, MN - .5, _ = 3°
CTNET VS. Cp, Constant O, 4 Blades, MN = .2
CTNE, VS. C,, Constant O, 4 Blades, MN - .5
PAGE
59
60
61646568
72
73
747576787980
81
82
83
85
94
95
96
98
I00
I01102
I07Ii0
Iii
112
113
114
115
L
vl
FIGURE
9-8
9-99-I0
9-ll
9-12
9-13
9-14
lO-I
I0-2
I0-3]0-4
10-5
10-6
10-7
]0-8
I0-9
lO-lO
lO-ll
lO-12
lO-13
lO-14
lO-15
lO-16
lO-17
lO-IB
]0-19
I0-20
ILLUSTRATIONS(Continued)
CTNET VS. Cp, Constant J, 4 Blades, MN = .8
CTNET VS. Cp, Constant J, 2 Blades, MN - .5, 9 = 3°
CTNET VS. Cp, Constant J, 8 Blades, _ = 0°CTNET VS. J, Constant Cp, 4 Blades, 9 = 0°
CTNET VS. J, Constant Cp, 8 Blades, _ = 0°
Comparison of Measured and Predicted Performance,4 Blades
Compar|son of Measured and Predicted Performance,8 Blades
Steady Pressure Test Rig Support StructureProp-Fan Steady Pressure Test Set-Up
LAP Blade Installation, Steady Pressure Test (2 Blade)
LAP Steady Pressure Blade (Camber Side)
LAP Steady Pressure Blade (Face Side)Steady Pressure Blade (Camber Side)
Steady Pressure Blade (Face Side)
Arrangement of the Pressure Taps on the Blade
Surface-Projected View
LAP SR-TL Steady Pressure Distribution of On-LlneData (Operating Condition No. 7)
LAP SR-TL Steady Pressure Distribution of On-Llne
Data (Operating Condition No. 2)
LAP SR-7L Steady Pressure Distribution of On-LlneData (Operating Condition No. 3)
LAP SR-7L Steady Pressure Distribution of On-Llne
Data (Operating Condition No. 4)LAP SR-TL Steady Pressure Distribution of On-Line
Data (Operating Condition No. 5)
LAP SR-7L Steady Pressure Distribution of On-LineData (Operating Condition No. 6)
LAP SR-7L Steady Pressure Distribution of On-Line
Data (Operating Condition No. 9)
LAP SR-?L Steady Pressure Distribution of On-Line
Data (Operatlng Condition No. 8)LAP SR-7L Steady Pressure Distribution of On-Line
Data (Operating Condition No. 7)
LAP SR-7L Steady Pressure Distrlbutlon of On-LineData (Operating Condition No. lO)
LAP SR-TL Steady Pressure Distribution of On-Line
Data (Operating Condltlon No. 11)
LAP SR-7L Steady Pressure Distribution of On-Line
Data (Operating Condition No. 13)
PAGE
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117
118
120121
122
123
128127
128130131
132133
134
140
141
142
143
144
145
146
147
148
149
150
151
vii
FIGURE
10-21
I0-2210-23
10-24
II-I
!I-2
11-311-4
II-5
11-6
I!-7
11-811-911-1011-11
ILLUSTRATIONS(Continued)
LAP SR-7L Steady Pressure Distribution of On-LineData (Operating Condition No. 12)
Repeatability Problem for Static Low Power Point
Repeatability Problem for Static High Power Point
Illustration of Repeatability at My = .20
Unsteady Pressure Test Rig Support StructureProp-Fan Unsteady Pressure Test Set-Up
LAP Blade Installation Unsteady Pressure Test (2 Blade)
LAP Unsteady Pressure Blade (Camber Side)
LAP Unsteady Pressure Blade (Face Side)Unsteady Pressure Blade, Transducer Locations
(Face Side)
Unsteady Pressure Blade, Transducer Locations(Camber Side)
LAPISR-TL Unsteady Pressure Blade Transducer Mounting
Unsteady Pressure Test Set-Up With Wake Generator
Examples of Measured Unsteady Pressures
Illustration of the Terms "Advancing" and "Retreating"
for Angular Inflow
PAGE
152
153
154
155
158
159
161
163164
165
166
167
168171
173
TABLES
TABLE
7-I
8-I
8-II
8-III
8-IV
8-V
8-Vl
8-VllI0-I
II-I
ll-ll
Splnner/Centerbody Drag Test Points vs. Corrected Mach No.
Operating Conditions for Mn >.l and SampleIRP Tabulations (2 Blade Configuration)
Operating Conditions for Mn >.l and SampleIRP Tabulations (4 Blade Configuration)
Operating Conditions for Mn >.l and SampleIRP Tabulations (8 Blade Configuration)
Measured Vibratory Peaks vs. LimitsP-Order Content of Strain Gage Output at 3° Inflow Angle
(2 Blade Configuration)
Comparison of IP Content of Strain Gage Output
Obtained by Two Methods
Comparison of Test vs. Analysis (IP)
Operating Conditions for Blade Steady Pressure Testing(2 Blade LAP Propeller)
Operating Conditions for Blade Unsteady Pressure Testing
(2 Blade LAP Propeller)Chart of Transducers from wh|ch Data was Acquired vs.
Test Condition
PAGE
62
86
87
8990
92
93
103
135
160
170
ix/x
l.O ABSTRACT
High Speed Wind Tunnel Testing of the SR-7L Large Scale Advanced Prop-Fan(LAP) is hereln reported. The LAP |s a 2.74 meter (9.0 FT) diameter,
8-bladed tractor type rated for 4475 KW (6,000 SHP) at 1698 RPM. It was
designed and built by Hamllton Standard under contract to the NASA Lewis
Research Center. The LAP employs thin swept blades to provide efficlent
propulsion at flight speeds up to Mach .85.
Testing was conducted in the ONERA SI-MA Atmospheric Wind Tunnel in Modane,France. The test objectives were to confirm the LAP was free from high speedclassical flutter, determlne the structural and aerodynamic response toangular Inf|ow, measure blade surface pressures (static and dynamic) andevaluate the aerodynamic performance at various blade angles, rotationalspeeds and flight Mach numbers.
The measured structural and aerodynamic performance of the LAP correlated
well with analytical predictions thereby providing confidence in the computer
codes used for design. There were no signs of classlcal flutter throughout
all phases of the test up to and including the 0.84 maximum Mach number
achieved. Steady and unsteady blade surface pressures were successfully
measured for a wide range of Mach numbers, Inflow angles, rotatlonal speeds
and blade angles.
No barriers were discovered that would prevent proceedlng with the PTA
(Prop-Fan Test Assessment) Flight Test Program scheduled for early 1987.
1/'2
Z
2.0 SUMMARY
This report describes the procedures followed and results obtained during
High Speed Wind Tunnel Testing of the SR-7L Large Scale Advanced Prop-Fan(LAP). The LAP Is a 2.74 meter (9 foot) diameter, 8 bladed advanced
propeller designed to attain high propulsive efficiency at flight speeds up
to Mach .85. The Prop-Fan achieves this superior speed and efficiency byemploying thin swept blades and high disc loading. The High Speed Wlnd
Tunnel Test was conducted in the ONERA SI-MA Large Atmospheric Wind Tunnel
facility In Modane, France.
Testing was accomplished during two separate tunnel entries. This wasnecessitated by a test rlg failure during the first entry. Prior to
Interruption of the test in early 1986,.a11 structural dynamic, aerodynamic
performance and a limlted amount of blade steady pressure testing was
completed. Testing resumed in early 1987, and culminated in the completionof the steady and unsteady blade pressure tests. A complete chronological
history of both tunnel entries Is provided in Appendix A.
The purpose of the Wind Tunnel Test was to confirm the LAP was free from highspeed classical flutter, determine the structural and aerodynamic response to
angular inflow, measure blade surface pressures (both steady and unsteady)
and evaluate the aerodynamic performance at various blade angles, rotational
speeds and Mach numbers. Results from these tests would assist in determiningthe readiness of the LAP and Its Instrumentation systems for follow-on
Prop-Fan Test Assessment (PTA) flight testing. The Wind Tunnel Test was
accomplished in two phases. In the first phase, structural dynamic and
aerodynamic performance data were collected concurrently, for the 2, 4 and 8
blade configurations, over a wide range of blade angles, Mach numbers and
rotational speeds. Due to rig constraints, structural dynamic evaluatlon ofthe LAP operating at a fixed angle of attack was limited to the 2 bladed
Prop-Fan configuration at a 3° inflow angle. During the second phase of
testing, a specially fabricated static pressure tapped blade was employed to
map the blade surface steady pressure distributlon for a range of operatingconditions. Due to drive system power limitations, testing was accomplished
utilizing a two bladed Prop-Fan configuration to provide blade loadingssimulating the take-off, cutback and design cruise conditions. The final
phase of testing, also utilizing the two blade configuration for reasons
mentioned above, employed another specially instrumented blade incorporating
high frequency response pressure transducers. Data from these transducers wasused to define and evaluate the blade surface unsteady pressure distribution
for the same operating conditions run during the steady pressure test.
Unsteady pressure testing included evaluating the effects of a wake In the
propeller inflow at a 3° angle of attack.
2.0 (Continued)
Results from the High Speed Wind Tunnel Test demonstrated that the SR-7L
Prop-Fan was free of high speed blade flutter over the entire operating
envelope tested. Additionally, all measured blade surface and blade shank
strains were well below allowables set prior to testing. Good correlationwas found between measured and analytlcally predicted IP strain sensitivities
for the SR-7L blade. Results confirmed that IP strain peaks Inboard on the
blade and lessens near the tip, and that, in general, blade strains were
found to increase with power and Mach number. Measured aerodynamic
performance for the four blade configuration corresponded well withanalytlcal predictions over the entire range of points tested. Though good
agreement for the eight blade configuration was found at Mach numbers of .70
and .73, performance was slightly underpredlcted at .5 Mach number. Steadyand unsteady blade surface pressure measurements were successfully collected
for a wlde range of Mach numbers, rotational speeds, blade angles and inflow
angles. In addition to confirming the presence of tip edge and leadlng edgevortex flows at Mach numbers of zero and .2, shock waves were evident at the
traiIlng edge at .7 and .78 Mach number. Unsteady pressure responses wereclearly present as evidenced by a dominant l-P response in the angular inflow
data and significant 2-P response in the wake inflow data. Sinusoldal
response was evident on the pressure (face) side of the blade for all casesexamined for angular inflow conditions and on the suction (camber) side under
low loading conditions. Under high loading condit|ons, the suction (camber)side exhibited non-slnusoldal response resulting from the presence of tip and
leading edge vortices.
3.0 INTRODUCTION
The Prop-Fan, a high speed, high efficiency alrcraft propulsion concept waslaunched during the "O|l Crunch" days of the mid 19?O's. In response to the
national need to reduce fuel consumption, Congress directed NASA to address a
series of aircraft related technologies aimed at increasing the fuel
efficiency of airline operation. In response, NASA created the Aircraft
Energy Efficiency (ACEE) program which addressed fuel savings through
advancements in both airframe and engine technology. The element of the ACEEprogram offering the greatest potential fuel savings was the Advanced
Turboprop Program (ATP) as described In Reference I. The NASA Lewls Research
Center had total responsibility for the ATP project which is summarized in
Reference 2. The objective of the ATP was to demonstrate technology readinessfor efficient, reliable and acceptable operation of advanced turboprop-powered
commercial transports at cruise speeds up to Mach 0.8 and at altitudes above
9,BOO meters (30,000 ft.) while maintaining cabin comfort levels (noise and
vibration) comparable to those of modern turbofan-powered aircraft. The
technology would also apply to possible new military aircraft for a variety of
missions. Out of this project evolved the Prop-Fan concept.
Although high propulsive efficiency from turboprops was nothing new, the
standards of high cruise speed and cabin comfort set by the contemporary
turbofan powered alrcraft were beyond the capability of any turboprop poweredaircraft. The Prop-Fan concept evolved to satisfy the requirements of high
speed and altitude with improved efficiency while maintaining a high degree of
cabin comfort. It is characterized by the large number of blades (8 or lO),
thin airfoil sections, and swept blade planforms (Figure 3-I).
Once the concept and its benefits were defined, NASA conducted a systematic
program to verify that the predicted benefits could be achieved and that
there were no unsolvable problems in Implementing the concept. The potentialbenefits of the Prop-Fan propulsion concept have been investigated in numerous
propulslon and aircraft systems studies conducted by the airframe and engine
manufacturers under NASA sponsorship (References 3 thru 9). These studies
have shown that the inherent efficiency advantage that turboprop propulsionsystems have achieved at lower cruise speeds may now be exte,ded to the
higher speeds of todays turbofan and turbojet powered alrc_aft. By applying
swept wlng/reduced dlameter technology to the design of prol=eller blades, and
by achieving higher disc 1oadlngs through the use of a greater number ofblades, It was found that the inherent fuel efflclencyof t_e propeller can
be extended to speeds up to .85 Mn. The efficiency of the l_rop-Fan should
a11ow aircraft to be designed that are 15% to 25% more eff$cIent than today's
most technologlcally advanced turbofan powered airlines (l_mference I0).
Since 1975, Hamllton Standard has been deeply Involved v_lfl_,the NASA Lewis
Research Center In the development of the Prop-Fan. I/_I_ re{ently, thiseffort utillzed a series of .622 meter (2 ft.) dlametermc_e_s which
Incorporated differing numbers of blades as well as chamg_ _n blade shape.These models underwent exhaust|ve testing in several wlm_ _mels at NASA and
OF POOR Q'J_kb! _ '_
-L
ORICINAL PAOE_.A,.,r_ A_'4D WHITE- PHOTOGRAPH
FIGURE 3-1. I_OP-FAN CONCEPT
6
3.0 (Contlnued)
United Technologies as wel] as on a NASA Jetstar acoustic research vehicle.
In these tests the targeted efficiencles were demonstrated, the source noisewas characterized, and structural dynamics verlf|ed. Detailed descriptionsof these tests and results have been the subject of numerous technical
papers; a summary of which can be found in References II and 12.
Although the results of the aerodynamic performance and source nolse tests can
be confidently scaled from model to product size, the structure of the solid
homogeneous model blades Is so different from that envisioned for a productthat extrapolation of model structural behavlor Is unacceptable. The
verification of the structural integrity of a large-scale Prop-Fan then
becomes the final major technical hurdle to be crossed before industry
acceptance of the Prop-Fan as a viable aircraft propulsion scheme. Thisverification was initiated in 1983, when Hamilton Standard, under the
sponsorship of NASA Lewis Research Center, embarked on a program to build andtest a 2.74 meter (9 foot) diameter, 8-bladed Prop-Fan. Thls Prop-Fan was
designated the SR-7L or the Large Scale Advanced Prop-Fan (LAP). The majorelements of the LAP program are depicted in the summary schedule shown in
Figure 3-2. Detail design and fabrication of the Prop-Fan components(blades, hub and blade retention, spinner, pitch change mechanism, pitchcontrol and instrumentation system) was initiated early In 1983 building on a
preliminary design conductedunder an earlier contract. Various bench tests of each component then
followed to verify key design characteristics.
Design, fabrication and test of an aeroelastlcally scaled .622 meter (2 foot)
Prop-Fan model was included in the LAP program to obtaln an early assessmentof the Prop-Fan's aeroelastic characteristics. This model has been
designated SR-7A. Other objectives of the SR-7A testing Included the
measurement of aerodynamlc performance and noise.
Testing of the SR-7L rotor under the LAP program included in-house whirl,static and high speed wind tunnel tests. Whirl Rig Testing was successfully
conducted on the G-5 rlg at Hamilton Standard. The objectives of the testwere to measure the stiffness of the blade retentlon system, evaluate the
control dynamic characteristics of the blade pitch change system and determinethe wear rate on blade actuation and retention hardware. Whirl Rig Testlng
was conducted using stub weights to simulate the Prop-Fan blades. The stubs
provided appropriate centrifugal loading but generated essentially no thrustor drag. The Static Rotor Testing, with SR-7L blades installed, was success-
fully conducted at the Wright Aeronautical Laboratory at Wrlght Patterson AirForce Base in Dayton, Ohio (Reference 13). The goals of the Static Rotor
Test were to measure the static aerodynamic performance of the LAP, assessstall flutter characteristics and investigate the structural behavior and
integrity of the SR-TL blades for static operating conditions. The final
component of the LAP test program, High Speed Wind Tunnel Testing, is the
subject of this report.
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3.0 (Continued)
The LAP program is comPlemented by another NASA sponsored program, theProp-Fan Test Assessment (PTA) program, which takes the large scale Prop-Fan
(developed under the LAP program) and mates it with an Allison Gas Turbinesupplied gas generator and gearbox to form a Prop-Fan propulsion system. The
ma_or elements of the PTA program are depicted In Figure 3-3. Following
completion of the Static Englne Test at Rohr's Brown Field Facl]ity, the
quick engine change (QEC) nacelle will be mated to the wing of Gulfstream II
aircraft which will ultimately serve as the f]|ght test vehicle for the
Prop-Fan.
This report addresses the procedures and results of the High Speed Hind Tunnel
Test conducted at the ONERA SI-MA Atmospheric _ind Tunnel in Modane, France.
This facility was selected for three reasons. First, it is capable of
reaching high cruise Mach numbers. Second, it is sufficiently large (8 meteror 26.25 ft. diameter test section) to avoid excessive wall interference
effects; and third, it has an existing model drive system. Although the power
capability of the drive Is only about one fourth of what the Prop-Fan is
designed to absorb, proper blade loading can be reached by running with apartial set of blades (eight, four, and two blade configurations). The
specific objectives of the test are listed below.
To conduct a careful and controlled search for any evidence of
classical flutter. Because of the greater air density of the :nd
tunnel, it is possible to more closely approach the flutterthreshold than at the 10,668 meter (35,000 ft) flight altitude. At
design Mach number the wind tunnel operates at an effective altitude
of about 4,267 meters (14,000 ft). Analytic predictions and tests
of the SR-7A model strongly suggested that classical flutter wouldnot be encountered.
To measure steady and unsteady surface pressure on the blade as wellas overall Prop-Fan performance. One blade was Instrumented with
465 static taps (20 chordal at 13 radial stations on the camber sideand |6 chordal at 13 radial stations on the face side of the blade)
to obtain a complete pressure map. Another blade had 26 dynamicpressure sensors (7 chordal at the 35 inch station and 6 chordal at
the 49 inch station, for each side of the blade) to assess unsteady
effects. These measurements will provide bench_rk data forunderstanding the physlcs of transonic flow over the blades and for
verlficatlon of analytic computer codes.
To determine the structural and aerodynamic response of the Prop-Fan
to angular Inflow. Analysis of data from this simple, known angularinflow condition will slgnlfIcantly contribute to the understanding
of Prop-Fan behavior in the more complex, airplane installed flowfield.
LARGE - SCALE PROPFAN
FROM LAP
UNIQUE TILTNACELLE
MOOIirlED TURBOSHAFT IIb / \ II
ENGINE GEARBOX /_
" _SYSTEM STATIC TEST
AT ROHR BROWN FIELD
AIRCRAFT.OOELTESTS\ / /
k; (.:._.'_ " /'_ _._ CAB,,,ENVIROMENT
1 T- 1.2f,_;_lN tillS" _ "-;-'_',_ EN-ROUTE NOISEL I _'I._JL_ - FLIGHT TESTING _"_-a'mlJ_a_-,'_Imb _ INSTALLED OPERATING
CHARACTERISTICS
FIGURE 3-3. PROP - FAN TEST ASSESSMENT (PTA} PROGRAM ELEMENTS
10
4.0 PROP-FAN DESCRIPTION
4.1 General Description
The Large Scale Advanced Prop-Fan shown _n Figure 4-1, is a 2.74 meter(9 ft.) diameter 8 bladed tractor type Prop-Fan rated for 4474 KN (6,000 HP)
at 1698 RPM. To achieve the program objective of verifying large scale
Prop-Fan structural integrity, a number of design requirements and goals were
established as summarized in Figure 4-2. The requirements Includecharacteristics Judged essentlal to meeting the program objective as well as
design features established from prior work. The goals, on the other hand,
represent design targets and were judged less Important to the program
objective. The LAP is designed to be mounted on a standard 60A spIined
propeller shaft for an existing turboprop gearbox. It has a hydrau11callyactuated blade pitch change system and a hydromechanlcal pitch control that
allows the Prop-Fan to operate In a speed governing mode or, with minormodifications, in a Beta Control mode, as was the case in this test. Beta
Control mode operation was chosen to provide the operator the capability toselect desired blade pitch angles while running. A pitchlock feature Is also
incorporated in the actuator. This feature malntalns the propeller blade
angle In the event of a loss of system operating oil pressure. The design of
the actuator and control is based on proven technology used in Hamilton
Standard's military and commercial propellers. A brief description of each of
the major elements of the LAP as depicted in Figure 4-I, is presented below.
4.2 SR-7L Blade
Features of the structural configuration of the SR-7L blade are shown inFigure 4-3. These Include a central aluminum spar which forms the structural
"backbone" of the blade, a multi-layered glass-cloth-reinforced shell
overhanging the leading and trailing edge of the spar, a nickel sheath which
covers the leading edge of the outer two-thirds of the blade, and anon-operational integral heater In the Inboard leadlng edge area. Though the
scope of the LAP testing never Included utillzation of the blade heaters, Itwas decided to Install the heaters to evaluate the structural response of a
blade closely resembling that of a typical blade configuration. The remaining
internal cavitles are filled wlth Iow-denslty rigid foam. The outboard
portion of the spar is Intentionally moved toward the blade leading edge to
increase stabillty by reducing overhung mass in the tip trailing edge, whileat the same time increasing the Integrity of the leadlng edge from the stand-
point of resistance to foreign object damage.
The blade design makes use of a NACA Series 16 airfoil outboard and a Serles
65 clrcular arc airfoil Inboard. Each blade has an activity factor of 227.3
wlth 45" of blade sweep at the tip. The blades were designed with
predeflectlon so that the blades will assume the desired aerodynamic shape at
the cruise operating condition (Reference 14).
11
BLADE
=:
BLADE RETENTION
CONTROL
ACTUATOR
SPI
FIGURE 4-I. LARGE-SCALE ADVANCED PROP-FAN
12O_!GINAV p_v
_Lf-,CK At, iD WHi-rE PHOIOGRAPH
_ UNITEDTECHNOLOGIES
DESIGN REQUIREMENTS
Configuration
DiameterNumber of blades
2.74 meter (9 ft)6
• Design point
Cruise Mach numberAltitudeTip speedPower loading (SHPID2)
0.810,668 meters (35,000 ft)244 meterslsec (800 ft/sec)32
• Structural Integrity
Flutter free over normal flight envelope (M <0.8)Stresses within allowable limitsOverspeed toleranceCritical speed margins
Safety featuresLeading edge projectionLighting protectionIcing protection (installed but not operational)Overspeod protection
• Reverse thrust capability
DESIGN GOALS
• Net efficiency (isolated nacelle)
• NoiseNear field (designpoint cruise, max.free field, 0.81))Far field
-78.6% AT M = 0.8, 10,668 meters (35,000 ftXCruise)-52.0% AT M = 0.2, SL (TO)
.144 db overall sound pressure level
•FAR 36 minus 10 db
• Stall flutter -None st 100% TO power and rpm; M = 0- 0. 2
• High speed (classical) flutter -None over extended flight envelope,(M __<0.85) 105% max operating speed
• Over speed limit (hub, blades,blade retention)
.120% max operating speed - no yield-141% max operating speed - no failure
Foreign object damageMinor-Birds up to 4 ozModerate -2" Hail; Birds to 2 Ib
Major-Birds up to 4 Ib
-No damage to pdmary blade structure-Some loss of matedal or airfoil distortion;
operate at 76% power for 5 minutes-Some loss of matedal or airfoil distortion;
maintain ability to feather
• Blade life .35,000 hr - replacement with scheduled malnt.-50,000 hr - meantime between unscheduled
removal
FIGURE 4-2. DESIGN REQUIREMENTS AND GOALS SUMMARY
13
z
FIGURE 4-3. FEATURES OF THE SR-7L BLADE CONSTRUCTION
14
4.2 (Contlnued)
Although some improvements in sweep/stress/stability trade-offs were
predicted through the use of advanced composites, it was decided not toinclude these In the final blade design. Their use would requlre the
development of new manufacturing technology, both in terms of suitable
construction methods and processes, and lengthy development of design
allowables to reflect the manufacturing process.
It was felt that the scope of the program would be best served by utilizlng
the service-proven combination of an aluminum spar enveloped with a
fiberglass shell for which processes and stress allowables are well known.
4.3 Pitch Change Actuator
The pitch change system Is comprised of two components, a pitch change
actuator and a control. The pitch change actuator Is the prime mover for
blade angle change and Is located within the Prop-Fan hub as shown inFigure 4-4. Its primary components are an internal stationary piston, a
translating outer cylinder with an integral yoke to engage each of the blade
trunnlons and a pltchlock and servo assembly which contains a four-way
metering beta valve assembly, a pitchlock screw, a ground adjustable lowpitch stop, a servo piston and a ball screw to drive the pitchlock screw andbeta valve (Reference 15).
As stated earlier, the LAP was modified to operate In the Beta Control mode
throughout all phases of the wind tunnel test. This was accomplished throughthe use of an electromechanlcal controller (D.C. motor and gearhead) mounted
on the front of the dome as shown in Figure 4-5. The controller provides the
rotary input to the pitchlock servo directly, replacing the rotary input from
the half area servo. In order to give the D.C. motor full control of the
blade angle, the servo and ballscrew must be disconnected from the pitchlockscrew by removing the quill shaft. This disables the control signal and
allows the motor to work directly on the pitchlock screw without having to
fight the servo output. This permits the operator to remotely position blade
angle as deslred while running.
The actuator was designed to present state-of-the-art technology and low
development risk technique that has been used on a number of existing
propeller systems. The design uses mostly steel for the load carryingmembers and all surfaces subject to sliding seal wear are chrome plated to
increase durability. The actuator was designed to conservative stress and
deflection levels to minimize development effort while maintaining a
reasonable but not minimum weight.
The pltch change mechanism was designed such that all malfunctions will
either cause the system to pltchIock or go to Feather. An additional safety
Feature on the LAP is a ground adjustable low pitch stop. This 11mlts the
minimum blade angle under all circumstances.
15
BLACK
FIGURE 4-4 LAP PITCH CHANGE ACTUATOR
ORIGINAL PAGE
AND WHITE pI_OTOGRAPIq
16
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17
4.4 Pitch Control
To minimize cost and development time, It was decided to utilize a modified
version of an existing turboprop propeller control. Based on the type of
engine and gearbox planned for use with the LAP, the control selected was a
modified 54460 control (Figure 4-6). The 54460 hydromechanlcal integral oiicontrol, which Is currently In use on the Grumman E-2 and C-2 aircraft, was
modeled after the Lockheed C-130 and P-3 controls. Since the first production
unit in 1956, there have been over 12,000 built and they have logged over
80,000,000 operating hours.
The primary func_ion of the LAP control, as modified for wind tunnel testing,was to generate the hydraulic pressure necessary to assure proper actuator
operation. Two gear type pumps located in the stationary control and driven
by the Prop-Fan shaft provided the system hydraulic pressure.
4.5 Hub and Blade Retention
The LAP hub assembly forms a seml-rlgid llnk between the blades, which
provide the thrust, and the engine shaft, which provides the torque(Reference 16). The hub and tailshaft is a one piece forged component which
is carburlzed, heat treated and machined (Figure 4-7). A single row ball
bearing retains each of the eight blades In the hub, while the tailshaft
secures the Prop-Fan to the engine shaft through two cone seats that are
preloaded against each other by the Prop-Fan retaining nut. The hub alsoforms the support for the pitch change actuator system, the control and the
spinner.
The retention transmits the loads from the blades to the hub while
accommodating changes in blade pitch. The single row ball bearing retentionprovides ease of maintenance by allowing individual blade replacement without
disassembly of the hub. It has a through hardened Inner race which seatsagainst the aluminum blade shank and an outer race which Is integral with the
barrel. The outer race Is carburlzed to achieve the hardness necessary to
support the ball loads. The balls are kept apart from each other by an
elastomerIc separator. The rotational speed of the Prop-Fan keeps the
retention submerged In oll which is contained in the hub by eight blade seals(Reference 17).
4.6 Spinner
The LAP spinner and rear bulkhead assembly Is essentially a reinforcedflberglass/epoxy shell, supported by the hub and actuator, and Incorporating
an aerodynamic shape to facilitate proper alr flow around the blade roots
(Figure 4-8). Its primary function is to insure proper Prop-Fan aerodynamicperformance. The rear bulkhead, which mounts on the rear of the hub arms, Is
the main structural support for the spinner and provides a mounting surfacefor much of the instrumentation hardware in the rotating field.
18
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FIGURE 4-7. SR-7L HUB
20
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BLACK A _'_n,-,,., '.,f.,'r4t"t-E -.I ,,- -._.--,D,,Du,_,...,,,j,...,,;..,
FIGURE 4-8 SR-7L SPINNER
21/22
r_
5.0 INSTRUMENTATION DESCRIPTION
5.1 General Description
In order to accomplish the objectives set forth in the Plan of Test, three
separate Instrumentatlon arrangements were employed to gather and record the
desired data. For the structural dynamic and aerodynamic performance tests,
the system described in sections 5.1.I and 5.1.2 was used. For the blade
surface steady pressure test, a system description is provided In section 5.2
and for the unsteady pressure test, the system used is presented insection 5.3. Figure 5-I provides a summary listing of all the
instrumentation, rotating and stationary, and indicates whether It was
provided by H.S. or ONERA. Additional detailed descriptions of the key data
gathering devlced required for specific phases of the test are presented in
the appropriate Test Procedure section of this report.
Common to all three instrumentation arrangements was the frequency modulated
multlplex electronic data acquisition system (DAS), as shown in Figure 5-2,
and several Prop-Fan system diagnostic monitoring devices.
5.1.1 Electronic Data Acquisition Systems (DAS)
The electronlc data acquisition system for the LAP provided the capacity totransmit 33 channels of information from the electronic measurement devices
on the rotating slde of the Prop-Fan to the data collection and monitoring
equipment in the stationary field. This was accomplished by employing an
eight rlng platter-type sllp ring assembly which provided the electrlcal
Interface for the DAS between the rotating Prop-Fan assembly and the
non-rotatlng control.
As illustrated in Figure 5-2, three rings were utilized to transmit data to
the stationary field. One of the three carried the output of a potentiometerwhich was mounted as shown in Figure 5-3 and used for monitoring blade angle
position. The other two rings transm|tted the remaining 32 channels of
information which consisted of signals from two miniature pressure transducers
for monitoring actuator high and low pitch pressures and combinations of upto 30 blade strain or unsteady pressure signals, as determined by the
particular test being conducted. Transmittal of these 32 signals on only two
sllp rings necessitated the use of FM multiplexing. The DC signals from the
strain gages and pressure transducers in the rotating field were divided intotwo groups of sixteen. The signals were then converted to frequency modulated
signals by two groups of voltage controlled oscillators. Each group was then
multlplexed by a mixer, allowing thirty two channels to be transmittedthrough two sllp rings. The two groups of sixteen channels were each
detranslated in the stationary field to four groups of four multiplexedchannels (IRIG Standard IA thru 4A). Each set of four channels was recorded
on one track of a standard Honeywell IOl tape recorder. Simultaneously,
eight discriminators were used to demodulate any two groups of four channels
for real tlme monitoring of data. One discriminator was tuned to the centerfrequency of each channel.
_G PAGE
23
H.S. INSTRUMENTATION
ROTATING:
HIGH AND LOW PITCH OIL PRESSURE MEASUREMENTBLADE ANGLE MEASUREMENT
BLADE VIBRATORY STRAIN MEASUREMENTS
BLADE SURFACE STEADY AND UNSTEADY PRESSURE MEASUREMENTS
STATIONARY:
CONTROL SUPPLY PRESSURE MEASUREMENT
CONTROL SUMP OIL TEMPERATURE MEASUREMENT
HEAT EXCHANGER _P MEASUREMENT
CONTROL AND SCANIVALVE ACCELEROMETER MEASUREMENTS
INSTRUMENTATION PROVIDED BY ONERA (ALL STATIONARY)
TEST RIG INSTRUMENTATION:
RIG ACCELEROMETER MEASUREMENTSPROP-FAN RPM PICK-UP
RIG BEARING TEMPERATURE MEASUREMENTS
BALANCE AND TORQUEMETER MEASUREMENTS (FIGURE 6-I0)
CENTERBODY PRESSURE MEASUREMENTS (FIGURES 7-I0, 7-11)SPINNER BULKHEAD PRESSURE MEASUREMENTS (FIGURE 7-]2)
TUNNEL INSTRUMENTATION:
STATIC PRESSURE MEASUREMENTS:
- 4 METERS UPSTREAM OF THE PROP-FAN PLANE OF ROTATION
- ADJACENT TO THE PROP-FAN PLANE OF ROTATIONSTAGNATION PRESSURE MEASUREMENTS
AIR DENSITY MEASUREMENT
STAGNATION TEMPERATURE MEASUREMENT OF FLOW
STATIC TEMPERATURE MEASUREMENT OF FLOW
FIGURE 5-]. H.S. AND ONERA INSTRUMENTATION LIST
24
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25
FIGURE5-3. LAPBLADEPITCHANGLEMEASUREMENTSYSTEM
26
5.1.1 (Continued)
The FM electronic instrumentation system provided inherent noise immunity for
data transmission. The frequency response for the system was DC to lO00 HZ.Overall accuracy of the system was ±3% RSS. Time correlation betweenchannels was +13.8 mlcroseconds.
The electronic Instrumentation system allowed for up to ten blade shell or
shank gages to be installed on any blade, though a maximum of 32 gages could
be active at any one time. Straln gaged blades were closely monitored during
flutter and critical speed testing as well as measuring strains at various
operating conditions. The blade shank strain gages were employed to measureblade bending moments.
A total of sixteen gages could be selected from blades one through four and
an additional sixteen gages could be chosen from blades flve through eight.
Selection of the desired combination of gages was accomplished using eightprogrammable connectors mounted on the Prop-Fan hub. Programming of the
connectors required using Jumper wires to connect the sockets of patch boards
in the connectors. The bridge completion circuits for the strain gages werelocated on circuit boards in the blade cuff.
Monitoring of instrumentation during the test was accomplished with an
oscilloscope, a spectrum analyzer and a vlsicorder. The oscilloscope
permitted a time domain display of four channels simultaneously. The
spectrum analyzer provided the capability to dlsplay any one channel in thetime or frequency domain. The analyzer also had transient capture and
playback features. The vlslcorder provided a hard copy plot ofinstrumentation signals versus time.
5.1.2 Prop-Fan Diagnostic Monitoring Instrumentation
In addition to monltorlng and collecting data from the rotating fleld, a
number of parameters intended to provlde protection for the Prop-Fan system
were measured and digitally dlspIayed by the stationary field instrumentation.These included control sump oil temperature, control supply oil pressure,
differential pressure across the heat exchanger and control/Scanlvalve
accelerometer measurements. The control sump temperature was measured by athermocouple installed inside the oll draln port of the control. Control
supply oil pressure was measured by a transducer Inslde the control. This
transducer slgnal was transmitted via an existing connector on the control.
The heat exchanger differential pressure was measured by a AP transducerconnected across the o11 exlt and return ports on the aft face of the control.
The control vibration was measured by two accelerometers mounted on thepropeller control houslng. One accelerometer was oriented to sense motlon inthe horlzontal dlrectlon and the other to sense motlon in the vertlcal
27
5.1.2 (Continued)
directlon. Additionally, during conduct of the steady pressure test, an
accelerometer sensing vertical motion, was installed on the Scanlvalve
fairing assembly. The "once per revolution" signal was orlglnally planned to
be provided by a pickup mounted on the control and triggered by a target on
the rotating propeller bulkhead. However, the system eventually used was
provided by an eddy current proximity probe targeted on a 59 tooth wheel(on a 60 t(_oth basis) mounted on the propeller shaft. The rotational speed
was averaged over ten revolutions.
5.2 Steady Pressure Measurement System
In addition to employing the instrumentation systems described In section
5.I.I and 5.1.2, a specially designed pneumatic instrumentation system was
used to collect and measure blade airfoil surface steady pressures. Thls
system consisted of a speclally fabricated blade with rows of pressure tapsInstalled at thirteen radial stations and a scanlvalve mounted on the nose of
the Prop-Fan. (See Figure 5-4.) The pressure taps were connected to the
scanlvalve by 36 capillary tubes run along the acutator dome.
The scanivalve Instrumentation system provided 36 channels for transmitting
steady pressure data from the surface of the blade to the stationary field..
The scanIvalve itself consists of a rotating and a stationary portion. The
rotatlng portion was attached to the front of the actuator dome. The radialtubes from the steady pressure measurement blade were connected to the
rotatlng portion of the scanlvalve. Each tube was connected to one channel
of the scanlvalve. The statlonary portion of the scanlvalve contained apressure transducer which monitored one channel of the scanIvalve at a time.
The stationary portion of the valve protruded through the leading edge of thepropeller spinner and was restrained against rotation. Switching of the
scanlvalve channels to be monitored by the transducer was controlled by a
pneumatic signal requiring a clean alr source of 150 psi. The scanning ratewas adjustable from 0.l to lO seconds per channel. The scanivaIve was
enclosed In an aerodynamically shaped fiberglass fairing to maintain a well
behaved inflow to the Prop-Fan. The umbilical, which connected the
scanlvalve wlth the control and monitoring equipment outside the tunnel, was
enclosed In a conduit with an airfoil shaped cross section. This alsominimized disturbance of the inflow to the Prop-Fan.
On 11ne monitoring of the scanlvalve pressure transducer sIgnals was
accompllshed uslng the digital readout from the scanlvalve controller. Inaddltlon to recording the measured pressures on magnetic tape, on-site data
manlpulatlon and prlnt-out generation was provlded by a Fluke DA computer
coupled wlth a plotter and dot matrix printer.
28
OF., I_0 OR QdAL_I ,Y
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29
5.3 Unsteady Pressure Measurement System
Collection of blade surface unsteady pressure data was accomplished by
utilizlng a specially instrumented blade, coupled with the instrumentation
arrangement described earlier in sections 5.l.l and 5,1.2. Twenty-six highfrequency response pressure transducers were |nstalled in two rows on each
side of the blade. The unsteady pressure signals were transmitted from the
rotating to the stationary field through the FM multiplex electronic data
acquisition system. The signals were monitored on the four-channe]
oscilloscope and recorded on the IRIG tape recorder.
L
30
6.0 TEST FACILITY DESCRIPTION
6.1 Wind Tunnel Description
The LAP High Speed Rotor Test was performed in Chariot No. 3 In the SI-MA
Large Wind Tunnel at the Modane-Avrleux Aerothermodynamlc Test Center
operated by the Office National D'Etudes et des Recherches Aerospatlales(ONERA), France (Figures 6-I and 6-2). The tunnel is a continuous, closed
loop, single return, atmospheric facility incorporating bleed slots In the
tunnel walls to a11ow air exchange capability and tunnel circuit screens tomlnlmlze turbulence, as shown In Figures 6-3 and 6-4. Tunnel cooling is
accomplished by alr exchange wlth outside air of up to 20% of the testsection mass flow rate. The tunnel Is driven by two coaxial counter-rotating
Pelton water turbines from a water supply wlth an 865 meter head, sufficient
to generate a maximum of 88MW (I00,000 HP). The test section velocity can be
continuously varied from 8 m/sec to Mach 1.0, well encompassing the range forthis test. The curves of Figure 6-5 glve a representative sample of the
tunnel driving power and Reynolds number as a function of test section Mach
number. Tunnel pressure altitude varies from approximately 1,100 meters
(3,600 ft.) at low speed, up to 6,000 meters (20,000 ft.) at Mach l.O. At
Mach 0.8 the tunnel operates at an effective pressure altitude of
approximately 4,260 meters (14,000 ft.). Stagnation temperature ranges from
-20°C to +60°C, depending on the external ambient temperature and the Machnumber in the test sectlon.
The tunnel has three Interchangeable test sections or chariots which are
positioned in the aerodynamic clrcult. Each of the chariots contain adifferent test section shape and size and are speclallzed for a wide variety
of test capabilities. The use of Individual chariots a11ows a test to beconducted in the tunnel while the next test Is being set-up In another
chariot located In l of 2 adjacent mounting stations. The LAP was installed
In Chariot No. 3 which has a propeller test rlg permanently mounted In its
test section (Figure 6-6). The test section Is 14 meters long, has a circularcross section with a diameter of about 8 meters and non-permeable walls.Several lateral fillers or Inserts were added to the test section specifically
for this test to assure the desired range of Mach numbers could be attained.
They serve the addltional purpose of area compensating for the propeller test
rlg blockage generated by the 970 mm diameter center body. A more detaileddescription of the tunnel facilities is contained in References 18 and 19.
6.2 Test Rig Description
The Prop-Fan was Installed on the test rig as shown In Flgure 6-6. The power
source for the Prop-Fan drive system Is twln Turbomeca gas turbines driving a
common gearbox, enclosed In a streamlined pod located approximately ? metersdownstream of the Prop-Fan plane, as shown In Figure 6-7. The engines were
rated for a combined maximum power of 1000 KW (1341) HP at standard
conditions (59°F, 29.92 In. Hg.) as depicted In Figure 6-8. Velocity
measurements made upstream of the engine Inlets Indicated no disturbance at
the Prop-Fan plane due to the operation of the engines. The gearbox was a
31
o 50 lOOM
\
FIGURE 6-1. MODANE - AVRIEUX AEROTHERMODYNAMIC TEST CENTER
32
ORtGfNAL PAGE
BLACK AND WHITE PHOTOC_RAPH
FIGURE 6-2 S1-MA LARGE ATMOSPHERIC WIND TUNNEL
33
_ 15_.M "-I
/I _. T.oT,_IR L,- - iil_l
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---t-_.-l-.-I_-l-l-.:l_tfl-.-I- , -_i--._ IIl-'b,_JI// I I IL I1.,"1 'i'T'_-FI i , L I". _! !','_',!
. .. ,_" ! ,. IT !r l lltIH i I!l]-._.:_l
I s ,,,ov.o,.E =,-,.R,OTS _ I _!i |i WATER
I "_-_.-_ -- I AIR INLET
F__l_-I i!_
L.__'-° ..... _ :li WATEr' PIPE
. FOR ENGINES !I;AH = 865M
FIGURE 6-3. S1-MA WIND TUNNEL SCHEMATIC
_4
BLACKORIGINAL
ANO WH ITE
p ,_ _, ,_,;-; _ ._
PI-.iOTOGRAPH
FIGURE 6-4. S1-MA WIND TUNNEL NOZZLE INLET "I1B_SE'¢TION
3.5
10
6
e (MWJ
0.5 MACH 1
lOO
8o
60
4o
2o
NOTE: THE CHART ABOVE IS FOR A REFERENCE CHORD
LENGTH OF 0.7 M
FIGURE 6-5. TUNNEL DRIVING POWER AND REYNOLDS NUMBER AS A
FUNCTION OF MACH NUMBER
36
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BLACK AND WHITE PHOTOGRAPH
= .
- O_--POOR QUALITY
FIGURE 6-7. PROP-FAN DRIVE SYSTEM ENGINIES_I_II_IF_.ARBOX
38
POWER
Kwi1200
1oo0
800
600
OPERATION LIMITED
(MAX, CONTINUOUS)
._'____'_"_'_
0 500 I000
MINIMUM PERMISSIBLE
PROP-FAN SPEED
TO 5 MINUTES
..
_NORMAL OPERATING REGION !:._j ....',_;..._,_._'i b
',-_'7;._:_";'_',7:,._,,. -_,-"'_':",_";:_.:.' _':'__.;_;_ t_-_• _,;"_'Z,; ":. :":".'4_';...........,,_....._._..;1";$>,,."_............ ,.7, .- .'-,_; ," ,',-'_ ,..":;'_" "'_'"""%% _;:'_!".,_'_
".::_!,.,::%_._:'_._._,_._'Z'._<_s;"z......_.._..........,._.,..
1500 2000 22.00 2460
PRPM
FIGURE 6-8. PROP-FAN TEST RIG POWER LIMITATIONS
39
6.2 (Continued)
speed reducer providing a reduction in RPM of approximately 3:1 from the
engine to the Prop-Fan. Power was transmitted to the Prop-Fan from the
gearbox through a long drive shaft and a balance (Figure 6-9). The drive
shaft housing was supported by rods and damping cables attached to the tunnelwalls. The particular rod and cable arrangement used In each test was
dependent upon the type of test being run and the number of blades
installed. These arrangements were established during a pre-test vibration
survey conducted In the tunnel by ONERA. Universal joints at each end of thedrive shaft allowed the Prop-Fan to be operated at various inflow angles
relative to the flow through the test section.
A balance and torquemeter, Installed in the drive system, allowed the thrust,
torque, sides forces and bending moments acting on the Prop-Fan to be
measured. A diagram of the balance is shown Is Figure 6-10. Forces andmoments are transmitted through the balance by six strain gaged elements
whlch are referred to as dynamometers. Thrust, yawing moment and pitching
moment are transmitted and measured by the three axial dynamometers (TI, T2
and T3). Balance frlctlon torque and the lateral and vertical forces aretransmitted and measured by the three tangential dynamometers (Zl, Z2 and Y).
The torque supplied to the Prop-Fan is computed by subtracting the measured
friction losses in the balance from the torque measured and transmitted by
the torquemeter. A11 six dynamometers have a capacity of 2000 daN. Two
flexible couplings decouple the transmission of the torque from the thrustmeasurement with one of the couplings being fitted with a gage bridge
(torquemeter).
A stationary aerodynamic fairing or centerbody was installed around the drive
system. This was also commonly referred to as the minimum body. The minimumbody provided a downstream extension of the aerodynamic contour of the
Prop-Fan spinner and was designed to reduce the air veloclty passing through
the root portion of the Prop-Fan rotor. At the inner blade radii thecombination of thick airfoil sections and the large number of blades presents
significant blockage to the flow, which could result in choking if the
velocity was not moderated. The minimum body Is approximatly 35% of the
diameter of the Prop-Fan rotor. A description of the minimum body pressuretap arrangement is presented In Section 7.0.
6.3 Test Riq Modifications
The LAP Prop-Fan High Speed Wind Tunnel Test was completed following the
incorporation of several structural improvements to the propeller test rig
drive system. These modifications, as shown In the comparison of Figures
6-II and 6-12, were added to preclude recurrence of a cone seat fretting
corrosion problem encountered early In the testing, which led to an
unanticipated interruption of the test. In addition, the modifications in
genera! provided a more structurally sound rig design, better able to handle
the loads Imparted by the Prop-Fan.
40
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U.Z0
I-
wI,-
Z
m
0
,2"rID
W
w
43
\
\\\
<
)" k-b. Z< bJ-r _.u) Id_. r_
0 l"
r_ _j
U bIbl ZL 0
0Z
0
OU
_z<
b_.j><
(2b.
c: 0b_wOu
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44
6.3 (Continued)
The most significant modification to the drive system was the elimination of
the cone seat retention configuration on the aft end of the prop shaft. It
was this area which experienced the severe fretting corrosion problem early|n the test. The new design provides for a one piece prop shaft mounted in a
substantially enlarged forward test rig housing incorporating increased
capacity cylindrical roller bearings. Since a safe structural operating
envelope for the Prop-Fan had been defined earlier in the testing, thebalance was removed to provide room for the test rig structural improvements.
45/46
r
7.0 SPINNER AND CENTERBODY DRAG MEASUREMENT
7.1 Test Rig Vibration Survey
Before testing was conducted In the SI-MA wind tunnel, the whirl flutter
stabillty of the Prop-Fan/test rlg system was evaluated. Stability of the
Prop-Fan/test rlg was a major concern because the LAP was the largestassembly ever Installed on the test rlg.
The Prop-Fan was assembled and installed on the propeller drive system in the
tunnel test section in preparation for the vibration survey. To e11minate
the possibility of damaging any of the actual LAP blades during this survey,
blade stubs were installed in place of the blades. Additlonal weights werehung from selected blade stubs to account for the weight difference between
the stubs and blades. The stubs were shimmed to minimize any movement in theblade retention during the shake test.
The survey was conducted using two electromechanlcal shakers positioned at
various locations along the axis of the mlnlmum body. Data was obtained
using two accelerometers mounted on the test rlg force balance Just aft of
the Prop-Fan. The data was used to determine the elgenfrequencles of the
various rlg vibration modes in the horizontal and vertical planes.
The results of the survey confirmed the need for an unsymmetrical rlg supportstructure which pre-test analytical whirl flutter stability studies hadsuggested. Though the vibration test demonstrated that the 4 bladed SR-7L
would operate whirl flutter free, it was agreed that rig damping
characteristics would be closely monitored during all test envelope expansion.In order to incorporate the maximum test rlg resistance to whirl flutter
onset, several different rlg support structure conflguratlons were employed,based on the number of blades Installed, as depicted In Figures 7-8, 8-I,8-2, lO-I and ll-l.
7.2 Wlnd Tunnel Corrections
Immediately following completion of the test rig vibration survey, ONERA
conducted the wlnd tunnel calibration test. The purpose of this calibration
was to remove the effects of the presence of the test rig on the measured
Mach number. The ONERA approach was as follows. First, a compressible flowcalculation was performed to establish the axial velocity distribution at theworking radius of the blade (.6R) through the plane of rotation of the
Prop-Fan. Next, two independent axial velocity surveys were made at the sameradial location, utllizlng a speclal pressure tapped plpe and scanlvalve
system. The results of these surveys are shown in Figures 7-) thru 7-6. The
uncorrected Mach number data represents actual measurements _alken, while thecorrected Mach number data is the difference between the ¢a_¢ulatlons and the
measurements of Flgures 7-I thru 7-6 added to the wall leach m_:mber measured
2 meters (6.56 ft.) upstream of the Prop-Fan plane of rotat_om. Thus, thecorrected Mach number data represents
47
r,,@
¢IO¢o:EDZ
3:U<X.I<
8..,I
Moo= .354
;RRECTED MACH NUMBER
I
UNCORRECTED MACH NUMBER
c
|o00 MM 500 MM
I
500 MM 1000 MM
BODY LENGTH, MILLIMETERS
FIGURE 7-1. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAM
MACH NUMBER .354
48
n,
n.,
en
Z
"!"u
J
u0.I
MACH NUMBER
CORRECTED MACH NUMBER
I
I
1!000 MM 500 MM 0 500 MM
BODY LENGTH, MILLIMETERS
FIGURE 7-2. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAMMACH NUMBER .615
1000 MM
49
a,tO
v
a,
m_r
Z
U<:[
.J<
,1.73
.72
I
.71
MACH NUMBER
MACFI NUMBER
!
,!000 MM 500 MM 0 _ MM
I BODY LENGTI'L MILLIMETERS
FIGURE 7-3. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAMMACH NUMBER .72I
! 000 MM
5O
Eio
¢Id
:i
z=u<2.J
u0,J
,8O
.79
.78
.77
.76 _,
.75
.74
.7_
Moo,, .754 I
I
CORRECTED MACH NUMBER I
""*__UUNCORRECTED MACH NUMBER
I
1000 MM 500 MM 0 500 MM
t"t'*"
I ooo MM
BODY LENGTH, MILLIMETERS
FIGURE 7-4. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAMMACH NUMBER .754
51
U}
m:E
Z
'I"u<:E.J<_J0..I
.SS
.84
.83
.82
.81
.80
.79
.78
,77
.76
MN .788
Iq.
, //,
1000 MM 500 MM
I
0 500 MM
I BODY LENGTH, MILLIMETERS
1000 MM
FIGURE 7-5. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAMMACH NUMBER .788
57
r,*
m_E
Z
Zo<X
,J<u0.J
,gZ
,90 ' Moo" . I
.89 829
.88
.87
.86 I
CORRECTED MACH NUMBER.85
.84
"" I -?---JFo.oo-_o,_o..o,.o.m
,81 _ _"_'_ '"
.80
1000 MM
FIGURE 7-6. AXIAL MACH NUMBER DISTRIBUTION FOR FREESTREAM
MACH NUMBER .829
53
7.2 (Continued)
the Influence of the tunnel walls on the axial velocity distribution. Note
there Is very little difference at the lower Math numbers between thecorrected Mach number data and the uncorrected Mach number data as measured
at the 2 meter wall location. In addition to the velocity calibration
described above, ONERA used the Prandlt-Young correction as discussed in
Reference 20 to further adjust the velocity data before computing an advanceratio, 3. This correction, as defined below, is a compressible correction to
the velocity to account for tunnel wall interference.
l
II l w
V
_4 _1
2(1 + 2 _4)I/2<1 - MN2) I/2
or
Vcoe m 1 - Vmeas
2(I + 2 _4)I/2(I - MN_) I/2
where
and
'1:4
THRUST
p x DISC AREA x V2
PROPELLER DISC AREA
TUNNEL CROSS-SECTION AREA
Note that this correction Is largest where the thrust Is largest and thevelocity the smallest, _.e. low Mach number testing.
54
7.3 Test Rig Corrections
7.3.1 Test Objectives
7.3.1.I To determine the aerodynamic drag on the Prop-Fan spinner as a
function of Mach number for Mach numbers ranglng from .2 to .85.
7.3.1.2 To determine the aerodynamic pressure drag on the test rigcenterbody as a Function of Mach number for Mach numbers ranging From .2 to.85.
7.3.2 Test Procedure
In preparation for the splnner/body tare test, the test rlg and associatedsupport structure were set up as depicted In Figures 7-7 and 7-8. In order
to isolate and evaluate the affects of the presence of the spinner and center-
body In the flow fleld, the LAP blades were replaced with blade stubs, whose
external contours were machined to match that of the spinner (Figure 7-9).
As defined in Figures 7-I0 and 7-II, the Forward end of the centerbody serves
as an extension of the external contour of the spinner, and was designed to
alleviate compressibility losses in the blade root section by reducing the
air velocity passing through the central portion of the Prop-Fan rotor. Atthe inner blade radii, the combination of thick airfoil sections and the
large number of blades presents significant blockage which could lead tochoked flow. The moderation of velocity caused by the centerbody reduces the
possibility of choked flow occurlng at the high subsonic Mach numbers at
which the LAP operates. Figures 7-I0 and 7-II also define the location of
the centerbody surface static pressure taps. There are four rows of
twenty-nlne taps per row positioned a]ong the length of the centerbodysurface. The rows of taps were spaced 90° apart, clrcumferentially.
Figure 7-12 defines the location of an additlonal set of static pressure taps
In the space between the spinner rear bulkhead and centerbody. These tapsconsisted of four rows, ten taps per row, extending radially outward, located
at the centers of equal areas. Each row was oriented 90° apart from another.
Collection of the static pressure measurements from the locations describedabove, while varylng tunnel Mach number, was accomplished u_ing a scanivalve
system. Splnner/centerbody drag data was collected at a total of 42 test
points. All of the data was collected at zero degree Inflow angle.Table 7-I lists the Mach numbers at which data was collected.
55
_g
FIGURE 7-7. SPINNER/BODY TARE TEST
56 .... v'_Pt_i _- PHO'i O&RAPI_I
_tALANCE
GEOMETRICAL
CENTER
BLADE PLAN _
AXIS
WIND TUNN___ ._vO -__513.4"-_
ROD
(_:35 ROD
L:47Z0 _ _1 ._:26 RODSiL._-,o,o _ _::':,o
_o tTi_oB cNOTE: ALL DIMENSIONS IN MILLIMETERS
_) = DIAMETER
L = LENGTH
FIGURE 7-8. TEST RIG SUPPORT STRUCTURE FOR SPINNER/BODY TARE TEST
5?
<Id
_->_0
Z<
r
v
m_
E_
=_XXXX
_ .zzzzz
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58
A
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i tN
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6 ........... ¢z <
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N
T_F_
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F_
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r_
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m_mm_mmm_mv
== _oo_oooo_o:::::::
I
: m v" _" =', m _ _= _" N N N N N N N N N N
Z
oI-U
0
<
Ob.
0
Z0
b,
0
,¢
-r<U 0
gl Ui.-r <
z0m
H-ULd
]-.1.6<
z
r_:E(/)Z0
H-
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hi
E
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I,Ll
<b.
m
1:1
I-ZLd
i
r_
0
60
w
utZ
I-u<0I-.l
I-
(_ NNNNNNNNNIq
i: ..........
0 0
\
m
_m0 N
N m ___m. _. _. m. m ?.--me, iN N N _l (q (q eq _ m |
o o 0 ooo0 00_ I
Z
oI-<u0.J
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W
fflUl
I1.
Um
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r_
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61
Table 7-I. Splnner/Centerbody Drag Test Points
SPINNER/CENTERBODYDRAG CORRECTED
TEST MACH
PT. NO.
24224424624825O2552572602622642662682702"/22742762872882892902912932942952962972993OO3023033053O830931031l312314315316317318
499
789
836
834
833
8OO848
.789
.739
.686
.638
•590
•494
•447
•348.244
.201
.201
.201
.201201298298298298298494494590590494
.298
.298
.298
.299
.298
.201.201.201.201.201
62
7.3.3 Discussion and Results
7.3.3.1 Spinner Draq
As depicted in Figure 7-13, the spinner drag (Ds) was measured directly from
the axial force measured by the balance (Fs) with corrections for the backpressure force (FBp) and the losses due to thermal effects in the flexible
coupling (FTH).
Ds = FB - FBp - FTH
The axial force applied to the propeller shaft (F,) was measured by three
dynamometers T,, Tz and T3 In the test rig balance as depicted In
Figure 6-I0. The losses (FTH) due to the thermal effects of the flexible
coupling were measured by strain gages bonded directly to the flex coupling.The back pressure force (FBp) iS the result of the difference between the
free stream static pressure (Po> and the integrated pressures fn the space
between the rear of the spinner bulkhead and the face of the centerbody
(PN), where aN IS an area weighting factor and A, Is the spinner basearea.
4O
FBp = ( _ (aNPN) -- Po) AsN=I
As - .507 mz
Figure 7-14 presents the spinner drag coefficlent (Cso) as a function of
corrected Mach number (McoR). The spinner drag coefficient Is given below
and Is expressed In terms of the spinner drag force (Ds), dynamic pressure(qo) and projected spinner base area (As).
Ds
CsD
qoAs2
As = .519 m
Since the spinner drag data was corrected for the back pressure effects, itrepresents the axlal components of the forces appl|ed by the static pressure
acting normal to the spinner surface and the spinner boundary layer wall
stress acting on the wetted area of the spinner. A quadratic equation wasfitted to the spinner drag data, as shown below and on Figure 7-14, and was
used to compute spinner drag corrections to measured thrust during Prop-Fan
performance testing.
Cso - 2.7699M" - 5.3225M 3 + 3.7727M z - 1.1215M + .2398co. co. co. co.
63
DCBT
•",.'OE.E.U..--T_,/ OS'_,_°P_T.(FTH) DCBT " f(Ps -- PO)dA
FLEXIBLE COUPLING LOSSES
A. SPINNER/CENTERBODY TARE TEST (WITHOUT BLADES)
TApP / / DCB
BALANCE j
MEASUREMENT (FB)
B. PROP FAN TEST PT. (WITH BLADES)
FIGURE 7-13, PROP-FAN TEST RIG FORCES
64
m
o _Ju
1,1,1m
Z
L_<
r_
r_0U
>
Z
u
<
r_
65
7.3.3.1 (Continued)
The spinner drag coefficient, as expected, increased with Mach number forMach numbers in the range from .3 to .85. However, an unexpected decrease indrag coefficient was observed between Mach .2 and Mach .3. The decrease indrag coefficient cannot be readily explained from the data that wascollected. Theoretically, the drag should vary as the square of thevelocity. The large number of data points taken at Mach .2 and Mach .3 doesconfirm that the decrease in spinner drag coefflcient is a real phenomenon.
7.3.3.2 Centerbody Drag
As mentioned earlier, the Prop-Fan was tested in the presence of a centerbody
which was designed to alleviate compressibility losses in the blade rootsections. With the force measurement as shown in Figure 7-13, it has been
shown that the Prop-Fan net thrust (TNET) cannot be directly measured onthe force balance (Reference 21). This Is true because, as discussed in
References 21 and 22, the thrust of the Prop-Fan blades changes the pressure
acting on the centerbody, thereby changing the pressure drag. The presence ofthe body also causes an increase in thrust on the rotor equal in magnitude to
the change In the centerbody drag. The change In centerbody drag is commonlyreferred to as the buoyancy force (BF). The measured Prop-Fan thrust has
been classically referred to as apparent thrust (TApp) and is the largestforce component sensed by the balance. Since the increase in centerbody drag
negates the increase In Prop-Fan thrust, there Is no net increase In thrust
produced by the Prop-Fan system. However, the thrust measurement (Fs), In
Figure 7-13 (Views A and B), does not sense centerbody drag, therefore,
measured or apparent trust (TApp) iS corrected by subtracting the buoyancyforce, which Is the difference between the centerbody drag measured at the
Prop-Fan operating point of interest (DcB) and the centerbody drag measuredat the same Mach number without blades (DcBT),
The centerbody pressure drag (DcBT), was determined by pressure integration
of the longitudinal rows of area-weighted static pressure taps. The
integratlon requires a simple summation because the pressures are measured in
the centers of equal annular areas.
44
DCBT " X (PN - Po)AN + E(PM - Po)AMN=I
WHERE: Po = FREE STREAM STATIC PRESSURE
PN, PN " STATIC PRESSURE AT TAP N, MAN : INCREMENTAL FORWARD CENTERBODY AREA AT TAP N
AM - INCREMENTAL AFT CENTERBODY AREA AT TAP M
66
7.3.3.2 (Continued)
As a result of several last minute profile modifications, by ONERA, to theaft end of the Prop-Fan test rig centerbody, the local static pressure tapspreviously requested by Hamilton Standard were omitted. Therefore, thebuoyancy force correction applied to the apparent thrust is In error.However, the change In the buoyancy force that would have occurred is smallenough to result in no significant effect to the data. Therefore, theequation above is further simplified to the following:
44
DCBT = _ (PN - Po)ANN:I
Figure 7-13 presents the variation of centerbody drag coefficient withoutblades (CcBoT) with corrected Mach number (McoR)- The centerbody dragcoefficient without blades Is given below and is expressed in terms ofcenterbody drag force without blades (DcsT), dynamic pressure (qo) andprojected centerbody area (Acs).
CCBDT
DCBT
qo AcBAcB = .22 m2
A quadratic equation was fitted to the centerbody drag, as shown below and on
Figure 7-15 and was used to compute the buoyancy force correctlon to measured
or apparent thrust during Prop-Fan performance testing.
CcsoT = 7.6393M" - 16.786M 3 + 12.124M 2 - 3.7165M - .2320COR COR COR COR
It is observed from the data that the centerbody drag coefficient decreases
with increasing Mach number. This indicated that the centerbody surface Mach
numbers are Increaslng at a faster rate than the free stream velocity.
7,3.3.3 Performance Corrections
With the Prop-Fan blades installed and thrusting, as depicted In Figure 7-13,
the balance measures the algebraic sum of the apparent thrust, the spinner
drag, and the back pressure force. Therefore, the apparent thrust of the
Prop-Fan Is obtalned as shown In the foIlowlng equation:
T, pp = Fa - F.p + Ds + FT,
67
./
/
ONMN
n,,0U
IE
I
./
,,¢,
0_u
IE
N
÷
n,.o
_u
tO
In,,0
-IE
u_
H
I--QmUU
?m.?
I-amU
U
?
o
c_
r_0u
o
0
r_Wr_
Z
U<
w
UI,I
F,,OU
i-Z[di
O
_D<n-
>OOmn,I,di-ZLdU
_K
ILl
_D
z
68
7.3.3.3 (Continued)
As defined earlier, the centerbody drag is obtalned from centerbody surfacepressure integrations"
44
DCB = E (PN - Po)ANN:!
Also, as mentioned previously, the buoyancy force was obtalned from thedifference between these and the tare run pressure Integratlons"
BF = DcB - DCBT
F|nally, the net thrust (TNET), which is defined as the propulsive force ofthe blades operating In the presence of the spinner and centerbody flow fieldwlthout the increase in thrust due to the mutual interaction, is obtained bysubtracting the buoyancy force from the apparent thrust:
TNET = TApp - BF
69/70
8.0 BLADE STRUCTURAL DYNAMIC EVALUATION
8.1 Test Objectives
8.1.1 Confirm that the SR-TL Prop-Fan was free of high speed blade flutter
over the portion of its operating envelope that could be run in the ONERASI-MA wind tunnel.
8.1.2 Evaluate the IP blade strain sensitivity of the SR-7L Prop-Fan for a
range of blade angles, rotational speeds, Mach Numbers and inflow angles.
8.1.3 Compare the measured and analytically predicted IP blade strain
responses of the SR-TL Prop-Fan for selected operating conditions.
8.2 Test Procedure
Testing of the SR-7L Large Scale Advanced Prop-Fan was conducted In the ONERASl atmospheric wlnd tunnel in Modane, France. The Prop-Fan was mounted so
that the rotor plane was located in the throat of the wind tunnel. The
tunnel throat was eight meters in diameter.
The drive system as described earlier in section 6.2, was supported by rodsand cables attached to the tunnel walls as illustrated In Figures 8-I and 8-2.
A balance and torquemeter, also described in section 6.2, provided the
capability to measure forces and moments acting on the Prop-Fan.
Test rig vibration was monitored by two sets of horizontal and verticalaccelerometers. The accelerometers were located on the drive train housing
in two planes aft of the Prop-Fan.
A stationary aerodynamlc falrlng was located downstream of the Prop-Fan. The
fairlng served as an extension of the external aerodynamic contour of thesplnner. The fairing resulted in an approximate 35% blockage of the flow
through the Prop-Fan rotor.
Power available from the test drive system was significantly ]ower than the
rated power of the Prop-Fan. Therefore, In order to simulate operation at
high power loading conditions, the Prop-Fan was run in two and four bladeconfiguratlons as well as with eight blades. Thls allowed power loadings perblade to be achieved, which correspond to intermediate and high power
operating points with elght blades. The disadvantage of operation with twoand four blades was the negation of the Inter-blade cascade effects which are
present in the eight blade design. These effects tend to be destabilizing inthat they lower the Mach Number at which the onset of classical flutter
occurs.
The mlsslng blades were replaced with stubs in the two and four blade
conflguratlon as depicted in Figure 7-9. The ends of the stubs were machlnedto match the external contour of the spinner. The Prop-Fan is shown in the
two, four and elght blade conflguratlons in Figures 8-3, 8-4 and 8-5.
71
?2
\
tnr_
==.J
Z
u_Z0
oZ
_3
FIGURE 8-3. TWO BLADE TEST CONFIGURAI"IOIM
74
ORIGINAL PAGE
BLACK AND WHITE PHOTOGRAPH
FIGURE 8-4. FOUR BLADE TEST CONFIGURATION
Bi_ACK #,_,;u WHITE PHOTOGRAPII75
FIGURE 8-5. EIGHT BLADE TEST CONFIGURATION
?6
B._'_._ /_,--_u li_,. L PHOT_O_RAPI-f
8.2 (Continued)
The strain gage arrangements for the two, four and eight blade configurations
are shown in Figures 8-6, 8-7 and 8-8. The blade surface gage locations werechosen either to correspond with the points of maximum strain for the blade
normal modes or to provide the distribution of strain along the entire span
of the blade. Blade shank gages were also used to measure vibratory bending
moments.
The Prop-Fan was operated in a Beta control mode during the high speed windtunnel testing. In this mode the blade pltch angle was selected by the
operator uslng a manual control. The test procedure consisted of starting
the Prop-Fan at a blade angle (B3/4) of 20° to 25° and increaslng power toobtain 1200 RPM to 1500 RPM rotor speed with the tunnel flow drive system not
operating. Mach number was then increased in increments. Foliowlng eachincrease In Mach number the blade angle was Increased to maintaln the rotor
speed In the 1200 RPM to 1500 RPM range. At the Mach numbers of interest,
test points were run at two or three different power settings and over a
range of rotor speeds. Blade straln gage data was recorded for thirty seconds
at each test point. Aerodynamic performance and test rig vibration data were
also logged concurrently with the strain gage data.
8.3 Discussion and Results
8.3.1 General Discussion
During initial balance runs of the Prop-Fan, which was conducted at zero Math
number, test rlg critical speeds were discovered at 360 RPM and 540 RPM.Both crlticals were highly undamped, allowing rig vibration to grow rapidly
if operation was attempted at these speeds. The critical speed at 360 RPM
corresponded to the predicted rig first critical. Since both of thesecrlticals were well below the planned test operating speeds, they did not
pose a problem. No other critical speeds were apparent within the testrotational speed range. For the low rotational speeds at which dynamic
balancing was accomplished, dynamometer vibratory stresses were the limitlngfactor rather than rlg vibrations measured by the accelerometers. This may
have resulted from relatively low accelerations causing large displacements
due to the low frequency of the response.
Figures 8-9, 8-10 and 8-II present a mapping of the test points run duringstructural dynamic testlng. The test points acquired for the four blade
configuration spanned the entire planned operating range. Power supplied to
the Prop-Fan by the turbines was limited to 800 KW due to the elevated
ambient temperature in the tunnel.
Operation between lO00 RPM and 1500 RPM resulted in high rig vibration forthe two blade and the eight blade configurations. Therefore, this operating
region had to be avoided. The vibration frequency was not IP and was not
believed to be caused by the Prop-Fan. With the exception of the 2 bladed
configuration at 3° inflow angle (_), operation at inflow angles of 3° orlO: also resulted In a high level, low frequency vibration which precluded
testing at those conditions.
77
FORWARD TO AFT VIEW
®
BLADE CAMBER SIDE VIEW
3F
®
GAGE r/R TIP
! .44
2 .77
3 .64
4 .69
5 .71
6 .84
7 .71
8 .84
9 .34
I0 .44
I 1 .44
12 .57
13 .78
GAGE NUMBERING CONVENTION - X XX
FIGURE 8-6. STRAIN GAGE ARRANGEMENT FOR TWO BLADE CONFIGURATION
78
®I
FORWARD TO AFT VIEW
®
BLADE CAMBER SIDE VIEW
5F
GAGE r/R TIP
! .44
Z ,77
3 .84
4 .69
5 .71
6 .84
7 .71
8 .84
9 .34
10 .44
I I .44
12 .57
I 3 .78
FIGURE 8-7.
GAGE NUMBERING CONVENTION -X XX
STRAIN GAGE ARRANGEMENT FOR FOUR BLADE CONFIGURATION
79
®!
FORWARD TO AFT VIEW
® /
®
o,,;BLADE CAMBER SIDE VIEW Z .77
3 .84
4 .69
5 .7!
6 .84
7 .71
8 .84
g ,34
10 .44
1 1 .44
12 .57
! 3 .78
GAGE NUMBERING CONVENTION -X XX
B-A°E.o!.GAGE NO/
FIGURE 8-8. STRAIN GAGE ARRANGEMENT FOR EIGHT BLADE CONFIGURATION
8O
1000
8OO
SO0
qoo
20O
8OO L06O
SR-7L 2 NRT TEST POINTS
+
×
+
Z_
12O0 I qO0 t600
RPH
1800 2000
3000
25OO
ici
z 2000
o
t 500
i
f 1000
SR-'7L 2-HI::I'I" TEST DFITR(HODFINE]
; cp
t_
O (
poO
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FIGURE 8-9, TEST POINTS AND BLADE SHANK MOMENTS (2 BLADE CONFIGURATION)
81
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82
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FIGURE 8"1 1. TEST POINTS AND BLADE SHANK MOMENTS (8 BLADE CONFIGURATION)
83
8.3.1 (Contlnued)
Mach numbers above .73 resulted in negative thrust for the eight blade
configuration due to the low power available from the turbines. Since the
balance was not designed for negative thrust, the maximum Mach number for the
eight blade case was limited.
8.3.2 Data Reduction
The strain gage signals were recorded on magnetic tape for each of the gages
shown In Figures 8-6, 8-7 and 8-8. These signals were then fed through a
peak stress (strain) converter, a device which determines the peak vibratory
strain occurring in a unlt of time (.l seconds for thls data). The output ofthe peak stress (strain) converter is then statistically analyzed to obtain
the average amplitude (mean) and standard deviation of each signal. The IRP
(infrequently repeating peak) Is defined, for the purposes of this report, as
the mean plus twice the standard deviation. It Is a conservative measure of
the strain amplitude normally used to estimate blade fatigue life.
Figures 8-9, 8-10 and 8-II show summaries of the conditions analyzed along
with plots of the IRP shank moments for the 2, 4, and 8 bladed configurations.It can be seen that the highest values recorded were for the two bladed
configuration wlth angular inflow (3 degrees). It Is also noted that there
are significant blade-to-blade dlfferences. However, it was found that the
data contained significant high frequency noise (several thousand HZ), whichtended to artificially increase the IRP values. In general, no significant
differences In strain gage signal were revealed between the two, four and
eight blade configurations at uniform inflow angle, for the same gage (3F).
Figure 8-12 shows an example of the signal to noise problem and itsvariability between different gages. Because of the low response levels, the
data reduction was not repeated with the noise filtered out.
Tables 8-I, 8-11, and 8-111 show the conditions of the runs made along with
some sample IRP tabulations, showing blade-to-blade variations.
Table 8-1V shows a tabulation of peak IRP values for each gage versus thelimits determined based on the threshold at which blade fatigue damage begins
to accumulate. Because of the high noise content, all of the IRP values
(Figures 8-9, 8-10, 8-II and Tables 8-I, B-If, 8-IIl and B-IV should beinterpreted as conservative upper bounds on the strains actually felt by the
blades. As such, the gages shown to be exceeding the Iimits should actually
be Interpreted as below the limits with the noise filtered out.
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ORIGINAL PAGE I_.
OF POOR QUALIT_
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89
Gage Description
TABLE 8-]_Z'. MEASURED VIBRATORY PEAKS VS. LIMITS
Peak IRPStrain observed
for 8-way(0' inflow)
fraction o! limit
Peak IRPStrain observed
for 4-way(O' inflow)
fraction of limit
Peak IRPStrain observed
for 2- way(O' inflow)
fraction of limit
Peak IRPStain observed
for 2- way(3' inflow)
fraction of limit
Radial Bending (44%r/R) Gage 1 .43 .33 .79
Radial Bending (71%r/R) Gage2 1.12" .48 1.05"
Radial Bending (84%r/R) Gage3 ,22 .45 .46 .88
Trailing Edge Bending Gage 4 .55 .6 1.01 *T. ,..
Shear (71%r/R) Gage5 .04 .07 .05 .11, i
Shear (84%r/R) Gage 6 .05 .07 .08 .12
Chordwlae (71%r/R) Gage7 .22 ,21 .16 .26
Chordwise (84%r/R) Gage8 .36 .23 .20 .33
Radial Bending (34%r/R) Gage 9 .41 .30 .72
Shear (44%r/R) Gage 10
Chordwise (44%r/R) Gage 11 .24
Radial Bending(57%r/R) Gage 12 .26 .45 ,49 1.11 *
Radial Bending(78%r/R) Gage 13 .29 .29
Flatwise Shank Moment - Gage F .29 .47
,22EdgewiseShank Moment - Gage E
.52 1.03"
.76.28
.17.19 .27
* It is fell that because of the noise present, these values can be Interpretedas within limits.
Notes: 1. The gage limits were set based on the strains reachinga Ikmslmld at whichblade fatigue damage begins to accumulate (vibratory strain _lmrimposedupon steady strain).
2. Becauseof the amount of noise in the strain signals, the Blipvalues listed here should he interpreted as conservativeu_ Immds onthestrains actually felt by the blades.
9O
ORIGINAI_ PAGE IS
OE POOR QUALIT_
8.3.2.1 Angular Inflow
As previously noted in section 2.0 and paragraph 8.3.2.1, the only angularinflow data obtained was for the two bladed configuration at 3° inflow
angle. A spectral analysis of each strain gage signal was performed using an
FFT (Fast Fourier Transform) algorithm. This analysis employed a Hanning
window to help overcome leakage effects. Correlatlons were then made with
the predictions of IP response.
Table 8-V shows the P-order content obtained for the 3 degree conditions
tested (shank moment and radial bending). It can be seen that the 2Pcomponent averaged about 25% of the IP component (with 2P/IP approximately
.54 near 1500 RPM) and that the 3P component was generally much smaller. It
Is noted that at least a portion of the excitation force driving the 2P
response may have come from the test rlg drive system. Twice per revolutionvibration Is characteristic of a shaft with universal joints. It is also
likely that nonlinearities In the aerodynamic excitation, possibly due to
observed vortex loading phenomena, or tunnel flow irregularities in the test
section, are causes of higher order excitation. The aerodynamic analyses
used for the predictions do not consider this effect.
The data digitized for the spectral analysis was sampled at constant time
steps. Even small variations in RPM are known to result in substantially
lowered P-order magnitudes. By synchronously sampling the data (digitizing
at constant fractions of a cycle instead of constant time step) this problemcan be overcome. This is known as a 'speed corrected' spectral analysis.
However, since the RPM trace was not available for this test, this was not
possible here. An alternative procedure was used for several strain gage
results to double check the amplitudes obtained from the non 'speed
corrected' spectral analysis. A calculation was made of the averageamplitude of a band pass filtered (IP frequency mid band) strain record.
This procedure is cumbersome but should produce conservative strains because
the amplitudes are those generated by a peak stress converter (with a finitereset rate). Table 8-VI shows that the band pass filtered values are
typically I0% higher than the spectral values. Even with no RPM variation,
the spectral magnitudes are known to be up to 16% low, due to the effect
known as leakage (with a Hannlng window as employed here) In digital signal
processing terminology. It is concluded that the actual magnitudes arewithin lot (higher) of the spectral values.
Figures B-13 and B-14 show the trends of IP response (shank moment and blade
bending) with respect to power, RPM and Math number for the 2 bladedconfiguration at 3° inflow angle. As expected, the response increased with
power and Mach number. Variations of RPM are influenced by system
resonances. Figure B-15 shows the variation of bending strain with spanwlselocation. Again, the trends are believable and consistent with the data
collected during the four and eight blade structural testing.
91
ORIGINAl5 PAGE Ig
OE POOR QUALITY
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96
8.3.2.1 (Continued)
The plotted values were normalized wlth respect to E.F. (excitation factor)which Is defined as:
E.F. = _ (VT1644.8)2(p/po)
Where _ is the inflow angle In degrees, VT the true airspeed In Km/HR,
p the alr density, and po the alr density at standard sea levelconditions.
Only the IP component of the response has been plotted to investigate trends.
The IRP values include the higher order content as well as unfiltered noise.
8.3.2.2 Analytical Method
A finite element model of the SR-7L blade was developed during the design
phase. The same model was used here to predict IP response for correlation
with test. The steady state aerodynamic loads were calculated using the HS
computer code H444. The steady alr loads along wlth centrifugal loads, wereapplied to the model which was analyzed using the in-house finite element
code, BESTRAN (Reference 23). From this solution the differentlal stiffnessmatrix was obtained and added to the structural stiffness matrix. The
dlfferential stiffness matrix represents the addltlonal stiffness of theblade under centrlfugal loading. The unsteady (IP) alrloads were then
calculated uslng the HS computer code H337. The loads were applied to the
finite element model and displacements and surface strains predicted.
Mohr's circle relationships were employed to calculate strains In thedirection of the gages for correlation with test. Shank moments were
calculated us|ng the reaction loads at the root of the blade and assuming a
linear variation of moment up to the shank location.
The H444 and H337 codes are both aerodynamic 'strip' analyses that have been
calibrated to predict the steady and unsteady alrloads on swept Prop-Fan
blades. The H444 code employs a Goldsteln formulatlon to calculate inducedvelocltles, whereas the H337 code uses a skewed wake theory. Two-dlmenslonal,
compressible alrfoll data was used by both codes to predict the aerodynamic
loads. Trlangular plate elements were used In the BESTRAN code to do the
finite element analysls. Each component of the composite blade (i.e., shells,spar, foam, sheath) were represented by separate plate elements. They were
a11 tled together using constraints based on plane-sectlons remaining plane.Further detail on the codes and their use can be found In the LAP Blade
Design Report (Reference 14).
Figure 8-16 Illustrates a Campbell diagram of the SR-7L blade. Shown are
BESTRAN frequencies compared wlth test values. It is assumed that the RPM
range tested was far enough away from IP resonance that the Influence ofdamping on strain magnitudes can be neglected.
97
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FIGURE 8-16 COMPARISON OF SR-7L NATURAL FREQUENCYTEST RESULTS TO PREDICTIONS
9B
@.3.2.3 Test vs. Analysis
Three cases were chosen to analyze and compare to test. The condltions werepicked to study the influence of power and Mach number variations on the
predicted strains. Case l (run #I176) was at 1673 RPM, 794 KH, .5 MN.
Case 2 (run #1190) was at 1698 RPM, 244 KN, .5 MN. Case 3 (run #1214) wasat 1707 RPM, 829 KW, .724 MN.
The IP gage strains were calculated using the methods outllned in the
prevlous section. Comparisons were then made to the values obtained from a
spectral analysls of the test data. See Figures 8-6, 8-7 and 8-8 for gage
locations. Figure 8-17 shows the variation of radial bending straln as afunction of spanwise 1ocatlon. As can be seen the correlation is quite
reasonable wlth the In-board predicted strains being about 30% too high.
Nowever, as previously noted in paragraph 8.3.2.1, the 'measured' strain
levels are felt to be up to I0% low.
Figure 8-18 shows the correlatlon of shank moments and root radial bending
(gages F and l) as a function of power, (holding RPM and Mach number). Both
the test and analysls show the expected trend of increased IP response wlth
Increased power. The test values are consistently below those calculated.
Figure 8-19 shows the same quantities plotted against Mach number (holdingRPM and power). The trends are consistent with expectations (higher response
with increased Mach number). Table 8-VII shows a comparison of all the
strain gage results for the three conditions. As noted from the plots, thecorrelation with radial bending strains and shank moments Is quite
reasonable. However, the correlation Is rather poor with the gage placed
near the trailing edge, and also with the chordwlse and shear gages. This Is
consistent wlth previous experience. It |s felt that these values are
affected more by local distributions of aerodynamic loads, whereas the radial
gages are strained by an integration of loads outboard of a given gage.Further study Is needed to better define the local load distributions In
order to more closely correlate wlth the measured strains.
8.3.2.4 Blade Flutter
Prior to the high speed wlnd tunnel test, an unstalled flutter analysis wasconducted for the two, four and eight bladed Prop-Fan configurations. The
analysis indicated that the Prop-Fan would be free from unstalled flutter
throughout the wlnd tunnel test operating envelope. The tests confirmed the
predictions In that no flutter or tendency to flutter was measured orobserved. The wlnd tunnel test results were also a good Indlcator that no
unstalled flutter would occur durlng the subsequent flight test program.
Though the maximum Mach number achieved for the elght bladed configuration
was 11mlted to .73, the resultlng dynamic pressure at the tunnel's effectivepressure altltude of 4,360 meters (14,OOO ft.) Is greater than the maximum
dynamic pressure expected for the f11ght test.
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FIGURE 8-19. TEST VS. ANALYSIS (EFFECT OF MACH NO.)
102
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8.3.2.4 (Continued)
The maximum Mach numbers achieved for the two and four bladed Prop-Fan
configurations were .80 and .84, respectively. Though these configurationsdid not permit the effects of blade cascade to be evaluated, they did
demonstrate the Prop-Fan's resistance to unstalled flutter onset for
subsequent blade surface pressure testing.
8.3.2.5 Whirl Flutter
Due to the flexibility of the ONERA test rig, whirl flutter was a concern forthis test. These concerns were reinforced when initial calculations of the
installations' stability showed whlrI flutter to occur within the operating
envelope. For the Inltlai computations, the forward support structure, shownIn vlew A of Figure 8-I and views A and B in Figure 8-2, had a three rod
configuration that resulted In undesireable symetrical pitch and yaw
stiffness. To stabilize the two and four bladed configurations, a two rod
support structure, as depicted In vlew A of Figure 8-I, was utilized. Thetwo rod configuration, which provided asymetrlcal pitch and yaw stiffness,
resulted In a more stabilized system. The computations also demonstrated the
need to add cables, as shown In view A of Figure 8-2, for the eight bladed
configuration to increase structural damping while improving the distributionof the pitch and yaw stiffness. The end result of these rig modifications
was successful completion of the wind tunnel test with no whirl flutterinstabilities encountered.
104
9.0 AERODYNAMIC PERFORMANCE EVALUATION
9.1 Test Objective
To measure the aerodynamic performance of the SR-7L Prop-Fan for a range of
blade ang]es, rotational speeds and Math numbers.
9.2 Test Procedure
Aerodynamic performance data (thrust and power) was measured concurrentlywith the structural dynamic data. As a result, the data Includes losses In
performance resulting from the presence of the strain gages and strain gage
wires on the aerodynamic surfaces of the blade. Installation of the gages
was accomplished so as to minimize these losses.
Power avallable from the two turbine engines driving the Prop-Fan in the wlnd
tunnel was slgnlficantly lower than the rated power of the Prop-Fan.
Therefore, testing was conducted using two and' four blade configurations as
well as eight blades.. The two and four blade configurations permitted
operation at power loadlngs per blade that correspond to high and Intermediate
power operating points respectlvely for the eight blade Prop-Fan deslgn. Thedisadvantage to thls approach is that the eight blade performance cannot be
easily extrapolated from the two or four blade test results. This is due to
the aerodynamic Interact|on between the blades. This interaction is more
significant for the eight blade design than for the two or four bladeconflguratlons, due to the reduced spacing between blades. Unfortunately the
aerodynamic performance data for the 2 blade configuration was llmlted, and
rather than attempt to project the full power elght blade performance from thethe two and four blade test results, the available results are compared to
analytical predictions. This serves to verify the analytical techniques andprovides confidence that the predictions for the eight blade high power
performance may also be correct.
The missing blades were replaced with stubs (see Figure 7-9) in the two and
four blade configurations. The ends of the stubs were machlned to match theexternal contour of the spinner.
The Prop-Fan was operated in the Beta Control mode during the aerodynamlc
performance testing. In this mode, Hamilton Standard personnel were able tochange the blade pitch angle during testing by means of an Increase/decrease
pltch switch located in the control room. For a fixed Mach number and a
constant power supplied by the turbines, the Prop-Fan rotational speed was
varied by increasing or decreasing blade pitch angle. At the Mach numbers of
interest, aerodynamic performance data was collected for two or threedlfferent power settings and over a range of rotational speeds.
Approx|mately 140 performance data points were collected.
105
9.3 Discussion and Results
9.3.] Data Reduction Procedure
The blade pitch angle (B314) was measured and recorded for each test point.
An electrical signal proportional to blade angle was provided by means of a
potentiometer mounted on the rotating side of the Prop-Fan. The potentiometershaft was positioned by the rotation of the No. 7 blade retalning ring
through a cable and pulley arrangement (Figure 5-3). Flgure 9-I demonstrates
the linearlty of the output of the blade angle measurement potentiometer as
monitored during a pre-test calibration check.
The Prop-Fan rotational speed was measured by use of a IP pickup. The sensor
was triggered by a gear _K)unted on the test rig drive shaft. The rotational
speed was averaged over ten revolutions.
The power absorbed by the Prop-Fan (BHP) was determined by multiplying thetorque supplied to the Prop-Fan (QcoR) by the rotatlonal speed (N). Torque
supplied to the Prop-Fan (QcoR) was computed by subtracting the measuredfrictional losses in the balance (Q_L) from the torque measured by the
torquemeter (QMEAS).
BHP = N x QcoR
Where: OcoR - OM_s - Q_L
The net trust (QNET) determlned during testing is the unlnstalled thrust of
the Prop-Fan rotor, operatlng in the presence of a spinner and centerbody andIs computed from the following equation"
QNET = F8 + Ft. - Fs, + Ds - BF
Where"
F, = axial force measured by the balanceFTH = losses due to thermal effects
F,p = back pressure force
Ds = spinner dragBF - buoyancy force
The temperature correction term (FTH) compensates for the effects of
changes In temperature on the balance strain gages.
The back pressure term (FBp) corrects for the increase In measured thrust
due to the differential between the pressure behind the spinner bulkhead and
the free stream pressure. The back pressure force Is calculated by
multlplylng the difference between the average pressure measured by the tapsshown In Figure 7-I0 and the free steam pressure by the pro_ected area of the
spinner bulkhead.
106
100.
-ZOO0.
I
-1500. -I000.
7_
BLAD£ ANGL£ (_3/4)
-500. O. 500.
PO'rENTIOMETER OUTPUT (mV)
!000. 1500. 2000.
FIGURE 9-1. LAP BLADE ANGLE POTENTIOMETER OUTPUT
107
9.3.1 (Continued)
The splnner drag force (Ds) is computed by multiplying the spinner drag
coefficient by the free stream dynamic pressure and the reference area of the
spinner, as defined in paragraph 7.3.3.1. The spinner drag coefficient was
determined as a function of Mach number during the spinner drag test
(Figure 7-14).
The buoyancy force term (BF) eliminates the apparent Increase in thrust
caused by the interaction between the Prop-Fan rotor and the centerbody. The
buoyancy force is determined by measuring the centerbody drag at theperformance test point of interest and subtracting the centerbody drag without
blades for the same Mach number. The centerbody drag was determined during
performance testing by integrating the difference between the pressures
measured by the taps shown in Figure 7-I0 and the free stream static pressure
over the surface of the centerbody. (See paragraph 7.3.3.2). The centerbodydrag without blades was computed by multlplylng the centerbody drag coeffi-
c|ent by the free stream dynamic pressure and the centerbody reference area.
The centerbody drag coefficlent was determined as a function of Mach numberduring the splnner drag test as depicted in Figure 7-15.
Mach number was established from the ratio of static pressure (measured four
meters upstream of the Prop-Fan rotor) to stagnation pressure. Staticpressure was also measured In the plane of rotation as a backup. The ratio
of static-to-stagnation pressure was correlated with data taken during a
pre-test ca]ibratlon in order to compute the Mach number (see Paragraph 7.2).
The Prandlt-Young correction was applied to the computed Mach number to
compensate for the effects of the tunnel walls and the thrust produced by theProp-Fan.
9.3.2 Data Presentation
The most complete aerodynamic performance data was acquired for the four bladeProp-Fan configuration. Operational problems encountered with the test rig,
while running the two and eight blade configuration, limited the operating
envelope for these configurations. The r|g operational problems wereaddressed earlier In section 8.0. The structural failure of the centerbody
during the test also significantly delayed the test program. Therefore, In
order to expedite the program, the test points were limited to the boundariesand a few Interior points of the operating envelopes for the two and eight
blade configurations.
108
9.3.2 (Continued)
The aerodynamic performance data was nondimenslonalized for analysis
according to the following set of equations.
(POWER COEFFICIENT)
Cp=
BHP (po/p)
5.674 (ND/IOOO)3D 2
(NET THRUST COEFFICIENT)
CTNET =
TNET(Po/p)
340.42 (ND/IOOO)2D 2
(ADVANCE RATIO) V
J = 60 --ND
Where:
BHP = power, KW
TNET = net thrust, N
D = Prop-Fan diameter, m
po/p = density ratio, sea level to ambientN = rotational speed, RPM
V = free stream velocity, m/sec
It should be emphaslzed that the performance data was acquired only during
structural testing where blade angle was continually varied to maintain a
constant power level. Accordingly, the data was plotted as curves of powercoeffIclent and net thrust coefficient versus advance ratio to eIimlnate
blade angle (63/4) as a variable. If desired, thls data can be converted to
efficiency (n) by the relationship:
(EFFICIENCY) CTNeT
q= X,,1Cp
The power and thrust coefficlent data as a function of advance ratio are
presented for the four blade conflguratlon in Figures 9-2, 9-3 and 9-4 andfor the two blade conflguratlon In Figure 9-5. The data was then
cross-plotted to derive the more conventional maps of CTNET versus Cp.Plots of CTNET versus Cp for the two, four and eight blade cases are
presented In Figures 9-6, 9-7, 9-8, 9-9 and 9-I0. The llmlted scope of the
two blade and eight blade data Is apparent. Data taken at 3° inflow angle
rather than at 0° Is presented for the two blade conf!guratlon, because abetter dlstrlbutlon of test points was run at that angle. Examlnatlon of
these plots shows that the net thrust coefficient exhlblts smooth consistent
varlatlons wlth power coefficient and advance ratlo.
109
SR-7L WIND TUNNELPERFORMANCE DATA
• O9
O8
.O7
.o6
.05
.04
CTNET
.03
.O2
.01 | •
/If
CTNET
Cp "" "--"
_Z_ --800KW
O =500KW
C] =2!;0KW
//
//
//
//
//
I P/
//
of/
f
1o
/
4J
0 im
0.6 0.8 1,0 !.2 1.4 1.6
ADVANCE RATIO, J
FIGURE 9-2. Cp AND CTNET VS J, 4 BLADES, MN = .2, _J= 0°
110
SR-7L WIND TUNNELPERFORMANCE DATA
.24_
.ZZ
.Z0
.1
.16
.14
CTNET
.IZ
.10
.O8
.06
.04
.02
1,0
_ .8
.7
Cp
o6
.5
.4
m ,3
.Z
m .I
1.8 2.0 Z.Z 2.4 Z.6 Z.8 3.0 3.2 3.4 3.6
ADVANCE RATIO, J
FIGURE 9-3. CpANDCTNETVS. J, 4BLADES, M N=.5,¢=0 °
111
SR-7L Wi ND TUNN ELPERFORMANCE DATA
.!o
.09
.o8
.o7
CT NET
.06
.05
.O4
.03
.02
.01
.10
.9
.8
.7
Cp
.6
.5
.4
.3
.2
.I
P = 800KW
O P = 500KW
[] P = 300KW
,,I l
f
3.2 3.4 3.6 3.8 4.0 4.2 _4 4.6 4.B
ADVANCE RATIO, J
FIGURE9-4. CpANDCTNETVS. J, 4BLADES, M N=.8,_=0 °
112
SR-7L WIND TUNNELPERFORMANCE DATA
CTNET
1.0 _ .50
.09 _ .45
.08 _ .40
.07 _ .35
.06 _ .30
Cp
.05 _ .25
,04 -- .20
.03 _ .15
.02 _ .IO
.01 _ .05
I I f'_
CTNET / .
Cp m_
/A P'- 800KW
O P'50OKW
[] P=250KW _ jJ
,. .t /
J
_i 1 i ,_
_m" /
1.9 2.0 2. I 2.2 2.3 2.4 Z.5 2.6
ADVANCE RATIO, J
FIGURE 9-5. CP AND CTNET VS J, 2 BLADES, MN -- .5, _=3 °
113
m
SR-7L WIND TUNNELPERFORMANCE DATA
.32
.30
.28
.26
.24
.22
.20
.18
CT NET .16
,14
.12
,I0
.08
.06
.04
,02
/
L_
,10 ,20 .30 .40 .50 .60 .70 .80 .!)0 1.0
Cp
FIGURE 9-6. CTNET VS Cp, CONSTANT J, 4 BLADES, MN = .2.
114
SR-7L WIND TUNNELPERFORMANCE DATA
CTNET
.20
.18
.16
.14
.12
.10
.08
.06
.04
.02
J
/j///,//,'
___lr J = 3.53.0
.J = 2.5
/ IJ=2,0
/,//
• 10 .20 .30 .40 .50 .60 .70 .80 .90 1.0
Cp
FIGURE 9-7. CTNETVS Cp, CONSTANT J, 4 BLADES, MN " .5, _=0 O
115
SR-7L WIND TUNNELPERFORMANCE DATA
,16
.14
CTNET
.12
.1o •
.08
.06 ....
04.02 4.4
J=3. =4.0
.10 .20 .30 .40 .50 .60 .70 .80
Cp
FIGURE 9-8. CTNET VS Cp, CONSTANT J, 4 BLADES, MN = .8, ¢ =0
116
SR-7L WIND TUNNELPERFORMANCE DATA
.IO
.o9
.o8
.07
,o6
CTNET
.05
,04
.O3
.02
.01
J=2.4
/
J/"
/
SJ,,
/
.I .2 ,3 ,4
Cp
FIGURE 9-9. CTNET VS Cp, CONSTANT J, 2 BLADES, M N : .5, _ : 3 °
117
SR-7L WIND TUNNELPERFORMANCE DATA
,26--
.:'4-
.22-
,20-
.18-
.16--
14--
CTNET
.12-
°10_
.08r
.06-
.04.,
.02-
/
J= .91r
J=.8
mMN=.t9
mmm,,m, MN =.50
J=2,64
//
/
J
/7 /
//"1
J
J=2.31
//
/
//
/
.10 .20 ,30 .40
Cp
//
.5o .60
FIGURE 9-10. CTNET VS Cp, CONSTANT J, 8 BLADES, _ = 0°
118
9.3.2 (Continued)
The four blade and eight blade data was cross-plotted again to obtain curvesof net thrust coefficient versus advance ratio for a constant power
coefficient. Comparison of Figures 9-11 and 9-12 shows that for the same
advance ratio and power coefficient, a higher net thrust coefficient isobtained with the four blade configuration than with eight blades.
Therefore, as expected, the four blade configuration, at lower power
loadings, is more efficient than the eight blade Prop-Fan.
Comparlsons of the calculated and experimentally determ|ned performance of
the four and elght blade Prop-Fan configurations are shown In Figures 9-13and 9-14. These calculations were made using a refined lifting l|ne method.
The predlcted and measured performance agree very well for the four blade
conf_guratlon over the entire range of test poInts. Although the performance
of the eight blade Prop-Fan design was underpredicted at Mach 0.5, good
agreement between measurement and prediction was obtalned at Mach numbersof .70 and .73. Moreover, the trends of thrust w_th power are predicted
accurately.
119
SR-7L WIND TUNNELPERFORMANCE DATA
0.28
Cp = 0.10
0.24
0.20
0.16
CT NET
0.12
A 0 Cp = 0.20m
[I_ _Cp = 0.30
Cp = 0.40
ADVANCE RATIO, J
FIGURE 9-! I. CTNET VS J, CONSTANT Cp, 4 BLADES, _ = 0°
120
SR-7L WIND TUNNELPERFORMANCE DATA
0.28
CTNET
0.24
020
o16 _\_0.12
0.08
0.04
0
OCp = 0.2El Cp = 0.24
Cp = 0.29
0 1.0 2.0 3.0 4.0I ADVANCE RATIO, J
5.0
FIGURE 9-12. CTNET VS. J, CONSTANT Cp, 8 BLADES. _J = 0°
121
SR-7L WIND TUNNELPERFORMANCE DATA
CT NET
.12
.11
.IO
.09
.o8
.07
.06
.05
.o4
.O3
.o2
//
/
/ //
/ I
c l /i /
I/ 1
/ //
/ /
I
///
Z_
//
/
//
,//v
f
CALCULATED
PERFORMANCE:
M= 0.2, J= 1.0
M: 0.5, J= 2.0 "" " "" --
M---- 0.5, J= 3.0 _=,,m
TEST DATA:
(_ M= 0.2, J= 0.95
r"IM= 0.5, J= 2.0
/_M= 0.5, J= 3.!
.01
0 .....
0 .05 .10 .15 .20 .25 .30 .35 .40 .45 .50
Cp
FIGURE 9-13. COMPARISON OF MEASURED AND PREDICTED PERFORMANCE, 4 BLADES
122
SR-7L WIND TUNNELPERFORMANCE DATA
CTNET
c iiIII / ,/CALCULATED PERFORMANCE:
M =.5, J = 2.17 i f
M =.5, J = 2.64 t
.11 ' .n M =.70, J = 3.0_
.= 73,J:36 / if" // /
I I I / oif ,/,
.10 /
TEST DATA : ifr
O M =.5, J = 2. I7 ii
.09 ° M =.5, J=2.64 C I /" /A M =.7, J= 3.04 it
I_ M =.73, J= 3.59 / / f.oe / if / i
/ ,/ /
.o, I ,"x' _tI,il
it// I
// I.o5 it
/ /.o4 / , I
I /.03
/
.oz
.oi
.05 ,10 .15 .20 .25 .30 ,35 .40 .45 .50
Cp
FIGURE 9-14. COMPARISON OF MEASURED AND PREDICTED PERFORMANCE, 8 BLADES
123/124
lO.O BLADE SURFACE STEADY PRESSURE MEASUREMENT
I0.I Test Objective
To measure the steady pressure distribution on the surface of the SR-7L
Prop-Fan blade for a range of blade angles, rotational speeds, and simulated
flight Mach numbers.
10.2 Test Procedure
In preparation for blade surface steady pressure testing, the Prop-Fan was
installed on the drive system, as described In section 6.2, and supported as
shown In Figures lO-I and I0-2.
As noted earller in sectlon 2.0, collection of blade surface steady_pressEre_
test data was interrupted during the first tunnel entry In early 1986 and hadto be rescheduled for a second tunnel entry In early 1987. Though the steady
pressure test data collected was quite limited as a result of theinterruptlon, enough Informatlon was obtained to indicate that a revised
pressure tap layout was desirable. A new steady pressure blade incorporating
an improved pressure tap layout was fabricated for use during the second
tunnel entry. It provided a much higher density of surface pressure taps inareas on the blade where the local steady pressures were more sensitive to
changing operating conditions.
The steady pressure measurement blade (S/N 009) was Installed in blade
posltlon 7. For balancing purposes, a counterweight blade (SIN 058) wasInstalled in position number 3, as shown in Figure I0-3. As a precaution,
signals from 2 shank mounted straln gages were monitored throughout the
test. Special contour matching blade stubs were Installed in the remalning 6hub arm bores. The test rig power capabllitles necessitated conducting all
testing using a 2 bladed Prop-Fan conflguratlon, thereby permittlng operation
at power loadlngs per blade corresponding approximately to the take-off andcruise condltlons of the eight blade Prop-Fan design.
The Prop-Fan was operated In the beta control mode during the entire test.In thls mode, Hamilton Standard personnel were able to change the blade pitch
angle during testing by means of an Increase/decrease pitch switch located inthe control room. For a fixed Mach number and a constant power supplied by
the turbines, the Prop-Fan rotatlonal speed was varled by increasing or
decreasing blade pitch angle.
The blade pitch angle (6314) was measured and recorded for each test point.
An electrlcal slgnal proportional to blade angle was provided by means of a
potentlometer mounted on the rotatlng side of the Prop-Fan as Illustrated in
Figure 5-3.
•.,r,_ /.L-._'._--_--_;"_ _-'-----N_ m_
125
BALANCE
GEOMETRICAL
CENTER
CABLEB \ A "lzo R
c
"\11"_'_ " I"-I I I _h'--'-'--"--_
CABLES
(_: 14, 3L:3700
NOTE:
ROD ROD ROD
...... +r-- t'
L::3550 : :: 00
A B L:_,,,o C,
AL-DIME",,ON"INM,,.'I'ETERS(_ = DIAMETERL = LENGTH
126--e'_,:l _kOV3.t-S +-01, 3}
FIGURE 10-2. PROP-FAN STEADY PRESSURE TEST SET-UP
127
ORIGINAL PAGE
BLACK AND WHITE PHOTOGRApHr
#5
/--- COUNTERWEIGHT BLADE
FORWARD TO / SIN 058
AFT VIEW #3 /
I
8
_ STEADY PRESSURE BLADE
S/N OOS
# 7 ,_1 I(BLADE ANGLE MEASUREMENT
SYSTEM LOCATION)
FIGURE 10-3. LAP BLADE INSTALLATIONSTEADY PRESSURE TEST (2 BLADE}
128
10.2 (Continued)
The Prop-Fan rotatlonal speed was measured by use of a IP pickup. The sensor
was triggered by a gear mounted on the test rlg drive shaft. The rotational
speed was averaged over ten revolutions.
The power absorbed by the Prop-Fan was determined by multiplylng the torque
supplied to the Prop-Fan by the rotational speed. Torque supplied to the
Prop-Fan was computed by accounting for the measured frlctional losses in the
test rlg relative to the torque measured by the torquemeter.
Collection of blade surface steady pressure data was accomplished by
utilizing a specially Instrumented SR-7L blade as i11ustrated in Figures 10-4
and 10-5, coupled wlth a Scanlvalve TM pressure measurement system.
Thirteen rows of pressure taps were located on both the face and camber sides
of the steady pressure measurement blade. The pressure tap location andnumbering scheme used for data acquisition and reduction are deplcted In
Figures 10-6, 10-7 and I0-8. This numbering scheme differs from that In the
referenced test plan. The pressure tap channels which span the blade were
fabricated by bonding a thln plastic skin to a channellzed adhesive layer.Each channel was connected to a tube embedded in the root of the blade which
led out to the blade shank. The tubes were connected to the Scanivalve
mounted on the dome cap of the Prop-Fan, protruding through the nose of the
spinner. One Scanlvalve channel was provided for each tube. The stationary
portion of the Scanlvalve contalned a pressure transducer that could be
scanned by remote command to monitor one channel at a time. This arrangementallowed pressure measurements to be made at only one radial station per run.
Pressure measurements were made by masking off all the rows of pressure taps
except at the section of interest. The Scanlvalve was then cycled through
all channels to record the pressures at one radial station. Thirteen runs
were required at each Prop-Fan operating point to obtain a complete pressuremap for the blade surface at the operating point.
The Scanlvalve was enclosed in an aerodynamic falrlng to ma}ntafn a wellbehaved inflow to the Prop-Fan. The umbilical, which connected the
Scanlvalve to the control and monitoring equipment outside the tunnel, was
enclosed in a conduit with an airfoil shaped cross-section. This alsominimized disturbance of the flow.
Table 10-I lists the operating conditions that were run _r_m_} the test along
wlth tolerances showing the maximum variation In the parameters a11owed whentestlng at different radial stations. In general, the p_=ce<Bure for setting
a speclflc test condition was to set Mach number and them a{IL1ustthe rotor
speed and blade angle, to obtain the desired power coeff_¢_emt am<l advanceratio.
129
ORIGINAL PAGE
BL&CK AND WHITE PHOTOGRAPH
FIGURE 10-4. LAP STEADY PRESSURE BLADE {CAMBER SIDE}
130 " AG_
OF __UA_rIp_-_
FIGURE 10-5. LAP STEADY PRESSURE BLADE (FACE SIDE)
131
ORIGtNALBL.._CK .At,JD WHITE
PACE
Pl-i !.]h__ _,_ p _,
20'2O
20
20
20
Z0
Z0
ROW ! 3
)W IZ
I]
OW 10
R o_w p.._ _I
ROW 8
oPRESSURE PORT LOCATIONS
.NUMBERING SEQUENCE
ROW 7
ROW 6I
20
2O
ROW 4
ROW 3
20
ROW Z
CAMBER SIDE
FIGURE 10-6. STEADY PRESSURE BLAJ[_
132
• PRESSURE PORT LOCATIONS
eNUMBERING SEQUENCE ROW I 3
ROW 12,
ROW 1 I
ROW 1
ROW
ZI
p3636
36
ROW 7
ROW 6
ROW 5 36
ROW 4
ROW 3
36
ROW 2
ROW !
FACE SIDE
FIGURE 10-7. STEADY PRESSURE BLADE
133
l,J- !,20
I,J= I, I
CAMBER SIDE
(BLADE OUTLINE MIRRORIMAGE OF FACE SIDE)
56.0 t 142.24
52.0 132.08
48.0 "121.92
44.0 1 1.76
40.0 01.60
36.0
16.0 '40.64
'30.48
20.32
4.0 10.16
0,0' ).00
INCHES CENTIMETERS
FACE SIDE
I, J = !, 36
NOTE: THE VERTICAL SCALE REPRESENTS THE DISTANCEALONG THE BLADE PITCH CHANGE AXIS FROM
THE PROPFAN CENTER LINE.
FIGURE 10-8. ARRANGEMENT OF THE PRESSURE TAPS ON THE BLADE SURFACE
-- PROJECTED VIEW
134
TABLE 10-I. OPERATING CONDITIONS FOR BLADE STEADY PRESSURE TESTING
(2 BLADE LAP PROPELLER)
M J /3 CpCondition Mach Advance Ratio Blade Angle Power Coefficient
No. No. +J2 _+}00 ' __2
NRPM
1 .01 .08 13.80' .0792 .02 .14 15.70 o .0933 .02 .15 18.780 .1524 .03 .18 21.600 .2046 .20 .88 25.650 .0986 .20 .88 30.40 ° .2517 .50 3.065 57.510 .6498 .50 3.055 54.95' .3609 .50 3.063 50.860 .108
10" .60 3.066 54.980 .22611* .70 3.055 55.00 ° .22912" .78 3.07 54.97 o .22313" .78 3.20 54.98' .112
12001200
_ L20o.1200
16651651
1186119011851436168518401782
* RADIAL STATIONS 2,4 AND I0 WERE NOTRUN AT THIS CONDITION.
135
10.2 (Continued)
After establishing the operating condition, the basic wind tunnel and
Prop-Fan parameters (tunnel static temperature and pressure, Math number and
rotor speed) were logged into the microcomputer, a record number was assignedfor filing purposes, and the scanivalve was activated. The scanivalve then
ran through a calibration sequence followed by the pressure data scan. The
data were then plotted in preliminary form and reviewed. If the data
contained questionable features, a second scan was performed or hand scanning
of individual suspicious points was made.
Test points 1 through 4 in Table lO-I approximate static conditions of
increasing power and were selected to provide information on leading edge
vortex flow (there was no applied tunnel flow for these points, althoughthere was some Prop-Fan induced flow). Points 5 and 6 were seledte¢ to - -
Investigate take-off conditions: points 7 through g covered a wide range of
power loading conditions at 0.50 Mach number and bracket through the design
cruise power loading. Points lO through 13 were selected to Investigate
transonic flow characteristics and were all at relatively low power due torig limitations. For these points, power coefficient and advance ratio wereheld constant while Mach number was varied.
I0.3 Discussion and Results
IO.3.l Data Reduction
lO.3.l.l Data Format
The pressure data were reduced to coefficient form and plotted during the
test using a microcomputer system to provide the test personnel with a basis
for Judging the quality of the data following each scan. This preliminarydata reduction included an approximate correction (described below) for the
centrifugal pumping In the tubes and channels that led from the scanivalve to
the pressure taps. The pressure coefficient formula that was used is:
Cp I
Pc - Po
0.sp(v +
Nhere:
Pc - Corrected Blade Surface Pressure
Po - Tunnel Static Pressure
p , Air Density
Vo - Tunnel VelocityVtm Tangential Velocity at Pressure Tap Radius
136
........... 7
10.3.1.1. (Continued)
For each radial station the mid-chord radius (for a prescribed blade angle of
50.0 degrees) was used in determining the approxlmate centrifugal correction
and the tangential velocity. It should be noted, however, that the radii ofthe pressure taps at each station vary a|ong the blade chord, and furthermore,
the radius of any given tap varies with b]ade angle. In addition, the va]ues
describing the chordwise distribution of the pressure taps, which were stored
in the microcomputer, were nominal rather than measured values, and did not
represent the precise distributions. The final data reduction at HamiltonStandard will account rigorously for these effects, and will be presented in
a separate NASA Contractor Report (Reference 24).
10.3.1.2 Discussion of Problems
Before revlew|ng the data, it is necessary to address three basic test
problems, each of which had some effect on the data presented in this
report. The problems are as follows in decreasing order of severity:
l , Particle impacts resulting in failure of the tape seal over rows of
dormant pressure taps or resulting in venting of individual channels.
, Crack formation In the plastic skin which seals the channels on theblade surface.
3. Reference pressure transients which interrupted the scan sequence.
Partlcle impacts on the steady pressure blade occurred on several occasions
durlng the test. In most cases the impacts resulted In little or no damage,
however, four cases of damage to the tape which sealed the dormant rows of
pressure taps occurred and two cases of channel venting occurred due todirect partlc_e Implngement on a specific channel. The data affected by tape
seal fallure were llmlted to radial stations 3, 6, 11 and 13 at Math numbers
generally greater than 0.20. Channel venting due to direct particle Impact
resulted on channels 26 and 29 and was repaired followlng the runs in whichit occurred. The data that were compromised by Impact events will be
e11minated from the final data package.
Prior to the second test run (radial station 13), cracks were found in the
camber side plastic skin layer, between radlaI stations 6 and 8, and 9 and I0,
on the blade. The cracks, which are believed to have been catalyzed by theappllcatlon of a cleaning solvent that "shocked" and embrittIed the plastic,
were not present following the first run (radial station 5); they were found
Immediately after applying the solvent. Continued blade checks showed high
leak rates oo_channels 2 through 8. To correct the leak problem a series oftape strlpswas applied to the cracked areas, sealing them off and
establishing acceptable leak rates. Testing continued with this conflguration
and leak rates were monitored following every other run. The fix was found
_.,_to_e satisfactory on all channels, although channel 2 required further repair_" as the test progressed. In general this problem is not considered to have-, compromised the data.
_- DI::_3G_AI_ PAGE rg
137 DE POOR QUAI/T¥
10.3.1.2 (Continued)
Reference pressure transients occurred during several data scans.
Fortunately, this phenomenon was quite evident in the on-line pressurecoefficient p]ots, so scans that were affected were either rerun entirely or
hand scanned to pick up the affected points. Therefore this problem isconsidered to be of minimal consequence to the test data.
10.3.2 Discussion of Result
The entire spectrum of data Is summarized in Figures 10-9 through 10-21; each
figure shows all of the measured pressure distributions for a specific testcondition, and is in the form of the on-line data reduction. These figures
are presented in order of increasing Mach number. For Mach numbers where twoor more power settings were run, the figures are sub-ordered by Increasing _-
power. Radial stations 2, 4, and lO are left blank for Mach numbers greaterthan 0.50 because no data were collected due to limited test time. Radial
stations with data polnts which may be eliminated or replaced with hand
logged values, for the final data package, are marked with an asterisk.
Figures 10-9 through 10-12 show the pressure distributions for the nominalstatic operating condition. The pressure loading is seen to increase with
increasing applied power, as expected, and the presence of leading edgevortex flow Is suggested in Figures lO-ll and lO-12 by the negative pressure
hump that spans along the camber side leading edge; the vortex then appearsto sweep across the chord In between the 90 and 95 percent radius, resulting
in very high loading in that region.
The data at these static operatlng conditions contain some inconsistencies.
For example, when running radial station 6 the data was found to diverge from
the trends at the neighboring stations. The cause of this inconsistency isnot certain, however, It is noted that the actual Mach number was not zero
(due to the Prop-Fan induced flow) and was variable during the "static" runs(Mach No. varied from 0.02 to 0.04). Though the reason for the Mach number
variation Is not known for sure, it was revealed during a daily tunnel
inspection that a portion of the tunnel flow straightening honeycomb had beenblown out. Subsequent running with the Prop-Fan driving the tunnel resulted
in somewhat higher Mach numbers. To investigate this problem further radialstation 7 was rerun for conditions 2 and 4. Comparisons of the pressure
distributions or these runs are given in Figures I0-22 and 10-23. It is
clear from this comparison that a repeatability problem existed for some of
the static point data.
Figures 10-13 through 10-21 and Figure 10-24 show the pressure distributions
for the remaining operating conditions. The data at these conditions show
good repeatability. For example, Figure I0-24 shows radlal station 8 whichwas scanned twice at 0.20 Mach number, first on the climb up to 0.78 Mach
number (in accordance with the test plan sequence), and then on the decent,
approaching shutdown. The results are seen to be nearly Identical.
138
10.3.2 (Continued)
Some of the aerodynamic effects that are apparent in Figures lO-13 through
10-21 are leading edge vortex loading at 0.20 Mach number for the take-off
case (Figure I0-14), inverted leading edge pressure distributions for the low
power cases at high Mach numbers (Figures 10-15, I0-18, 10-19, I0-20 and
lO-21) and evidence of trailing edge shock waves at the outboard stations at
high Mach numbers (evident by the trailing edge pressure jump in
Figures I0-19, I0-20 and I0-21. The inverted leading edge pressuredistributions, as noted above, are typical for cambered airfoil sections
operating at Incidence below the design angle of attack value.
139
RADIAL STATION I
0.591
2, 0.474
RADIAL STATION 1 .
r/R z 0.341
1,000 0.2 0.4 0.6 0.8 I 0
BLADE CHORD, X/C
FIGURE 10-9, LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
¢
140
.... • i iNOTES: I. TOLERANCES SHOWN DENOTE MAXIMUM 0] !__|2
PARAMETER VARIATION FROM ONE RADIAL / /
S,TATION TO THE NEXT. -05. "_Z. -'PLOTS CONTAIN DATA POINTS WHICH MAY _ I . .
BE REPLACED OR ELIMINATED. 01 ! I r I ,J'
-0.5
:. •o-z o: o _ 9 X CAMBER SIDE"
,,i _-_-_ to rACES'DE i-1.0' 0 L I ,_'
-05;
-,o. o_
-0.1_
-t,O _ 13, 0.9911
5 12, 0.960
-2.0 -0]--Q t I , 0.975
0.964
-1.5 0.938
0.903-1.0 0
0.861
-0.5 0.609
5, 0.747
0.482
i2.0 0_2
-I .5 0.5_
-I .0
-:'.5 -0.5
L_ -Z,O 0
_" -1 .S
¢_ O.S
o.5 RADIAL STATION 1
1,00 0.2 0.4 0.4 0.0 1.0
BLADE CHORD. X/C
_$, 0.591
RADIAL STATION I.
r/R = 0.341
FIGURE 10-10. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
141
0.59 1
RADIAL STATION I
02 04 04 0,° 1,0I_AD[ CHORD, XIC
RADIAL STATION I .
r/R -- 0.341
FIGURE 10-1 1. LAP SR-TL STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
142
NOTES:
13
FIGURE 10-12. LAP SR-7L STEADY PRESSURE DISTRIBUTION {ON-LINE DATA)
l_t3
NOTES:
TEST CONDITION NO. 5
M N _ O.ZO= ZS.05 ° + 1.00 °
RPM - 1665 + I 0
J =, 0.88 + 0.0Z
Cp : 0.098 + 0.0Z
TOLERANCES SHOWN DENOTE MAXIMUM
PARAMETER VARIATION FROM ONE RADIAL
STATION TO THE NEXT.
* PLOTS CONTAIN DATA POINT_J WHICH MAY
BE REPLACED OR ELIMINATED,
CAMBER._IDE ....
FACE SI DE
2, 0,474
RADIAL STATION I,
r/R _ 0.341
FIGURE 10-1 3. LAP SR-TL STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
144
0. 0.2 0.4 0.6 0.11 1.0
BLADE CHORD. X/C
FIGURE 10-14. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
145
NOTES:
TEST CONDITION NO. 9
X HAMBER SIDE JO FACE SIDE
3, 0.59 I
2, 0.474
RADIAL STATION I,
r/R : 0.341
FIGURE 10-15. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
146
NOTES:
T£Er CONDITION NO. 8
M H = 0.50
" 54.95" _- 1.00"
RPM "IIg0±10
J = 3.055 + 0.0Z
Cp = 0.360 +_0.0Z
I , TOL£RANCE;S SHOWN DENOTE MAXIMUM
PARAM£TIER VARIATION FROM ON£ RADIAL
STATION TO THE NEXT.
2. • PLOTS CONTAIN DATA POINTS WHICH MAY
BE R£PLACIED OR £LIMINAT£D.
CAMBER SIDE
-I ,0_
-0.$ •
0,$
-0.5_o _2
0._ _
-0.5_ | ,I,
o,i_ |
-o.
10 ixo.sJ O
ol .O
-0.5
0.5
0.
13*
(CAMBER SIDE iFACE SI'D£ " -- -
op==,==:_-.--.--._ 6 __-- I,. o.,_
-05 °'s_ _-'*_-'-----_ _-.._ Io,o.J,.4
,--._ _7, 0,SEt
-0, / /o__-'.. I I __-::--.. 4*/ J__
-05 0'5= 3
-0.s] o,__ _ =,0,474
05_ ' RADIAL STATION _,u -0.'=t " _ ___'_ r/R - 0.34I
_ os___j" STATION I_
• 0 02 0.4 0,8 0,8 1,0
BLADE CHORD, X/C
FIGURE 10-16. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
147
TEST CONDITION NO. 7 "I'0 l
MN - o.3o " t r_'T- F-T-_- $?,51° + 1,00 ° 1 I I I I J
J - _I,0S_ + 0,0Z -1.01 t _ NV/ "
Cp - 0.649 + 0,02 |
NOTES: I, TOLERANCES SHOWN OI"NOTEMAXIMUM _0,5t _'_,._2"
PARAMETER VARIATION FROM ONE RADIAL ] J l l J J,
STATION TO THE NEXT. 01 _ - _ ;,
2, * PLOTS CONTAIN DATA POINTS WHICH MAY " }
ro:::::o.i
°2.0 .0.5
FIGURE 10-! 7. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
148
TEST CONDITION NO. 10
M N ', 0.60j_ = S4,98 o +_. 1.00 °
RPM = 14:36 + 10
J = 3.066 + 0.02
Cp = 0.226 + 0.0Z
NOTES: I. TOLERANCES SHOWN DENOTE MAXIMUM
PARAMETER VARIATION PROM ONE RADIAL
STATION TO THE NEXT.
Z. " PLOTS CONTAIN DATA POINTS WHICH MAY
BE REPLACED OR ELIMINATED.
-o.5.oI_______ !*0.5
-o_ 11_
0.5
lO
FACE SIDE
0.082
-- 3.0.59 1
FIGURE 10-1 8.
RADIAL
STATION I
RADIAL STATION I,
r/R = 0.341
LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
149
NOTES: 1.
TEST CONDITION NO. I I
0 _
0.5
M N " 0.70
- 55.00 _ + 1.00 °
RPM - 1685+ IO 1| _
J = 3.055 +_ 0.02
Cp = 0.229 +_ 0,02 o
TOLERANCES SHOWN DENOTE MAXIMUM
PARAMETER VARIATION FROM ONE RADIAL
STATION TO THE NEXT. 10
2. * PLOTS CONTAIN DATA POINTS WHICH MAY
BE REPLACED OR ELIMINATED.
-0.5
o:I Lx
13"
CAMBER SIDE
FACE SiDE
0. S9 !
2.0.474
FIGURE 10-19. LAP SR-TL STEADY PRESSURE DISTRIBUTION {ON-LINE DATA)
150
NOTES:
TEST CONDITION NO, 13
M N = 0.78
]3 = 54.98" + 1.00
RPM : 1782 + 10
J = 3.20 + 0.02
Cp : 0.1 I Z +_ 0.02
I. TOLERANCES SHOWN DENOTE MAXIMUM
PARAMETER VARIATION FROM ONE RADIAL
STATION TO THE NEXT.
2. * PLOTS CONTAIN DATA POINTS WHICH MAY
BE REPLACED OR ELIMINATED.
-1.O-
wh: -0"5_O_Z
o
a. _ 0.5U
1.00 0.2 0.4 0.6 0.8 1,0
BLADE CHORD. X/C
3
0.5
lO
CAMBER SIDE D "
FACE SIDE
0,996
0.986
1 I, 0.975
0,964
0,938
0,903
0.841
6, 0.009
0.747
0.602
2, 0.474
RADIAL STATION I,
r/R -- 0.341
FIGURE 10-20. LAP SR-7L STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
151
TEST CONDITION NO. 12
N - 0.78- 54.07 ° + 1.00"
RPM : 1040 + 10
J - 3,07 _+ 0.0Z
Cp = 0.223 _+.0.02
TOLERANCES SHOWN DENOTE MAXIMUM
PARAMETER VARIATION FROM ONE RADIAL
STATION TO THE NEXT,
* PLOTS CONTAIN DATA POINTS WHICH MAY
BE REPLACED OR ELIMINATED,
4,]
oi] 13"
-0.5,
o_ I1"
-0.5
0.i 1-_
-0,5-
0.$"
10
9 •
CAMBER 51D£ jFACE SIDE
13, 0.986
12. 0.986
I !. 0.975
0.904
0.038
0.803
O. 801
0.809
0.747
0.082
0.591
2, 0.474
RADIAL STATION 1,
r/R = 0.341
FIGURE 10-21. LAP SR-TL STEADY PRESSURE DISTRIBUTION (ON-LINE DATA)
152
-2.5
-2.0
-1.5
z
-_ -l.o
_m "0. 5
0.0
0.5
1.0
LAP (SR-7) PRESSUREDISTRIBUTION
Radia] Sta.: 7 Record no.: 114
Beta (de9): 15.3 03-10-1987 2h48:01
Hach No. : _02
Speed (rpm): 1209
x comb_ side
io ?ace i_ide
-2.5
-2.0
-I,5
d_
wZ
_-I,0
°0.5
kd
0.0
0.5
Io0
LAP (SR-7) PRESSUREDISTRIBUTION
Radial St_ : 7 Record no. : 204
Beta (deg): 14.3 03-13-ig87 20:55.30
Moch No, = 0.02
Speed (rpm): IIQ3
x combei side
o-Fac_ :ide. _
........ 1 ....
0.0 0.2 0.4 0.6 o.e 1.0 0.0 0.2 0.4 0.6
BLADE CHORD. x/c BLAOE CHORD. x/c
o.e 1.0
FIGURE 10-22. REPEATABILITY PROBLEM FOR STATIC LOW POWER POINT
153
-2.5
-Z.O
-].5
&
-z.o
U
ua -0.5
0.0
0.5
hO
LAP (SR-7) PRESSURE OISTRIBUTION
Radial Sta.: 7 R_cord no.: [03
Beta (de9): 22.4 03-i0-xg87 18:26:06
Mach No. : 0.02
Speed(rpm): |203
i i dx comber sl e
Io Face bide
E
f
0.0 0.2 0.4 0.6 0.8 1.0
BLAOE CHORO. xlc
-2.$
-2.0
-1.5
zW
-hn
N(..I
_-O.g
LAP (SIR-7)PRESSURE OISTRIBUTION
Radial 5ta. : 7 Record no. : 205
Beta (de9): 21.4 03-13-igB7 21:04:30
NOch No. _ O.04
Speed (rpm): 120!
i
Ix combe- sideo ?ace ;ide
0.0
0.5
1.0
0.0 0.2 0.4 0.6 0.8
BI_AOIZC.IO_. xlc
1.0
FIGURE 10-23. REPEATABILITY PROBLEM FOR STATIC HIGH POWER POINT
154
-Z.5
-Z.0
LAP (SR-7) PRESSUREDISTRIBUTION
Rodiol 5to.: B Record no.: 172
Beto (de9): _6. e 03-[2m19_7 20: 2_: 50
Hoch No. : 0.20
Speed (rpm): 1670
x combe side
o Foce _ide
-65
-l.OtJ
-0.5
0`0
0.5
l°O
0`0
-& 5
-Z.0
-1.5
_-].0tJ
-0.5
0`0
0.5
1.0
LAP (SR-7)PRESSURE DISTRIBUTION
Rodio] 5to.: B Record no. : [B|
Beto (deg): 25.0 03-12-1887 21:5g: 4B
Moch No. : 0.20
Speed (rpm): 1701
x combe side
- != r_:e _Q-_
I
0.2 0,4 0,6 0`e hO 0`0 0,2 0.4 O.e 0`8
BLAOE CHORD. x/c 8LAOE CHORD, x/c
FIGURE 10-24. ILLUSTRATION OF REPEATABILITY AT 0.20 MACH NUMBER
155/z56
II.0 BLADE SURFACE UNSTEADY PRESSURE MEASUREMENT
11.1 Test Objectives
11.1.1 Measure the unsteady pressure distribution on the surface of the
SR-7L Prop-Fan blade for a range of blade angles, rotational speeds, and
simulated flight Mach numbers for axial and angular inflow conditions.
11.I.2 Evaluate the effect of a wake in the propeller inflow on the unsteady
pressure distributlon on the surface of a blade.
11.2 Test Procedure
In preparatlon for blade surface unsteady pressure testing, the Prop-Fan wasinstalled on the drive system as described in section 6.2, and supported as._
shown in Figures ll-l and ll-2. Table ll-I shows the conditions that weretested.
The unsteady pressure measurement blade (S/N 054) was Installed in blade
position 3. For balancing purposes, the steady pressure measurement blade
(SIN 009) was Installed In position number 7, as shown In Figure ll-3. As a
precaution, signals from 2 strain gages located on the shank of the unsteadypressure blade were monitored throughout the test. Specla] contour matchingblade stubs were installed in the remaining 6 hub arm bores. The test rlg
power capabilities and tunnel time constraints necessltated conducting all
testing using a 2 bladed Prop-Fan configuration, thereby permitting operation
at power Ioadlngs per blade corresponding approximately to the take-off andcruise conditions of the eight blade Prop-Fan design.
The Prop-Fan was operated in the beta control mode during the entire test.In this mode, Hamilton Standard personnel were able to change the blade pitch
angle during testing by means of an Increase/decrease pitch switch located inthe control room. For a fixed Mach number and a constant power supplied by
the turbines, the Prop-Fan rotational speed was varied by increasing or
decreasing blade pitch angle.
The blade pitch angle (B3/4) was measured and recorded for each test point.An electrical signal proportional to blade angle was provided by means of a
potentlometer mounted on the rotating side of the Prop-Fan as illustrated in
Figure 5-3.
The Prop-Fan rotational speed was measured by use of a IP pickup. The sensor
was triggered by a gear mounted on the test rig drive shaft. The rotational
speed was averaged over ten revolutions.
The power absorbed by the Prop-Fan was determined by muItlplying the torque
supplled to the Prop-Fan by the rotational speed. Torque supplied to the
Prop-Fan was computed by accounting for the measured frlctlonal losses in the
test rig relative to the torque measured by the torquemeter.
157PRECEDING PAGE BI.NNIKNOT FILib_.O
I-0Z
158
OF_ Y
OR,IGINAI. PAGE
BLACI'( AND ',,llIHITE P'r,KTrtY_APH
FIGURE1 !-2. PROP-FAN UNSTEADY PRESSURE _ SET-UP
159
TABLE IT-l.OPERATING CONDITIONS FOR BLADE UNSTEADY PRESSURE TEST
(2 BLADE LAP PROPELLER)
M J _ Cp NCondition Mach Advance Ratio Blade Angle Power Coefficient RPM
No. No. +.02 +1.00 ° +.02 ±10
2 .02 .14 15.700 .093 12003 .02 .15 18.78 o .152 12004 .03 .18 21.60 ° .204 12005 .20 .88 25.65 o .098 16655A .20 .88 27.190 .15 16845B .20 .88 29.50 ° .20 16846 .20 .88 30.40 ° .251 16517 .50 3.065 57.51 ° .649 11868 .50 3.055 54.950 .360 11909 .50 3.063 50.860 .108 1185
10 .60 3.066 54.980 ,226 143611 .70 3.055 55.00 ° .229 tB85
160
#5
F UNSTEADY PRESSURE BLADE
FORWARD TO //S/N 054
AFT VIEW #3 /
8
STEADY PRESSURE BLADE
SIN 009
_7_-" I{BLADE ANGLE MEASUREMENTSYSTEM LOCATION}
FIGURE 11-3. LAP BLADE INSTALLATIONUNSTEADY PRESSURE TEST (2 BLADE)
161
11.2 (Continued)
Mach number was established from the ratlo of static pressure, measured four
meters upstream of the Prop-Fan rotor, to stagnation pressure. Static
pressure was also measured in the plane of rotation as a backup. The ratio
of static to stagnation pressure was correlated wlth data taken during apre-test calibration in order to compute the Mach number. The Prandlt-Young
correction was applied to the computed Mach number to compensate for the
effects of the tunnel walls and the thrust produced by the Prop-Fan.
Collection of blade surface unsteady pressure data was accomplished byutilizing a specially instrumented SR-7L blade as illustrated in Figures II-4
and 11-5, coupled with an FM multiplex data acquisition system. Twenty-six
high frequency response pressure transducers were installed in two rows onthe face and camber sides of the unsteady pressure blade. The p÷esgur_ ---
transducer location and the numbering scheme used for data acquisition and
reductlon are depicted in Figures 11-6 and 11-7.
The pressure transducers were mounted flush with the blade surface as
depicted In Figure II-8. The signal and excitation wires from each
transducer were connected to slgnal conditioning electron|cs located In the
cuff of the blade. The slgnal wires also passed through attenuatingresistors mounted on the blade root. The function of the attenuating
resistors was to establish the gain for the pressure slgnals.
The unsteady pressure signals were transmitted from the rotating to thestationary field through the FM multiplex data acquisition system provided
for the SR-TL Prop-Fan. The signals were monitored on a four-channel
oscilloscope and recorded on a 14-track IRIG tape recorder.
The frequency response of the system was DC to 1000 Hz. Prior to the High
Speed Wind Tunnel Test an evaluation program was conducted to determine the
sensitivity of the transducers to temperature, strain, vibration, and
centrifugal 1oadlng. The results of this test program indicated a maximum 2_
of full scale error due to temperature in the range from 0 to 130°F and amaximum .92% of full scale error due to all other factors.
Generation of the wake In the Prop-Fan inflow was accompllshed by erecting a
vertlcal steel cylinder upstream of the rotor. The cyllnder was lOOmm (3.93inches) In diameter and was located such that its centerllne intersected the
Prop-Fan axis of rotation at a distance of 1.372m (54.02 inches) upstream of
the rotor plane. The wake generated by the cylinder was Intended to create a
twice per revolution (2P) disturbance for the Instrumented blade to pass
through. Figure ll-9 shows the Prop-Fan with the cylinder _n place.
162
• i
ii
ii
ii
!I
Ii
m
I
C20977
FIGURE 1 1-4. LAP UNSTEADY PRESSURE BLADE(CAMBER SIDE)
163
ORICINAL PAGE
8L,_C.K .A,'qD'_-_,ITEPI.K)T'b_GR,kYi,I
ORIGINALPAGEAND WHITE PHOTOGRAPH
~
k _
FIGURE 1 1-5. LAP UNSTEADY PRESSURE BLADE (FACE SIDE)
164
7,
FIGURE ! 1- 6 UNSTEADY PRESSURE BLADE, FACE SIDETRANSDUCER LOCATIONS
165
PT14C_ T18CPT16C
I
' I IPT15C PT19CPT17C
PT23C
PT21C I II
' I ' IPT24CPT20CPT22C PT26C
FIGURE 1 1- 7 UNSTEADY PRESSURE BLADE, CAMBER SIDETRANSDUCER LOCATIONS
166
\\
\\
\
PETG (GLYCOL MODIFIEDPOLYETHYLENE TEREPHTHALATE
SHEET} 0.030 THK-.j,
J_ L_._
SHELL PLIES
\
\
\
]t/
//
/
p TRANSDUCER/ (FLUSH MOUNTED}
,/J
J
FIGURE ! 1- 8 LAP SR-7L UNSTEADY PRESSURE BLADE TRANSDUCER MOUNTING
167
t.V" <_,--_"_,-V;..I.?_.TY
WAKE GENERATOR
FIGURE 1 1-9. UNSTEADY PRESSURE TEST SET-UP WITH WAKE GENERATOR
168
ORIGIN'A/ PA_['
6LACK AND _'HITE PHO'IOGRAPId
=
11.2 (Continued)
The complete range of Prop-Fan operating conditions that were tested is givenin Table ll-I. Because of time limitations, the range of conditions tested
was reduced for the zero degree inflow cases with and without the cylinder.
In addlt_on, transducer wire problems resulted in intermittent and
nonexistent signals on several sensors. The resulting series of testcondltions for which signals were recorded is given for each transducer in
Table ll-II.
11.3 Discussion and Results
To provide an illustration of the data collected during the test, examples
are given in Figure ll-lO of the periodic variations In pressure (withcorresponding frequency spectra) which were measured on the camber side ok -the blade at the 90% radius, 56% chord point (pressure transducer number
PTI6C). Here the three inflow cases are compared for the Prop-Fan operating
condltion defined below.
Mach number, MN - 0.20
Advance Ratio, 3 = 0.883
Power Coefficient, Cp = 0.250
Blade Angle, 8 = 32°
This operating point Is representative of the Prop-Fan take off condition.
The pressure versus time plots at the left in Figure II-I0 were obtained from
a sIgnal enhancing waveform analyzer. Sampling was Inltiated by the recordedonce per revolution plp signal and waveforms from I024 revolutions were
averaged. Thus the repetitive portion of the pressure waveform is enhanced
and the random part Is suppressed. The spectra, shown at the right In
Figure ll-lO, were obtained vla digital Fourier transform analysis.Successive time slices were transformed and averaged for 4.8 seconds. Each
spectrum conta|ns 400 frequency points spaced linearly from 0 to 500 Hertz.
Prel|minary Interpretation of Figure ll-lO is as follows:
In interpreting the waveforms, a transducer can be considered to be scanningthe inlet flow as it rotates. Since the blade position Is known as a function
of time, the time axls can be converted to angular position. The six and
twelve o'clock positions are indicated on the bottom trace. For the trace at
the top of Figure ll-lO, representing clean inflow, the signal level shouldbe low, corresponding to a low distortion level. However, a small slnusoldal
component can be seen In the waveform and spectrum that must be caused by aresidual flow angularity In the tunnel. This can be consldered a backgroundlevel and must be subtracted from the data for the 3° angular inflow and for
the cylinder wakes.
169
F-.
:+0
+,
+<
li:l--
hoI,-,,
<"I-(,,3
+i
li,i,.,.Illi<I--
<7
O__i u
...i Z
:+,-+
_zl-o+
I+,1,11:3.-12:;
I<_
I++!-
a _m-_ z<_ 0
WI-_ I"
hi
<">a
m" o <
l_-"tl._ . _,_ _ a
": _i
- _+_
170 +7<©C+p++++++r+,+,+-
N N N
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171
11.3 (Continued)
In the data for the 3° angular inflow, the angle of attack seen by theinstrumented blade should be nearly a pure sine wave at the once per
revolution (IP) frequency. Simplistic analys_s wou}d _ndlcate that the bladepressure response should also be sinusoidal. The waveform and spectrum showthat this is far from true. Figure 11-11 illustrates the terms "advancing"and "retreating" for angular inflow. Further evaluation of the data at otherpositions on the blade is required to identify the source of thisnon-slnusoidal behavior.
For the data with the cylinder wake, the blade pressure should respond with a
pulse each time the blade passes through a wake at the top and bottom of therevolution. This behavior Is observed in the bottom trace, but the pulse
magnitudes are surprlsingly different at the top and bottom positions_ ....
Another Interesting feature of the data for cylinder wakes is the oscillating
response after the wake pulse.
Slnusoldal response was observed on the pressure (face) side of the blade In
all cases examined for angular inflow conditions.
Sinusoldal response was also observed on the suction (camber) side of the
blade under low 1oadlng conditions. However, under high loading conditions,
non-slnusoldal behavior is present. The non-slnusoldal response appears to
be a result of leading edge and tip vortices which may be distorting theresponse. Another possibility Is the formation and breakdown of the vortices
as the angular Inflow or wake inflow modulates the angle of attack.
An analysis of the blade unsteady surface pressure data will be presented ina separate NASA Contractor Report (Reference 25).
172
ZRETREATING BLADE
INFLOW
i, IL
% _ w
ADVANCING BLADE
FIGURE 11-1f. ILLUSTRATION OF THE TERMS "ADVANCING" AND "RETREATING"
FOR ANGULAR INFLOW
173/174
12.0 CONCLUSIONS
The High Speed Mind Tunnel Test has provided an extensive evaluatlon of the
operating characteristics of the SR-7L Large-Scale Advanced Prop-Fan. Allthe test objectives, set forth in the Plan of Test 267X-135 Rev. D, regarding
acquisition of data were accomplished and I02 hours of operating experience
were attained. No problems were uncovered that would have been considered an
impedance to the planned follow-on PTA Flight Test. ONERA drive system
constraints and testing problems precluded running all desired test points,
however, those polnts successfully run did provide a wealth of aerodynamic
performance, structural dynamic, steady and unsteady blade pressureInformatlon. Further areas of investigatlon are indicated that should
ultimately result in highly accurate aerodynamic design and performance
prediction methods. The conclusions and recommendations derlved from each
phase of the High Speed Wind Tunnel Test are presented in the follow_ng -- -sections. Two separate low number NASA Contractor Reports will be published
providing a more detailed evaluation of the blade surface steady and unsteady
pressure tests.
12.1 Blade Structural Dynamic Evaluation
The SR-TL Prop-Fan was found to be free of high speed blade flutter over the
entire operatlng envelope tested. All the measured blade surface and bladeshank strains were below the a]lowables set prior to testing. These allowable
levels were set to avoid accumulation of fatigue damage to the blades,
therefore, no fatlgue damage to the blades was incurred.
Reasonable correlation was found between the measured and analytically
predicted IP blade bending strains for the SR-7L blade. Results confirmed
that IP strain peaks inboard on the blade and lessens near the tip. Also,
blade strains were found to Increase with power and Mach number (all othervariables held constant).
The Interpretation of IRP (infrequently repeating peak) strain values was
made difficult due to a significant amount of high frequency (>25P) noise in
the signaTs. In generaI, this type noise Is easfTy filtered out, however,this could not be done for the zero inflow angle data due to the low signal
response levels. The angular inflow (3°) strains were determined fromspectral analysis, the amplitudes of which were not affected by "noise"
outside the frequency range of interest.
Signiflcant 2P blade vlbratory response was also measured for the 3° angularInflow case. The amplitude of the 2P response averaged approximately 25% of
the IP component response. It is concluded that at least a portion of the
excitation force driving the 2P response was generated by the test rig drivesystem. Twice per revolution vibration is characteristic of a shaft with a
universal Joint. It Is also noted that nonlinearities in the aerodynamic
excitation, possibly due to observed vortex loading phenomena or tunnel
inflow irregularities, are causes of higher order excitation.
175
12.1 (Continued)
Because RPM traces were not available for the data reduction effort, the
harmonic content of the blade strains was determined from a "non-speed
corrected" spectra] analysis. The spectrum values were compared to valuesobtained from calculations of amplitude of the IP bandpass filtered signals.
It was concluded that the blade strain magnitudes from the spectral analysis
were withln I0% (low) of the actual magnitudes.
For comparison purposes, IP strain predictions were also made for the angularinflow cases. Due to test constraints, angular inflow test data collection
was limited to a 2 bladed Prop-Fan configuration at 3° inflow angle.
Predictions were 16 to 31% higher than test data for the inboard response(flatwlse shank moment and radial bending) and up to 13% lower than test for
the outboard bending response. Th|s is consldered reasonable Correl_tton,
especially when it is considered that the measured values were perhaps up to10% lower than correct. (See prevlous conclusion). As w|th earlier testingof the SR-7A 2 ft. diameter aeroelastlc model, correlation of root response
was marginal for the trailing edge, shear and chordwlse gages (Reference 26).
12.2 Aerodynamic Performance Evaluation
Measured aerodynamic performance of the SR-7L Prop-Fan corresponded well with
analytical predictions for the four blade configuration over the entire range
of points tested. Similarly, good agreement between measured and predictedperformance was found for the eight blade configuration at Mach numbers of .70and .73. However, performance of the elght blade configuration was slightly
underpredlcted at .50 Mach number. The characteristic shape of the LAP
performance curves were slmiiar to those observed for the SR-7A aeroelasticmode] wlnd tunnel tests (Reference 27).
12.3 Blade Surface Steady Pressure Measurement
Steady pressure distributions were successfully measured on the blade surfaceat a11 of the 13 radial stations for Mach numbers of 0.03, 0.20 and 0.50, and
at all radial stations except 2, 4 and lO for Mach numbers of 0.60, 0.70 and
0.78. Subsequently, an uncertainty analysis was performed, which demonstratedthat the measurement errors and uncertainties involved In th|s test, were
acceptable.
Studles of the sensitivity of the correction for centrifugal 1oadlng on the
column of air In the blade's pressure tap channels showed that the
assumptions used In processing the data were valld.
During operation at approximate statlc rotor conditions, the Inflow Machnumber was 0.03 +0.015, where +0.015 Is the maximum statlon-to-station
variation In the-Math number. -The pressure distributions were found to be
very sensitlve to Mach number for these conditions, so some respectivestation-to-station pressure dlstributlon inconsistencies exlst In these data.
176
12.3 (Contlnued)
The chordwlse pressure loading Is found to moveaft with increaslng relatlve
Mach number for the 0.60, 0.70 and 0.78 Mach number cases.
Evidence of tlp edge and leadlng edge vortex flows were found in the statlcrotor data and the 0.20 Mach number data.
The inverted leading edge pressure distributions observed during the low
power high Math number conditions, are typical for cambered airfoil sections
operating at Incidence below the design angle of attack value.
Evidence of trailing edge shock waves Is present at the outboard radialstations In the 0.70 and 0.78 Mach number data.
12.4 Blade Surface Unsteady Pressure Measurement
Unsteady blade surface pressure data were successfully measured over the
followlng range of conditions"
Angular Inflow (3°) 0.02 ! MN _ 0.70Unlform Inflow (0°) 0.02 < MN < 0.50
Inflow with Wake 0.03 < MN < 0.50
The uniform Inflow data shows evidence of distortlon. This appears to be a
result of test section Inlet asymmetry.
The angular Inflow and wake data clearly shows unsteady pressure response.
dominant once-per-revolutlon response is evident in the angular Inflow datawhile the wake data i11ustrates twlce-per-revolution response as the
Instrumented blade passes the wake generating post.
A
Slnusoldal response was observed on the pressure (face) side of the blade in
all cases examined for angular Inflow conditions.
S1nusoldal response was observed on the suction (camber) side of the bladeunder low loading conditions. However, under high loading conditions,
non-slnusoldal behavior Is present. The non-slnusoldal response appears tobe a result of leadlng edge and tip vortices which may be distorting the
response. Another posslbillty Is the formation and breakdown of the vortlces
as the angular Inflow or wake Inflow modulates the angle of attack.
177/178
LIST OF SYMBOLS
_mm,-=__# r']_ ..... ,,
A
AN
AM
AF
&
BHP
BF
CCBDT
CCBDW
CSD
Cp
CT
CTNET
C
D
D
EF
LIST OF SYMBOLS
area
Incremental forward centerbody area at tap N
Incremental aft centerbody area at tap M
blade activity factor = 6250 f"°(b/D) X3dX
Hub/tip
area weighting factor
brake horsepower, KW
buoyancy force, N
centerbody drag coefficient without blades =
centerbody drag coefficlent with blades
sp|nner drag coefficient =
Ds
qo (As)
power coeffic|ent =poN3D s
thrust coeffIcient -
poN2D"
net thrust coefficlent =
TApp - BF
poNZD 4
speed of sound, m/sec
drag, N
dlameter, m
excltatlon factor = 9(VT/644.8)Z(p/po)
DCBT
qo (Ac,)
DcB
qo (AcB)
179
F
IRP
M
N
P
P
PA
Po
PN, PM
PT
PSTAG
Q
qo
r
r/R
SHP
T
TSTAG
Ts
T_
V
q
force, N
Infrequently repeating peak
V
advance ratio = 60 --
ND
Mach number
rotational speed, RPM
power, watt
pressure, N/cm 2
pressure forces in the form (P-Po) Area, N
freestream statlc pressure, N/cm 2
static pressure at tap N, M
total pressure, N/cm z
stagnation pressure, N/cm 2
torque, N'm
dynamic pressure, N/cm z
radius, m
fractional radlus
shaft horsepower
thrust, N
stagnation temperature, °K
static temperature, °K
total temperature, °K
velocity, mlsec
blade angle, deg
effIclency
180
w
Subscrlpts
APP
B
BP
CB
CBT
C, COR -
FL
MEAS
N
NET
0
S
t
T
TH
TNET
mass density, Kg/m 3
inflow angle, deg
thrust
p(d_sc area)V 2
dlsc area
tunnel cross-section area
apparent
balance
back pressure
centerbody
centerbody tare
corrected
funct|onal losses
measured
number
net
free stream
spinner
tangential
true
thermal effects
net thrust
lsl/ls2
REFERENCES
mmmJ,_m ....fl_'_J
REFERENCES
I. J.B. Whitlow and G.K. Sievers, "Fuel Savings Potential of the NASA
Advanced Turboprop Program," NASA TM-83736, 1984.
2. J.F. Dugan, Jr., NASA LeRC, B.S. Gatzen and N.M. Adamson, UnitedTechnologies, "Prop-Fan Propulsion - Its Status and Potential", SAE Paper780995, November 1978.
3. "Cost/Benefit Tradeoffs for Reducing the Energy Consumption of the
Commercial Air Transportation System, Douglas Aircraft Co; NASA AmesContract NAS2-8618, NASA CR-137925 Summary Report, June 1976.
4. "Study of the Cost/Benefit Tradeoffs for Reducing the Energy Consumption
of the Commercial Air Transportation System", Lockheed-CalifOrni_ - - -
Company, NASA Ames Contract NAS2-8612, NASA CR-137927 Summary Report,
August 1976.
5. "Energy Consumption Characteristics of Transports Using the Prop-Fan
Concept", Boeing Commercial Airplane Company, NASA Ames ContractNAS2-9104, NASA CR-137938 Summary Report, October 1976,
6. "Study of Unconvent|onal Aircraft Engines Designed for Low Energy
Consumption", Pratt-Whitney Aircraft, NASA Lewis Contract NAS3-19465,
NASA CR-135065 Final Report, June 1976.
7. "Study of Cost/Benefit Tradeoffs for Reducing the Energy Consumption of
the Commercial Air Transportation System", United Airlines, NASA AmesContract NAS2-8625, NASA CR-137891, June 1976.
8. "Study of Unconventional Aircraft Engines Designed for Low Energy
Consumptlon", General Electric Co., NASA CR-135136, December 1976.
9. "Fuel Conservation Merits of Advanced Turboprop Transport Aircraft",
Lockheed-Californ|a, NAS2-8612, NASA CR-152096 August 1977.
lO. NASA Tech. Mem. 83736 "Fuel Savlng Potential of NASA Advanced TurboProp
Program", September 1984.
11. D.C. Mlkkelson, G.A. Mitchell, and L.J. Bober, "Summary of Recent NASA
Propeller Research", NASA TM-83733, 1984.
12. NASA CR 3505 "Evaluation of Wind Tunnel Performance Testing of an
Advanced 45° Swept Eight Bladed Propeller at Mach numbers from .45 to
.85", Hamilton Standard Dlvlslon, Windsor Locks, CT, March 1982.
183
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
C.L. DeGeorge, J.E. Turnberg, H.S. Wainauski, "Large-Scale AdvancedProp-Fan (LAP) Statlc Rotor Test Report", Hamilton Standard DivisionUTC, Windsor Locks, CT (NASA CR-180848).
W.E. Su111van, J.E. Turnberg, J.A. Vlolette, "Large-Scale Advanced
Prop-Fan (LAP) Blade Design," Hamilton Standard Division, _indsor Locks,CT, (NASA CR-174790).
R.A. Schwartz, P. Carvalho, M.O. Cutler, "Large-Scale Advanced Prop-Fan(LAP) Pitch Change Actuator and Control Design Report", HamiltonStandard Division, Windsor Locks, CT, January 1986 (NASA CR-174788).
M. Soule, "Large-Scale Advanced Prop-Fan (LAP) Hub/Blade Retention
Design Report," Hamilton Standard Division, Windsor Locks, CT _ _(NASA CR-17¢786).
B. Huth, "System Design and Integration of the Large-Scale AdvancedProp-Fan, "Hamilton Standard Division, Windsor Locks, CT, August 1984,(NASA CR-174789).
Pierre, M. "Characterlstlques et Possibillte de la Grand Soufflerie
Sonlque de Modane-Avrleux," ONERA Technical Note 134, 1968.
Masson, A. "Essair D'Hellces dans ia Grande Soufflerle de
Modane-Avrieux," ONERA Technical Note 161, 1970.
Pope, A., "Wind-Tunnel Testing," John Wiley and Sons, 1954.
R.M. Reynolds, R.I. Sammonds and G.C. Kenyon, "An Investigation of aFour Blade Single-Rotatlon Propeller in Combination with an NACA
2-Serles, D-Type Cowling at Mach numbers up to 0.83", NASA RM A53B06,
April 13, 1953.
H.G1auert, "Airplane Propellers, Body and Wing Interference", Volume IV,
Division 2, Chapter VIII, Aerodynamic Theory, W.F. Durand, editor Juluis
Springer (Berlin), 1935 (Dover reprint 1963).
"BESTRAN User's Guide, Auxiliary Program ST570", D.C. 3ennings,
December 14, 1977.
Bushnell, P.R., "Measurement of the Steady Surface Pressure Distribution
on a Single Rotation Large Scale Advanced Prop-Fan Blade at Mach numbersfrom 0.03 to 0.78", (NASA CR-182124).
184
25.
26.
27.
Bushnell, P.R., Gruber, M.E., and Parzych, D.J., "Measurement of the
Unsteady Surface Pressure Distribution on a S|ngle Rotation LargeScale Advanced Prop-Fan Blade at Mach numbers from 0.03 to 0.70",(NASA CR-I82123).
D. Nagle, S. Auyeung, 3. Turnberg, "SR-7A AeroelastIc Model Design
Report", Hamilton Standard Divls|on UTC, Windsor Locks, CT(NASA CR-174791).
G. Stefko, G. Rose, G. Podboy, "Wind Tunnel Performance Results of anAeroelastlcally Scaled 2/9 Model of the PTA Flight Test Prop-Fan",AIAA-87-1893.
lss/ls6
,i
APPENDIX A
CHRONOLOGICAL HISTORY OF TEST
APPENDIX A
CHRONOLOGICAL HISTORY OF TEST
Testing of the LAP Prop-Fan started on February 20, 1986 and continued until
April 9, 1986. The test was terminated at this point following the discoveryof severe fretting corrosion in the ONERA drive shaft retent|on area. During
this time period, 55 hours and 3 minutes of test tlme were accumulated.
Testing resumed on February 27, 1987 and contlnued until March 19, 1987. Theintent of the second phase of testing was to collect blade steady and
unsteady surface pressure data utlIizlng a 2 blade configuration. Testing
was successfully completed after accumulating an additional 47 hours and lOmlnutes of test time.
The following tabulation provides a chronological history of the entire HighSpeed Nind Tunnel Test conducted at the ONERA facil|ty In Modane, France.
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1, ReDort No. 2. Government AcceSsion No. 3. Recipient's Catalog No.
NASA CR 1821255. Report Date
JULY 1988
4. Title and Subtitle
LARGE-SCALE ADVANCEDPROP-FAN (LAP)IIIGHSPEED WIND TUNNEL TEST 6. Performing Organization Code
73030
7. Author(s) 8.
William A. Campbell Peter ArseneauxHarold S. l_ainauski
9. P_f_ming Orgini_tion Name end AddressHAMILITON STANDARD DIVISION
UNITED TECHNOLOGIES CORPORATIONONE HAMILTON ROADWINDSOR LOCKS, CT 06096
Sponsoring Agency Name and Addrm
NASA Lewis Research Center21000 Brookpark Rd.Cleveland, Ohio 44135
Performing Or_nizauon Report No.
HSER 11894
12.
10. Work Unit No.
._35-03-01
11. Contract or Grant NO.
NAS3-23051
13. Type of Report and Period Covered
CONTRACTORREPORT
14. Sponsoring Agency Code
15. SuD_emen_ Note=
Project Manager: J. Notardonato, Advanced Turboprop Project OfficeNASA Lewis Research CenterCleveland, OH 44135
16. Abs_a_
High Speed Wind Tunnel testing of the SR-7L Large Scale Advanced Prop-Fan (LAP) is hereinreported. The LAP is a 2.74 meter (9.0 FT) diameter, 8-bladed tractor type rated for 4475 KW(6,000 SHP) at 1698 RPM. It was designed and built by Hamilton Standard under contract tothe NASA Lewis Research Center. The LAP emp_ys thin swept blades to provide efficient propulsiorat flight speeds up to Mach .85.
Testing was conducted in the ONERA SI_ Atmospheric Wind Tunnel in Modane, France. The testobjectives were to confirm the LAP is free from high speed classical flutter, determine thestructural and aerodynamic response to angular inflow, measure blade surface pressures (staticand dynamic) and evaluate the aerodynamic performance at various blade angles, rotational speedsand Mach numbers.
The measured structural and aerodynamic performance of the LAP correlated well with analyticalpredictions thereby providing confidence in the computer prediction codes used for design.There were no signs of classical flutter throughout all phases of the test up to and includingthe 0.84 maximum Mach number achieved. Steady and unsteady blade surface pressures weresuccessfully measured for a wide range of Mach numbers, inflow angles, rotational speeds andblade angles. No barriers were discovered that would prevent proceeding with the PTA (Prop-FanTest Assessment) Flight Test Program scheduled for early 1987.
17. Key Woads(Suggested by Author(s))
Advanced TurbopropProp-FanEnergy Efficient PropellerHigh Speed Wind Tunnel
T_is docu_ wri11 remain under distributi(
limitation uat_I July 1989.
lg. Security Oauif. (of this report) 20. Security Qa_if, (of this I_l t Z_t. No. of Pages 22. _ice"
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