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TRANSCRIPT
sMO TR-81-61
$At DOCUMENT NO, SAI-06.'$IR-051
0 (PERFORMANCE TECHR4OLOGY LEVELrAi
OO (PTP-8 Ill
O VOLUME IX"4<
EVALUATION OF REENTRY VEHICLE NOSETIP TRANSITIONAND HEAT TRANSFER IN THE AEDC HYPERBALLISTICS TRACK G
SCIENCE APPLICATIONS, INC.APPLIED MECHANICS OPERATIONEL SEGUNDO, CALIFORNIA 90245
JANUARY., 1981 ,
::::: REP::: FOR PERIOD DECEMBER 1977 - EPTOSIER 1979
""i CONTRACT NO. F34071-77-C-0126
APPROVED FOR PLIBIC RELEASEJ DISTRIBUTION UNLIMITED$.
"AIR FORCE BALLISTIC MWSSILE OFFICENORTON AIR FORCE BASE, CALIFORNIA 92409 S 204
8112 28,t
This final report was submitted by Science Applications, Inc., 1200 ProspectStreet, La Jolla, California 92038, under Contract ilumber F04701-77-C-0126with the Ballitic Missile Office., AF54C, Norton AFB, California. MajorKevin E. Yelmgren. BJMO/SY1T, was the Project Officer in charge. This tcchnicalreport has been reviewed and is approved for publication.
005I r YELMRN ,adU
Chief, Vehicle Technology BranchReentry Technology DiidsiornAdvanced Ballistic Reentry Systems
FOR THE COMMANDER
.4 04L UM SAt%) SF
Director, Reentry Technology DivisionAAdvanced BAllistic Reentry Systems
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Vol . IX, "Evaluation of Reentry Vehicle Nose- FnlRpr:1/797tip Transition and Heat Transfer in the AEDC ** PERFORMING ORal REPORT NUMOCERIHyperballistics Track G SAI-063-81R-051
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is. SUPPLEMENTARY NOTES ~,,
It. KEY WQR"~ (Contirnue on #,vteo*sie Hj. f nocoety 4" amlnd i by block i=30
Reentry vehicle, nosetip, boundary layer transition
*ABSTRACT (Contn aue an a.,.... side 11 "0040e*f a"d Identify bV bloek nS"e)
Zfoundary layer transition and surface heating distributions on graphite fineweave carbon-carbon, and metallic nosetip materials were derived from surfacetemperature responses measured in nitrogen environments during both free-flightand track-guided testing in the AEDC Hyperballistics Range/Track G.
Innovative test procedures were developed, and heat transfer results were vali-dated agiinst established theory through experiments using a super-smooth tung-
OD ,~ 173 DITIN O I OV 615 BSOETEUNCLASSIFIED
SECURITY CLASSIFICATIOM OF THIS PAGE (19mwia Do&. sntfed)
- - OWNERS
-UNCLASSIFIED
.... sten model. Quantitative definitions of mean transition front locationswere established by deriving heat flux distributions from measuredtemperatures, and comparisons made with existing nosetip transitioncorrelations. Qualitative transition locations were inferred directlyfrom temperature distributions to investigate preferred orientationson fine weave nosetips. Levels of roughness augmented heat trans-fer were generally shown to be below values predicted by state of theart methods.
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SECUmiy CLASSIFICATION OF THIS PACrGE(wh Doae Egn;u.)
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TABLE OF CONTENTS
P age
LIST OF FIGURES .. .............................................
LIST OF TABLES .................................................- xii-
1.0 INTRODUCTION ............................................... 1.
2.0 FACILITY AND MODEL DESCRIPTION .. .......................... 4
2.1 Facility Description .. .............................4
2.1.1 Range Description. .. ..................... 4
[2.1.2 Range Instrum~entation .. ................... 6
2.2 Track G Models .... .................................6
3.0 MODEL SURFACE PREPARATION AND CHARACTERISTICS........10
3.1 Model Surface Preparation .. ........................ 10
3.1.1 Tungsten Models. .. ....................... 10
3.1.2 Graphitic Models .... ..................... 10
3.2 Model Surface Characterization .. ................... 15
t3.2.1 Pre-Flight Surface Characterization . . . 22
3.2.2 Post-Flight Surface Characterization .. 27
4.,0 TEST DESCRIPTION .. ...................... ............... 49
4.1 Test Matrix .. ...................... ............... 49
4.2 Surface Temperature Data Reduction .. ............... 49
4.3 Data Uncertainty. ........................ ......... 55
4.4 Track G Reiated Problems .. ......................... 55
4.4.1 Luminosity Effects from Extraneous .3ources 55
4.4.2 Other Considerations .... .............. 66
5.0 ANALYSIS .. ........................ ..................... 67
5.1 Data Format. .................. ....................... 67
5.2 Modeling and Data Reduction. .. ..................... 70
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PasteW
5.3 Cudes Utilized in the Analysis ... ........... ... 76
5.3.1 The ABRES Shape Change Code (ASCC) .... 76
5.3.2 The SAI CAPER-2D Code . . ......... 79
5.3.3 Matrix of the Analyzed Shots. . ........ 80
6.0 TRANSITION RESULTS ......... ................... ... 86
6.1 Nosetip Transition Correlations ... .......... .. 86
6.1.1 Anderson - PANT Correlation .. . . . 86.
6.1.2 Biship's Transition Correlation ..... .. 86
6.1.3 Dirling's Transition Correlation ... ..... 87
6.2 Data Inferred Transition Location .. ......... ... 88
6.3 Comparisons with Transitiun Correlations ........ 92
6.3.1 Comparisons with the PANT Correlation . 92 j6.3.2 Comparisons with Dirling's and Bishop's
Translktion Correlations .. ......... ... 93 '
6.3.3 Comparisons with Reda-Raper TransitionData .......... .................. ... 93
S7.0 HEAT TRANSFER RESULTS ........ .................. ... 123
7.1 Stagnation Point Heat Transfer Augmentation . . .. 123
7.1,1 Augmentation Factors ........... .. 123
7.1.2 Comparisons with the PANT and Phinney'sLaminar Augmentation Correlations .... 125
7.2 Transitional/Turbulent Heat Transfer Augmentation. . 127
7.2.1 Data Derived Rough Wall Heat TransferAugmentation Factors . . . . . . .. .. 139
7.2.2 Comparisons with the ASCC TCurbulentHeating Augmentation Methodology ..... 163
8.0 SUMMARY AND CONCLUSIONS ........ ................. ... 179
9.0 RECOMMENDATIONS ............ ..................... ... 186
10.0 REFERENCES . . . ........................ 188'
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LIST OF FIGURES
Figure 2.1 Track Photopyrometer Arrangement . 5
Figure 2.2 In-Flight Laser Photos at 152M fromRange Entrance .............. 7
Figure 2.3.A Model Design Detail . . . . . . . . . 9
Figure 2.3.B Track Constrained Model. . . . . . . . . . 9
Figure 3.1 Aerotherm Operating Envelope forStagnation Point Model Testing in Air. , , 11
Figure 3.2 CAPER-ID Pretest Predictions . . . . . . . 13
Figure 3.3 Definition of Surface CharacterizationParameters..... . . . . . . . . . 21
tFigure 3.4 Typical GE 223 CC Nosetip . . . . . . . . 23
Figure 3.5 Recovered Tungsten Nosetip (W/SSR 30). . . 24
Figure 3.6 Preconditioned GE 223 Nosetip. . . . . . . 25
Figure 3.6.C full Section Photomicrograph of GE 223Nosetip . . . . . . . . . . . . . . . . 26
Figure 3.6.D Cross Sectional Photomicrograph of GE 223Preconditioned Nosetip . . . . . . . . . . 28
Figure 3.7 Cross Sectional Photomicrograph ofPre-Roughened Tungsten Raplica Nosetip . . 29
Figure 3.8 Full Section Photomicrograph of 994-2Preconditioned Model .... .. . . . . . 30
Figure 3.9 Cross Sectional Photomicrograph ATJ-SNosetip . . . . . . . . . . . . . . . . . 31
Figure 3.10 Cross Sectional Photomicrograph of 994-2Nosetip . . . . . . . . . . . . . . . . . 32
Figure 3.11 Microroughness Height Distribution forPreconditioned GE CC-223 . . . . . . . . 33
Figure 3.12 Microroughness Height Distribution forGrit Blasted Tungsten . . . . . . . . . . 34
IT_ _LXJI
Figure 3.13 Microroughness Height Distribution forL Preconditioned 994-2 ................ 35
Figure 3.14 Microroughness Height Distribution forPreconditioned ATJ-S . .............. 36
Figure 3.15 Full Section Photomicrograph of CC-223Nosetip - Shot 4871 ............. 38
Figure 3.16 Full Section Photomicrograph of CC-223Nosetip - Shot 4880 ....... ........... 39
Figure 3.17 Full Section Photomicrograph of 994-2Nosetip - Shot 4886 ..... .......... .. 40
Figure 3.18 Cross Sectional Photomicrograph of CC-223Nosetip - Shot 4871 .... ........... ... 41
Figure 3.19.A Microroughness Height Distribution forGE223 - Shot 4871 (Post Flight) . . . .. 43
Figure 3.19.B Microroughness Height Distribution for " IGE223 - Shot 4871 (Post Flight) ...... .. 44
Figure 3.20 Cross Sectional Photomicrograph of 944-2Nosetip - Shot 4886 .... .......... ... 45
Fiture 3.21 Microroughness Height Distribution for994-2 - Shot 4886 ....... ............ ... 46
Figare 3.22 Pre and Postflight Photographs of 994--2
Graphite ............................ ... 47Figure 4.1 Arrangement of Points of Projection for
* .Temperature Recording (Viewing Up Range) . 56
Figure 4.2 Brightness Temperature Distribution on0.4 Radius Tungsten Hemisphere .......... 58
Figure 4.3 Measurement of Low Temperature Surface,Adjacent to High Temperature Zone . . . . 59
Figure 4.4 Minimum Measurable Temperature vsDistance from High Temperature Background 60
Figure 4.5 ICC Photographs of Graphite and Carbon-Carbon Nosetips. ..... ............... 62
Figure 4.6 ICC Photographs of Tungsten Nosetips . . • 63
Figure 4.7 Brightness Temperature Distribution on1/4. Radius Turgsten Hemisphere .... 64
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o (m• -• •• .•-: TL•,. ... ..
-a~ -an n ___ ____ ____ __r_
Figure 4.8 Mean Brightness Temperature DistributionAround Hemispherical Nosetip ....... 65
Figurc: 5.1 Nosetip Mean Temperature Prof4 .1 .. .. 68
Figure 5.2 Mean, Maximm, Minimum and StandardSDeviation Profiles . 9
Figure 5.3 Wall Temperature Interpolation . . . . . 74
Figure 5.4 Heat Transfer Rate Calculations . . . . . 75
Figure 5.5 Computed Heat Transfer Distributions Usedin Evalua-ions of Data Reduction Procedures 77
Figure 5.6 Computed Surfaca Temperature DistributionsUsed in Evaluations of Data ReductionProcedures ..... ............ .. . . 78
Figure 6.1 Transit..on Statistics for Tungsten(Shot 4882) ....................... 95
SFigure 6.2 Transition Statistics for ATT-S (Shot 4974) 96Figure 6.3 Transition Statistics for 99w-2 (Shot 4951) 97Figure 6.3 Transition Statistics for 994-2 (Shot 9501) 978[ Figure 6.4 Transition Statistics for 994-2 (Shot 5018) 98
[ Figure 6.5 Transition Statistics for CC-223(Shot 4880) ........ ............... ... 99
Figure 6.6 Transition Statistics for CC-223(Shot 5068) ........ .............. ... 100
Figure 6.7 Transition Statistics for CC-223(Shot 5069) .... ........... ....... 101
Figure 6.8 Circumferential Distributions of TransitionLocations for Tungsten (Shot 4882) . . . . 102
Figure 6.9 Circumferential Distributions of TransitionLocations for ATJ-S (Shot 4974) ..... ... 103
Figure 6.10 Circumferential Distributions of TransitionLocations for 994-2 (Shot 7951) ..... 104
Figure 6.11 Circumferential Distributions of TransitionLocations for 994-2 (Shot 5018) ..... 105
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Figure 6.1.2 Circumferential Distributions of TransitionLocations for CC-223 (Shot 4880) .... 106
Figure 6.13 Circumferential Distributions of TransitionLocations for CC-223 (Shot 5068) ..... 107
Figure 6.14 Circumferential Distributions of TransitionLocations for CC-223 (Shot 5069) ..... 108
Figure 6.15 Comparison Between Data and PANT Corre-lation for KMS (Free-Flight) . . . . . . 111
Figure 6.16 Comparison Between Data and PANT lor W,ATJ-S and 994-2 and Kman . .. ,. 112
Figure 6.17 Comparison Between Data and PANT forCC-223 and Kmean .............. 113
Figure 6.18 Comparisons Between Mean Transition FrontData and Dirling's Correlation forK - Kuean ......... ............... ... 116
Figure 6.19 Comparison Between Mean Transition Frontata and Dirling's Correlation for
- /n(1+3 5OKmean/RN) ]...........117• ~ ~ ~ ~ -eanea " """
Figure 6.20 Comparison Between Mean Transition FrontData and Bishop'.s Correlation ........ .. 118
SFigure 6.21 Mean Transition Front Data vs. FreestreamF6Pressure ........ ................. ... 119
"Figure 6.22 Mean Heat Flux-Inferred Transition FrontData vs, Freestream Pressure ......... ... 120
Figure 6.23 Mean Transition Front Data vs. (KO) . . 121
Figure 6.24 Mean Transition Front Data vs. Freestream~~Reynolds Number .. .. .. .. .. . .. 122
Figure 7.1 Comparison Between Data and PANT Predictionsfor the Stagnation Point Heat TransferAugmentation Factors .. .. ..................26
Figure 7.2 Comparison Between Data and Phinney'sCorrelation for the Stagnation PointAugmentation Heat Transfer Factors . . . . 128
Figure 7.3 Stagnation Point Augmentation Factors forCC-223 vs. Freestream Reynolds Number . 129
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Paset
Figure 7.4 Stagnation Point Augmentation Factorsfor 994-2, ATJ-S, W vs. FreestreamReynolds Number ...... .... ... 130
Figure 7.5 Stagnation Point Augmentation Factorsfor CC-223 vs Normal Shock ReynoldsNumber . . . . . . . . . . . . . . . . . 131
Figure 7.6 Stagnation Point Augmentation Factorsfor 994-2, ATJ-S, W vs. Normal ShockReynolds Number . ............ 132
Figure 7.7 Stagnation Point Augmentation Factorsfor CC-223 vs. K/O . . ........... 133
Figure 7.8 Stagnation Point Augmentation Factorsfor 994-2, ATJ-S, W vs. K/O ........... 134
Figure 7.9 Stagnation Point Augmentation Factorsfor CC-223 vs. K/6* .... ......... . . 135
Figure 7.10 Stagnation Point Augmentation Factorsfor 994-2. ATJ-S, W vs. K/6* . .... .... 136
Figure 7,11 Data-Inferred and Predicted Smooth WallHeat Transfer Distribution Around Nose-tip - Shot 4882 .... ............. .... 140I Figure 7.12 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundNosetip Shot 4882 ........... . 141
Figure 7.13 Data-Inferred and Predicted Smooth WallHeat Transfer Distribution Around Nosetip-Shot 4963 .................... .... 142
Figure 7.14 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundNosetip - Shot 4963 ... ........... ... 143
Figure 7.15 Data-Inferred and Predicted Smooth WallHeat Transfer Distribution Around Nosetip- Shot 4953 ............ .............. 144
Figure 7.16 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundNosetip - Shot 4953. . . ........... 145
Figure 7.17 Data-Inferred and Predicted Smooth WallHeat Transfer Distribution Around Nosetip-Shot 4974. . ............. 1.46
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Figure 7.18 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundNosetip - Shot 4974 . . . . . . . . . . . . 147
Figure 7.19 Data-Inferrsd and Predicted Smooth WallHeat Transfer Distribution Around Nosetip- Shot 4951 ... . . . . . .. . . . . . . . 148
Figure 7.20 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundNosetip - Shot 4951. . . . . ......... 149
Figure 7.21 Data-Inferred and Predicted Smooth WallHeat Transfer Distribution Around Nose-,tip - Shot 5018 . . . . . . . . . . . . . . 150
f Figure 7.22 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundNosetip - Shot5018.......... . . . . . 151
Figure 7.23 Data-Inferred and Predicted Smooth WallHeat Transfer Distribution Around Nose-tip - Shot 4871 ........................ 152
Figure 7.24 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundS iNosetip - Shot 4871 . . . . . .. .. . .. 153
Figure 7.25 Data-Inferred and Predicted Smooth Wall
Heat Transfer Distribution Around Nosetip- Shot 4880 ......... .............. .... 154
SFigure 7.26 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundNosetip - Shot 4880 ..... ............ .. 155
Figure 7.27 Data-Inferred and Predicted Smooth WallHeat Transfer Distribution Around Nose-tip - Shot 4954 ....... .............. ... 156
Figure 7.28 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundNosetip - Shot 4954 ..... ........... ... 157
Figure 7.29 Data-Inferred and Predicted Smooth WallHeat Transfer Distribution Around Nose-tip - Shot 5068 ....... .............. ... 158
Figure 7.30 Ratio of Data-Inferred to Predicted SmoothWall Heat Transfer Coefficients AroundNosetip - Shot 5068 ..... ............ ... 159
Figure 7.31 Data-Inferred and Predicted Smooth Wall"Heat Transfer Distribution Around Nose-tip -Shot 5069 ........................ 160
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riFure 7.32 Ratio of DatA-Inferred to Ptedicted SmoothWall Heat Transfer Coefficients Around Noss-trip - Shot 5069 . . . . . . . . . . . . . . . •
Figure 7.33 Ratio of Data-Inferred to ASCC PredictedRough Wall Heat Tranafer CoefficientsAround Nosetip - Shot 4882 ....... . 167
Fituire 7.34 Ratio of Data-Interred to ASCC PredictedRough Wall Heat Transfac CoefficientsAround Nosetip - Shot 4953 168
Figure 7.35 Ratio of Data-inferred to ASCC PredictedRough Wall Heat Transfer CoefficientsAround Nosotip - Shot 4974 ........ 169
Figure 7,36 Ratio of Data-Inferred to ASCC PredictedRough Walt Heat Transfer Coaffc•c.entsAround Nometip - Shot 4951 ..... . 170
Figure 7.37 Ratio of Data-Inferred to ASCC Pre0ictedRough Woli Heat Transfer CoefficientsAround Nosetip - Shot 5018 ............. 171
I Figure ".38 Ratto of Data-Inferred to ASCC PredictedRough Wall ||wAt Transfer CooffcientsSAround Nosettp - Shot 4871 ... ........ .. 172
Figure 7.11) Ratio of Data-Inferred to ASCC PredictedRough Walt Neat Transfer CoefficientsAround Nosetip - Shot 4880 ....... . 173
Figure 7 40 Ratio of Data-Infe'ared to ASCC PredictedRough Wall Heat Transfer CoefficientsAround Nosetip - Shot 4954 . .... 174
Figure 7.41 RatLo of Data-lnftrred to ASCC PredictedRough Walt Heat Transfer CoefficientsAround Nosetip - Shot 5068 ........ 175
Fi ur ,' ,42 Rat to of Data-inferrod to ASCC PredictedRough Wall Heat Tranufer CoeaffteentsAround NooetLp - Shot 5O9. ........ 176
YFgure .. 41 Ratio* of Data Derived to ASCC-PredietedRough Wall Heat Transfer Coefficients . , 177
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LIST OF TfABLES
Page
Table 3.1 APG Operating Conditions ... ........... ... 16
Table 3.2 Model Number Designations for AblationPreconditioning Tests ......... ............ 17
Table 3.3 Ablation Preconditioning Results ......... ... 18
Table 3.4 Sumnary of Preflight Surface MicroroughnessCharacteristics ..... ............... ... 48
Table 4.1 Test Matrix ....... ................. ... 50
Table 4.2 Test Conditions ....... .............. ... 53
Table 5.1 Matrix of Analyzed Shots ... ........... ... 81
Table 5.2 Launch Conditions of Analyzed Shots ..... .... 82
Table 5.3 Trajectory Information of Analyzed Shots . . . 83
Table 6.1 Nosetip Transition Front Location ...... .. 89 ITable 6.2 Transition Location Statistics. ........... 109
Table 6.3 Calculated Flow Properties at MeasuredTransition Location ......... ............. 110
Table 6.4 Inferred Roughness Results using PANTCorrelation ....... .................... 114
Table 6.5 Comparison of Range Data with TransitionModels ........ ................... ..... 115
Table 7.1 Stagnation Point Heat Transfer Augmentation. 124
Table 7.2 Heat Transfer Augmentation Factors .... ..... 162
Table 7.3 Maximum Roughness Reynolds Number on Nosetip 164
Table 7.4 Ratio of Data Derived to Computer HeatTransfer Rates ..... ............... ..... 166Ii I!
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ACKNOWLEDGEMENTS
This research was sponsored by the Space and MissileSystems Organization (SAMSO) of the Air Force Systems Command.Capt. R. Chambers was the project officer. The project monitorwas Mr. W. Portinier of the Aerospace Corporation. The authorswish to express their thanks to Mr. Richard Raper and the staff ofthe AEDC VKF Range G facility for their competent support and manyvaluable contributions to this research effort.
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1.0 IYTRODUCTION
Analysis of the flight data base, coupled with key'qround test programs, has shown that nosetip related effectsmarkedly influence reentry vehicle performance. In clear airenvironments e':he twc major phenomena affecting performance arehosetip transition and heating augmentation. The understandingof these effects is essential for both advanced ballisticvehicles 'and for evolving maneuvering vehicle designs.
itNosetip transition is triggered by some roughness height
characteristic of the particular material. Since real materialsposseas-a definite statistical distribution of roughness heightit is not certain which roughness height causes transition.Also, real material may not have a uniform surface roughness
distribution, which in turn results in asymmetry in the transitionprocese. For woven carbon-carbons, for example, there mayexist a transition-front shape which is related to the weavepattern of the material. Once the onset of transition occurson the nosetip, transitional and turbulent flows exist andcause higher surface heat transfer rates (which can be furtheraugmented by the surface roughness) which will accelerate theablation process. These effects will bring about geometryvariations in the nosetip that will greatly affect the vehicleaccuracy and survivability.
To properly understand whether or not nosetip Pliapingpresents a problem to any vehicle design, and to intelligentlypursue materials and designs for advanced system nosetips, itis necessary to have at least an adequate understanding of
those physical processes which control the tendency of thenosetip to change shape.
The desire to predict the fine structure of nosetiptransitio;n, and the rough wall heat transfer that follows it,has been part of reentry vehicle work for a decade. tE arly wellregarded work under the ABRES PA14T program generated volumes of
SZzb -1-
wind tunnel data on rough surface heat transfer and transition.
Similar wind tunnel experiments were carried out under ABRES
STREET programs, and complementary data has been collectedunder a number of other DOD and university programs. Untilrecently, however, this data base has been almost exclusively
wind tunnel derived, using metallic models with artificial
roughness in low Mach number/marginal Reynolds number flowg.
The need for a valida..ion of this existing data~ base via
experiments using real materials in flight-replicating environments
has been recognized for some time. Holden, (Ref. 1) at CALSPAN,
has been conducting high Mach/Reynolds number experiments on
room temperature metallic models in' his shock tunnel facility,
and has produced results in serious conflict with the existing
data base. Reda, (Ref. 2) at NSWC, has utilized the ballistics
range to study transition and heat transfer on graphitic models,
and has generated results contrary to the earlier wind tunnel
findings. SAI (Ref. 3) collected transition data on tungsten
and graphite models at the AEDC ballistic range and showed
results that are also contradicting to the wind tunnel data.
I Several key questions remain unanswered, and these are:
1) What is the form of the transition processin flight environments.
2) What is the characteristic material roughnessheight that triggers transition.
3) What are the magnitudes of heating augmentationbrought about by the surface roughness.
The object of the present program is to address the
above questions. Tungsten and graphitic materials were tested
in simulated flight environments. These environments are char-
acterized by higih enthalpy, stagnation pressure and freestream
Reynolds numbers similar to those of flight environments. The
object of the testing was to:
1) Investigate roughness effects on nosetip transition.
2) Investigate roughness effects on heating augmentation.
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-I3) Verify state-of-the-art predictive models as to their
applicability to flight environmenits.
This document presents the analysis of the data collected
in the
AEDC Track G facility as well as some free flight data.
The
I ~various problems encountered in the development and calibrationl
of
the track are briefly discussed. Transition and heat transfer
results
are presented and compared with state-of-the-art predictive
meth-
odologies. Recommendations are made for future facility utilizations
and directions to be taken for additional work.
ii -3--7,x-,
2.0 FACILITY AND MODEL DESCRIPTION
2.1 Facility Description
2.1.1 Range Description
The AEDC Hypervelocity Track G is an advanced ground
test facility in which a gun-launched test article is confinedto a straight-line trajectory by four surrounding guide rails(Figure 2.1a). The facility is designed so that models can
be launched at velocities up to 20,000 ft/sec, guided through
1000 feet rf controlled test environment, and recovered withoutdamage in & 500 foot long decelerator tube. This unique facility,which became operational in early 1977, overcomes hyperballistic
[ range limitations associated with dispersion of the free-flight
model and the absence of a .aodel recovery capability.
The major subsystems comprising the Track G test facility'S~are:
e 1) A model launcher device;
r 2) a model guidance system including tha trackV and its ancillary hardware;3) a model recovery system that is used to dis-
sipate the kinetic energy of the test articlewithout significant damage;
4) a test model that is either a full- or reduced-"scale flight vehicle;
5) an environmental system, the basic componentof which is the 10-foot-diameter range tank,to provide a wide range of environmental simu-lation such as high altitude flight in clearair, erosive particle encounter, or specialchemically inert environments;
6) an instrumentation system capable of in-flightdata acquisition.
Detailed description of the facilities have been pco-vided in Reierences 4 and 5. The basic capabilities of the range
are similar to the Range G free flight test configuration used
for the studies in Reference 3.
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2.1.2 Range Instrumentation
The same basic instrumentation was used for the TrackG tests as was used for the free flight tests discussed inReference 3. In-flight surface brightness temperatures ofthe model nosetip were measured by five photopyrometer systems
(References 3, 4, 5) Figure 2.1b, that could be located at the
following down-range locations:
Photopyxometer (ICC) Distance from RangeStation Entrance, m (ft)
1 10.59 ( 34.73)4 30.38 ( 99.63)
11 72.55 (237.48)
19 121.38 (398.11)29 181.85 (596.48)
41 255.02 (836.48)
For the current series of launches, the ICC Station 11was not available. The opticc• 1. system was arranged such that
the luminous nosetip was viewed from 5 degrees off the flightaxis, thus minimizing motion blur and shadowing effects. As
in the free flight configuration, extraneous shock layer radia-tion and surface chemiluminescence which may influence the sur-
face temperature measurements were surpressed by purging theregion where the image was to be recorded by helium.
Various other high speed photographic, electro/opticaland electronic systems used in aeroballistics ranges are also
used in Track G (Reference 4). Given the accurate location of
the in-flight model, stereo-photographic techniques permit moni-
toring the noseitp surface conditions photographically (Figure
2.2).
2.2 Track G Models1 "IA Track G model is required to withstand acceleration
loads of up to 200,000 g's during launch as well as extreme
S- . .. .•- :" •:•-• • - • • • i • •••i •,• ! ••.-
-1
4880 4886 4955
223CC 994-2 FWFCC
iii
4882 4883 4884
W!PR W/SSP 30 W/SS
4887 4963W/SSR 15 WASS
FIGURE 2,2. IN-FLIGHT LASER PHOTOS AT 152 M FROM RANGE ENTRANCE
-7
r\
heating rates, encounters with erosive fields, thermal degra-dation over the 1000-foot test environment, and finally decel-eration loads of up to 120,000 g's during recovery. In addition,the carrier for the test specimen (nosetip) must continually
* reconform to the launch tube bore to compensate for carrierwear during the launch process and to form an adequate bearingsurface while in contact with the rail guidance system and
recovery tube. The external model components (Figure 2.3a)are (a) carrier, (b) carrier heatshield, (c) specimen holder,and (d) test specimen. Attachment of the specimen carrier tothe specimen support shaft is by means of a shock-absorbingthread pattern, whereas t:.. test specimen is attached usinga swaging process. The test specimens comprising the nosetipof the model were 4" ond 0.4" radii hemispheres fabricated fromh i a variety of materials depending on the specific test objectives.Phenomenology tests used the better characterized tungstenand graphitic materials, ATJ-S and 994-2. Real-materialassessments were made on carbon-carbon composites such as GE223 and fine weave pierce fabric (FWPF) woven materials. A
photograph of one of the current models in flight guided bythe rails is shown in Figure 2.3b.
Details of the model surface preparation and characteri-zation are discussed in Section 3.0.
!
: -8-
'i °-
Specimen Internal StrutHolder / (Titaniu'n)
Z•est Specimen
Heat Shield (PolycarbonateResin)
FIGURE 2,3A. MODEL DESIGN DETAIL
F
FIGURE 2,3a, TRACK CONSTRAINED MODEL
- . I
• • -9-
3.0 MODEL SURFACE PREPARATION AND CHARACTERIZATION"
Model nosetips were fabricated by AEDC from tungsten,ATJ-S, 994-2 graphites and carbon-carbon composites, GE 223and fine weave pierced fabric (FWPF). The tungsten models were¼" and 0.4" radii hemispheres and the graphitic models were3/8" prior to preflight preparation in the Aerotherm Arc Jetand approximately 0.4" after preconditioning.
Samples of each type of model was sent to the SAIMaterials Sciences Division in Santa Ana for careful characteri-
zation.
3.1 Model Surface Preparation
f 3.1.1 Tungsten Models
Tungsten models were preconditioned in the same manneras previously described in Reference 3. The resulting surfaceswere termed 1) nominally smooth for surface roughnesses of rmsvalues of about 25 pin.; 2) super smooth for a surface roughness
of rms values less than about 10 pin., and 3) preroughenedsurface using grit blasting to achieve surface roughnesses inthe range of 100-300 win. Only limited tungsten models wereused in the current test series.
3.1.2 Graphitic Models
To simulate pre-transition surface roughnesses occurringin actual flight, graphitic models were preconditioned in a plasmaarc facility. Conditions were chosen such that laminar ablationoccurred on the nosetip. The preconditioning tests were per-formed at the Aerotherm 1.5 MW Arc Plasma Generator (APG) Facility.Figure 3.1 describes the operating performance envelope of thefacility. Test conditions were chosen such that the nosetipmodels would experience laminar sublimation and the stagnationpoint would recede approximately 30-40 mils.
"I 0
_ -- O-
18NOZZLE EXIT DIAMETER, IN, :•- I 3 , I
•- 0.4 0.6 -1-- 0.3 41 .00- 1500 "•TT SEC -- Typical SO,000 FT/SEC.
400 Manned Lifting Reentry•,..200 .. .•- • /Trajectory
100 , , Shuttle Worst Heati Ing Rate-
j:• • • Supesonic Anode
Z--
•' •_ • "., . •Subsonic -
L0 10.. 10"1 110.1 10" 100 10, 10,
MODEL STAGNATION PRESSURE. P1. ATM
(Convective Heat Flux to a IVA Inch DiameterFlat Face Model as a Parameter)
FIGURE 3.1. AEROTHERM OPERATING ENVELOPE FOR STAGNATIONPOINT MODEL TESTING IN AIR
L -11.-
3.1.2.1 Ablation Pretest Predictions in Arc-Jet Environment
CAPER-ID predictions were made using nominal arc-jet
conditions (Ht = 45,000 btu/ibm, Pst*g - .07 atm, =cw 33302btu/ft sec) to determine recession rates and thus, enabling
test times to be chosen to obtain the desired recession. TheAerotherm (Ref. 6), G. E. (Ref. 7), and Kratch (Ref. 8) ablation
* models were utilized and the resulting stagnation point reces-
sions are depicted in Figure 3.2. Since the Kratch model best
represented stagnation point recession data in previous arc-jettests (Reda, Ref. 2), test times were chosen using this model.Nevertheless, in the present tests, CAPER-ID with the Kratch
ablation model overpredicted stagnation recession by 30%.
3.1.2.2 Models
Graphitic models and three model-sting adaptors weresupplied by AEDC for the preconditioning tests.
Preconditioned models D23, 175-9994-4 and MT-4411A3ware then sent to the Materials Science Division of SAI forsurface characterization and the remaining models were returned
to AEDC for ballistics range tosting.
* 3.1.2.3 Ablation Test and Flow Calibration
The tests were performed in Aerotherm's Arc Plasma
Generator (APG) Facility and utilized the vacuum chamber testleg with a segmented constrictor arc heater. A supersonicanode configuration combined with a 3.5 inch exit diameter
nozzle produced a test stream with a uniform core of at least
2.0 inch diameter. In all cases, the test gas was simulatedair, 76.8 percent N2 and 23.2 percent 02. The stagnation pointof the test model was positioned ½ nozzle exit diameters (1.75inch) from the nozzle exit plane, insuring immersion of the model
in the jet uniform core.
-12-
7.. .... .. . . ..... ..... ... •••r •• '• •i• ,,•.: .. . .-•
H = 45,000 Btu/lbmt
S.12S" - = 0.07 atmstag
.10 - 3330 Btu/ft-sec
c .09
AEROTHERMI ABLATIONH--.08 MODEL
S~z0HN .07
.06 G.E. ABLATION
E-4 MODEL
* o .05
- z0H . 04
St. .03i• DATA
S02 -KRATCH ABLATION'• MODEL
.01
0 2 4 6 8 10 12 14 16
EXPOSURE TIME (SECONDS)
FIGURE 3.2. CAPER-ID PRETEST PREDICTIONS
-13-
F
The stagnation heating rate, q, was measured on the
jet stream centerline at the point corresponding to the modelnosetip through the use of a 2 - ¼ inch diameter, flat faced,
transient calorimeter with a 1/8 inch corner radius (=
.421 ft0). The stagnation pressure was measured with a hinch diameter pitot tube also positioned at a point correspond-
ing to the model nosetip. Stagnation pressures were determinedto the nearest .001 atmosphere.
The centerline enthalpy was inferred from the center-
line measurement of heating rate and stagnation pressure through
the equation
HC = 23.8 q reffS•Pt2
wherekH = local enthalpy on centerline (Btu/Ib
q= cold wall calorimeter heating rate (Btu/ft 2 -sec)
P t= stagnation pressure (atm)
Reff = effective calorimeter nose radius (ft)
The test model surface temperature was measured with
an infrared pyrometer. For all tests, the pyrometer sensingarea was centered on the 430 point (corresponding to an axialdistance of .1 inch from nosetip). Because the model movedforward in its holder, approximately .05 inch when under test
conditions, the pyrometer would be aligned at an axial distanceof 0.05 inches from the nose. The spread in surface temperatures
could be accounted for by variations in axial movement by the
model from test to test. the pyrometer viewed the model througha quartz window and the pyrometer output was corrected for the
transmittance of the quartz. The surface emittance was assumed
to be 1.0 for all tests.
-14-
L- " " . --
Each model was sized and weighed before and after each
test firing to determine the stagnation point recession and the
total mass loss. Model exposure times were nominally 18 seconds,with the exact times determined to the nearest .1 seconds from
oscillograph data. Only the very first run had a test time
appreciably lower than 18 seconds in order to check for unifomn
ablation around the nosetip.
The arc-jet operating conditions and the model test
results are summarized in Tables 3.1 through 3.3 respectively.
3.2 Model Surface Characterization
Surface roughness characterizations were performed bytaking photomicrographs of model surface cross-sections or
K. their replica and measuring the surface macro and micro rough-
ness. Surface macroroughness and microroughness are defined
as the surface roughness measurements made under a magnification
of 50X and 350X respectivley. (The definition of macroroughness
here is different from that used for weaved carbon-carbon mat-
erials, e.g., CC-223.) The measurements taken were roughnesselement height h, roughness element width w, and the spacing
between the roughness element peaks, L (Figure 3.3). Thep
equivalent sand roughness, ks, has been calculated from themeasurements using the method presented in Reference 9. The
average element spacing, D is given by
D =(3.1)
and the sand grain roughness by the equations
kJ
ks - 1.90k - 139.0 X , X > 4.93
k (3.2)= 0.0164 X3 . 7 8 X > 4.93
-15-
r
16A- A
-c M%w c
w -j
LJ j c U o N o cm.a
II000
J-1
C' In In "ri n ~In ein IN- ior -W -it D
-JL
cc U.) Nn %o "') In r.0% -
cc InC
W)JO im 6 n ul UO in W) In 6) L) n V
W 1W - - - 6 - ; - - ; 6 -
-j I " " N - N - N7 -' ND w N %D ) -n NCID 10 0 0 40 0m 0 0) m C0 0 0 0
c.,- Ifc? I N) 6 "n "f in in m. -W~0% 0w c" Ien m' C") m' ' ' ' ' C4) C'
-16-r
TABLE 3,2. MODEL NUMBER DESIGNATIONS FOR
ABLATION PRECONDITIONING TESTS
MATERIAL MODEL NO.
223 Carbon-Carbon MT4411 Al
MT4411 A2
I "MT4411 A3
MT4411 A3
ATJ-S F23i " F45
D23
994 17S-9994-117S-9994-2
17S-9994-3".17S-9994-4
-17-
g .*fI
t U ,, t
3 i ii JI z
4-I -- ,
-J ii "-i l I lii!i
. .,___ _ _ _ _0J
II
Z .L.. .- ,
I-
-J4-18-S-,,l i . -- ""-'-' ..- '_7,•'r' ."'T ..:.- ";..... ,L ... . -"-- ' ' • -`:'•' " -
i. . . ii 'IAI- w
4c - 4 4 i 4q g o ° °o# N 4o
- ir w .--* 'I,-. * -
rJ to #
-- -o -N
it.
z - a' z- - ý qz6 . -p. . t W V e
0 4c
4 a I
w ii -a 9 9&
s ii i tA f%
IA a Z~A~INA__ _ __ _ 0
a. S ~ ~ #4 #'IPA#
uh I gU E
'ASri b6-1S Ita 0w AD- %A -
C-1w ; c
CA IV ' W% .@o
4 I0 in qq Ur *
4.0 0.
OA 4 - 0; 0ý a; 0; 0 C
aA 0 wl CI 0 " 0m cc N"
CM V ~ '
b- 4 % i %a %a4 4L g 0 4%0% f a. p. ~ * i
ft~C "r CDa - C6
CO f ~ 0 4 CM C4 C4 (1k ChLU. li0p P. W a. P. 0 0 I 0.~.* WI W WI I WI WI W WI I W
W 0- a '%a toIQ .
OZ 42 0 A P 0 U nCý 40 0% cc WI co 0% F.. 0oo cc 4 0 cc. cis Go W- P.40 CD
cc) CD **
42 .
o e WC0
-C N N N
LU-w -W -w -L -L a, -
ce
i Q-D 40C
rx 0 wl Ln 0l 0% 1. -
0 - 0 C20 0 C
-20-
h IOPTICALLY DEFINED APPARENT SURFACE
if Average of all roughness heights measured
I Adjusted roughness height (by a factor of 4/v)
ks Equivalent sand roughness
W Average width of roughness elements
rp Average peak-to-peak spacing between roughnessp eelements
FIGURE 3-3. DEFINITION OF SURFACE CHARACTERIZATION PARAMETERS
Ij
-21-
740
The scaling parameter, X, is given by
Xa2.9 D(3.3)
Where the average roughness height1 1 is given by
(3.4)
The 4/rt factor accounts for the failure of a plane cross-sec-tion to pass through the peak of a hemispherical roughnesselement.
Since the roughness characterization of the tungsten
models was to be performed on a sample which was not to be
damaged, a replica was made which accurately duplicated the
tungsten sample's surface. Photomicrographs were then taken and
analyzed. Since the graphite and carbon-carbon samples were
I. ~made in duplicate, an actual sample was cut, polished andIcharacterized using standard techniques. Both preflight andpostf light characterizations were performed on the arc jet
pre-conditioned model nosetip. Typical preflight and recovered
postf light nosetip photographs are shown in Figure 3.4 and 3.5.
3.2.1 Pre-f light Surface Characterization
Three graphite models, pre-conditioned in the Aerotherm
Arc Jet and a tungsten model were characterized. Stereo, macro-photographs were taken of two graphite specimens (994-2 and ATJ-S)I
one carbon-carbon sample (f rom billet 4411-3) and a tungstenmodel which had been grit blasted. The carbon-carbon sample wasunusual in that it had matrix material as the least ablated phase
on the face of the sample (Figure 3.6a). Photographs taken at450 (Figure 3.6b) indicate a stair-step pattern in this orien-
* . tation. The extension of the matrix phase above the surface of
* the specimen is also indicated in low magnification photomicro-
graphs (Figure 3.6c). Photomicrographs at approximately 340X of
-22-
FQ.
CL
ct a
C-
La.
u-I.
In 1- n
U-
-
w
-23-
.J~24 - -'7
)
1 9.
I a
FIGURE 3.5. RECOVERED TUNGSTEN NOSETIP (W/SSR 30) - SHOT 4883
-24-
as. ~VUU~ ~ask
A)
B)
FIGURE 3.6. PRECONDITIONED GE 223 NOSETIP A) HEADON VIEW
B) 450 VIEW
-25-
a LLJ
-4ý- r7U 'r .
am A"-S 2'** #~t~i6~' .,
- -U
.OI
LU
-26-
the 223 specimen cross sections (Figure 3.6d) were taken with
polarized light and clearly show the presence of transverse
oriented graphite in the z yarn bundles. This is evidenced by
the brightly appearing globules between the individual fila-
ments in the z yarn bundles. Very few transverse yarn bundles
were on. the surface of the specimen which is an unusual con-
dition for a laminar CC-223 model.
An epoxy replica was taken of the tungsten specimenwhich was then cross-sectioned, mounted, and polished. A 340
magnification photomicrograph is shown in Figure 3.7. The
actual tungsten sample was then submitted for ballistic rangeItests. The remaining two graphite samples were also sectioned,mounted, and polished. A low magnification photomicrograph of
the 994-2 model is shown in Figure 3.8. Photomicrographa of
the sections were taken at approximately 340X are shown in
Figures 3.9 & 3.10. Measurements were then made to determine
the microroughness on the surface of the specimen.These measurements are summarized as distribution plots
showing cumulative roughness heights (Figures 3.11 - 3.14). As
can be seen, the roughness height for the two graphite speci-
mens are very similar. The average roughness heights for
both the carbon-carbon material and the tungsten replica are
nearly the same but the distributions are not.
3.2.2 Post Flight Surface Characterization
Models from the following shots were recovered from
the track tests and were submitted for microstructural char-
acterization.
Shot Number Material
4871 CC-22 3
4880 GE CC-223
4886 Graphite 994-2
-27-.
NJ jli
00
I-
C',,
10
I--
-28-,
It 21
-LJ
0~
.-JU,
CC
CL
0 u
P=
0i
41, -
-29-.
U-
-APw
-- van w<1 5
9' w
va.. -L~cro
W j f rD $S
* k 4 v1 S * *4~0.
.4& ~V 5 j 4 w-4
"A.; z-
Zz U...
. L
R .. t4
!, oc... 142
tat -itJq-4 izrN
LL.
00&'
~ -' LL
- -~ ~ zc*.u. - -30-
00
0j
CD,
I I VA
40-
0U
V. J4 0I' cn)
auk.0I-
0Z
r 7 -~- - -. (
F -.
I-
x
-Z-
C.)
(onLL
C4
C/)
-32-
LO
(I)m
0 La
cn0 (D
ol~ 0
cicz
co LU
4uoi A49nioC
-33-i
f
:.i00
80
64 60
040
20
0.0 0.05 0.10 0.15 0.20
Roughness Height (10-3 inches)
FIGURE 3.12. MICROROUGHNESS HEIGHT DISTRIBUTION FOR GRIT
BLASTED TUNGSTEN (PREFLIGHT)
-34-
100
80-
4. 40U
0 .2 .4 .6 .8
I3
Roughness Height (10 inches)
FIGURE 3.13, MICROROUGHNESS HEIGHT DISTRIBUTION FOR
PRECONDITIONED 994-2 (PREFLIGHT)
-35-___ ____,_ ___ ___--,
--
100-..
80
a . 6 0 t
40!Q
U 20
0
0 .2 .4 .6 .8-3
Roughness Height (10 inches)
/ FIGURE 3.14. MICROROUGHNESS HEIGHT DISTRIBUTION FOR
PRECONDITIONED ATJ-S (PREFLIGHT)
-36-• .:
. • -. -;7:
Each of the models were sectioned and low magnification photo-micrographs (10X) were taken to determine the general appearanceof both the surface and indepth microstructure. These photo-micrographs are shown in Figures 3.15 - 3.17. The appearance
is not significantly different from the preflight photomicro-graphs 3.6c and 3.7. Model 4880 exhibited gross damage insome of the Z yarns of the type generally attributed to pro-cessing. This type of damage has been observed and reported(Reference 10) in most 223 ablation models that have been
evaluated over the past two years. Unless this type of damageintersects the surface, it does not appear to affect ablationrecession or transition performance. The general yarn spacingand pore distribution were similh: in all of the 223 models.
Only the shot 4871 model was characterized in detail, since4880 is quite similar.
Measurements of the surface roughness on model 4871made from CC-223, were made on a microscale at higher magnifi-cations (approximately 200X). ?hotomicrographs chowing thetypical appearance of the matrix Z yearns and transverse yarnsat the surface of the ablation model are shown in Figure 3.18.What appears to be a localized tensile fracture is evident in
one of the Z yarns. This may have been caused by some isolateddebris in the ballistic range. In general, it appears thatthe Z yarns eroded uniformly and preferentially and the matrixmaterial was most resistant to ablation recession. These photo-micrographs were taken under polarized light in order to observethe amount of transverse oriented graphite (TOG) contained ineach of the Z yarn models. The TOG appearance in Z yarn bundlesof CC-223 is characterized by globules of very bright materialwhen examined under polarized light. Only a small amount ofthese globules appear to be present in these models which isgenerally characteristic of material taken from central loca-tions in carbon-carbon 223 billets. Cumulative plots of the
r-- -37-
L Z.
OLLI
. . '.:-' -. .- .. ..~ . I
>e.%
LL- vi
-. U~d ~ ~ ~ LA..
-38-
FC 4
0
oz
w 00
0j
... ~~U (4* ~
(.D
-39-
F Ia
3.-
... 0
0.A. ',.- 0-.41
0 U-Kow-AT 00
u 0
6L.~J4 R -1~I~ 0~,~.w'~--"C;--
~ ~ . *.¶; .4 ~ * Li - j~.
U-
-40--
M7.
0
C.C
06
a.
U-
roughness heights measured for each phase in the composite and
for a summation of the phases are shown in Figure 3.19. Al-
though the roughness distributions (Figure 3.19) are narrower
for the post flight models, the limited samples and test condi-
tions do not permit a general conclusion concerning the effects
of flight on roughness distribution.
High magnification (200X) photomicrographs were also
taken of model 4886, made of 994-2 graphite, for roughness
* measurements. Typical photomicrographs taken using polarized
light are shown in Figure 3.20. Both the small grain size and
the highly uniform distribution of porosity is evident in these
photomicrographs. The largest roughness elements appear to
occur when pores intersect the surface, rather than from pre-
ferential etching of the individual grains in the graphite.
Accumulative distribution of roughness heights measured is
shown in Figure 3.21 and compares closely with the curve ob-
tained from the preflight sample. However, relatively large
isolated defects do develop in the 994 graphite surface during
flight as seen in Figure 3.22 of preflight and postf light nose-tips. The pits in the postf light model may result from range
debris or thermal stress in the graphite from the large tem-
perature gradients, their influence on transition or heat transfer
is not known. While the actual mean roughness height of each
* element is greater than the heights measured in the CC-223
model, the large width and spacing of each element gives the
model a much smoother surface appearance. At this time, however,
appropriate means for considering the effect of roughness spacingdifferences on a microscale level have not been developed.
A summary of the relevant roughness parameters are given
in Table 3.4.
-42-
U))
0
F ~>4 *
LLI
Q)
z x
o00
CC1L uI-0 4Icn
r-I-X
wuFWW
4U)
00
L)
(ACD
0 L1L
a co w
-43-
* 4 JJ
(iJ)
LLI
coo
- Lo
CD r-I
zc~
Cu *j
-44-
L Z/~~ ~ .1 . ...... 7 - 1-11 __ ___ ___ __
ALL- - k - -_ _ _ _ _ _ _ _ _ _ _
0-
C)
00
C-4
U-
owwz
-45-
lO0-
SPost Flight
-. . .- 'reflight80
j• 60
CU 4° • 40-
20-
0 .2 .4 .6 .8
-3Roughness Height (10 inches)
FIGURE 3,21, MICROROUGHNESS HEIGHT DISTRIBUTION FOR 994-2, SHOT 4886FULL SECTION, (POST FLIGHT)
- 6
,' -46-
,r7-•@
11U
~CL'-cc
cc 0
0 Cc
LIL
cr..
ce.
-47-
N-ri-
(D wzi
wnw>Li r-4 M 0
~~.r4 '0 H 4
cowt
-J48
4.0 TEST DESCRIPTION
4.1 Test Matrix
The complete test matrix for the guided track G shots is
listed in Table 4.1 in sequence. as they were launched. Testconditions are listed in Table 4.2.
4.2 Surface Temperature Data Reduction
Photographic pyrometry techniques used here have been
adequately described in the literature (References 11,12,13).
The basic data reduction procedure consists of:
1. Recording on film, nearly head-on, imagesof self luminous models with a high speedimage converter camera (ICC).
2. Recording the image of a carbon arc cali-bration source on identical film and pro-cessing it simultaneously with the modelphotograph.
3. Microdensitometer scanning of the film andconversion of the film density of themodel image to temperature distributionusing the carbon arc calibration photo-graphs.
Geometric projection of the nosetip into the film plane yields
a one-to-one relation between points on the photographic image
and the actual hemispherical nosetip. The brightness tempera-
ture data obtained from AEDC are arranged on the nosetip, pro-
vided a reference point (the stagnation point) on the photo-
graphic image and the projection can be related. For the last
few shots small holes were drilled in the graphitic nosetip
equally spaced approximately 600 from the stagnation point.
This did not affect the overall performance of the nosetip, but
did serve as reference points which were identifiable in the
temperature contour plots. The use of raised pegs may be more
effective.
-49-
V 04, 4.1(0 CZ 44
to 0 w : 0H* 0 41 r-
0'41 4J 0 4. J -
4) 4) U 4 4
4-) *44)~ 04)
0 HW:N ~ 4J
ti M w- 0 r0* 4 UW .
w) A4.. 0
0 0 00) 4N F- 44 C HIV 4
xJ 0) 41 -4H $ ) o 0 0 t 4 ))
r4 0 0~r4~ W(U F-4r-4 0 )W
I- H *V.~rP-4 4) .. 4.r0(U4) I) N .14H 0) 4) v4
04 9) (D 41 :3 4J 4 TA M-H r- r-4o'o 4 0 041 m H M 0 4 )
I- 4) I t 0 ) t D ý )r
0 41) 4) 0) 0) ' 0) N4 1:71~ i
it ()X , A 0)- )r4k ~0)
0)0)0~ ~ ,-H N( (Ch 0% m m
m r-4 Z % IV4 H %.. HH H:$ HL (% M. 0%C C 0 rN~4b" (H r- -4Hr- H Hr- P-4 HH P-H4r4 - -
M% Hr-i H N %" HHH P -4 f - t- r- r- F- 1- r- r4 F-4 V4
(a m -4 r4 lP4 r- r-4r4 r46 q-4 r46 r-
0 0
0ot z~~ 04 0.4 04 4~~I I2% ~
W 0% N A ~ % 3: 2c Ch a: h (
u .n r- w- r- 0 0 m voz ON1ý c c o o () I ~0 0% L0%
064 co6666 44666666 4
-50-
(a -4
-Aq~ QJ a 'V -
U) 0410 ~ u)
~ '-4 .,4*-,41 4M-W 41 41
0) r4 V-4 * 4 J U ) r-4 001 11 F1 41 0
>) 00) X f
4O0 to4 4)~40 0 ON*'4 *,,4) M M U 1-4 4r
~4 0 ( .4 - 4 J r-4 P- r11-4 C4Jz 0). r-0'-4 4J F" (U (
(a 0 4.)4 a 4 5
1 4)(a w 0) r. 4 to0 r4U' U4H3L :1 r4 r4U
J.4O41 m r -)r44- H t 41$ 0 '41 f-4 V-4
ý4 0 m.4.r 4) *r4W 0) 4144-) M 4 W4J Q) V )t () r-4
to 41 1- - 4 - 4
-H MU4 10 U4 4 4) P1 -
tfl ( ý4 . P-4 4 4
0 :. v 4 Qfl a4 ) rq Q
N m VC4e N N, NN C4
a% a%0' ' 00
r-4 4 4 T-4 "i F-4li LAL
E-51
0 ýk
A?W 4 4 )
0%~
in 4) q)
zU 41 0
0) 4)
.ch :3 :3 -W[I~ ' (a 0~4.G 0
- W-
LUU
4)-
'-4 'ti
-E-4
-52
41)
L4.4 0 000 0 0 a 0 0 0 0 0O00000
v-4
MU~~4I 0 N 0 V L 0 0% t- 0% V ~ M 0% 0% C4 N % M - 0% M.
.4.1 4,14 L ~ 4
'4Ul M.. r4 L 4 0ý '. t, c- c4 rý LA r, N r~r*4 'U Lm r-0 N - M~ L -4 a0 0 a ' !n o 0 t Ln %n N m 45 m 4
.4.4 .M4 ko co v r. * - U) N co w . . GO M %a * 0 en
P-i r-4 r- '.0 r- oA 45 u) v- % N r- r-'0 ~ 45 r-4 0% r-4 0%
4Z HJ 'W C'ON MNocor N Nw
U E-4 14g U 41 ý4 C 1 A C ýL 4L
(L 1 - r4 r- - -4 r- -4 P1 --4 r- 4 -4 ALA I V4 r- - r- -4LA r-4 r-4
w>
r- J 0
- >.~U1 00%U0%0%OS0 rL00%0u4~N
4J 04 m N N v w m 0 N rI 0 v I ý * - m Ln N S ý *
-r4 -4, L4 9 0 0 r- , N Nr L O AL L
H) > -4 r-4 P-4 P -4 s-4 r -4 r- r4 r-4 r-4 P -4 P -4 P -4 P -4 v -4 rA P- -4 r--4
Ir, 4UJ
uQ -4 .N .L A mA mA co0 '0 N. v '0 N.0 t-4 LA m. 0 LA a. LA4 a%
C:0)0)QocoC N N0 m cn 00 ON0 IV N N m q O f 00:3 4 - q U LAL V LfALn LA 4 LA45'0 %a %.D0 % % %0 '.4Ln
0z -
0J I" NI NI I Il %N IN I, II--, g g0% NA ON C14'. r3L 4 0:
-53-
.44
NC~ LAH
W 44.44- w
00M
to
Mr r.4J1L LA 4 HA
CL4 0 1
r-4. Ch 0 0O~
W~ cn4 0H
8~ tp0,0 c ;400wwa
ZU rnp4 $
ro-c) 00 4
K I-(0 r4() r
LU 4-) 0l1-j.IU zUo- -4 f-> u~e
0M0
C4-)
LA 0-4 rA C
04 44
-5-'
o----------
The brightness temperature data on the nosetip is arranged according
to the schematic in Figure 4-1. Temperature is given in 30
increments in both e (streanrviise around nosetip) and 0 (aximuthal)
and is stored on magnetic tape to facilitate computer processing.
4.3 Data Uncertainty
Several sources of error in the temperature measurements
were considered in the free flight test report (Reference 3).
Except for some problems unique to the track tests, as will be
discussed below, the uncertainty in the temperature measurements
is still approximately ± 150 -±2000K. As in the free flight
tests, -the estimated uncertainty (95 percent confidence level)
of the free-stream measurements and calculated stagnation con-
ditions are as follows:
Estimated RandomBias Error
Parameter Percent Percent
Free Stream Pressure 0.7 0.3
Free Stream Temperature 0.1 0.2IFree Stream Velocity Negligible 0.5
Total Enthalpy Negligible 1.0
Stagnation Pressure 0.8 1.2
4.4 Track G Related Problems
A number og problems unique to the operation of the
track were encountered at the beginning of the track test series,
these were considered sufficiently serious that the brightness
temperature results were in doubt. A brief discussion of each
of these problems and their solution is given in this section.
4.4.1 Luminosity Effects from Extraneous Sources
Early in the test as in the free flight series, smoothtungsten models were flown (Shots 4875, 4877) to validate the
-55.-
TOP OF RANGE
FIGURE 4.1. ARRANGEMENT OF POINTS OF PROJECTION FORTEMPERATURE RECORDING (VIEWING Up RANGE)
test and data reduction procedures by reproducing a fully
laminar flow over the nosetip. Results from the super smooth
shot 4875, U'iqure 4.2, are significantly higher than predicted
* by laminar theory at the early ICC's and furthermore the distribu-tions are not consistent with either laminar or turbulent
predictions. Since similar free flight data were in excellent
agreement with laminar theory, the present data were considered
suspect. Extraneous illumination sources were thought to
influence the apparent brightness temperature of the nosetip.F The possible sources considered by AEDC were as follows:
1) Exposure flare: internal reflections withintube allow transfer of light from hot imageareas to adjacent cool image areas.
2) Lens reflections; reflections between elementsF of optical system produce out-of-focus images
which overlaiy the primary image.3) Extraneous illumina.tion of model: system can-
not distinguish between self luminosity andlight reflected from model surface.
4) Bleed thru: finite transmission of image con-
verter tube when "of f" allows exposure of modelIand wake to overlay model image.
5) Presence of self luminous, reflecting orabsorbing particulate or gaseous cloudoverlaying surface.
The first source appeared to be the most serious and the effect
is shown schematically in Figure 4.3. The magnitude of the
effect varies with the image intensifier tube, e.g., Figure 4.4
shows the exposure flare for the System 82, Glen I! 'tube for
various background brightness temperature levels.
The temperature for unity signal to noise ratio refers
to the temperature for which the radiance of the signal is
twice that of the background. Thus, even though the minimum
detectable temperature is 1200 0K, if the position of the mea-
sured temperature of 1200 0K, is, say, 0.25" away from a temperature
-57-
IC ST T4,-N 7 5 'lq 2
IF 5 TPTI ON 11. 72' !
IC STRTILN 19 121M tIC STPT I.N 2 9 18 MY 32 IC s5R T rN 41 2,!' r- -i
k -i
• Czl
4i -2.'4 --!
S-- ,.cr -
2.14
M~ *. ! r "I
- -. . .4 -E -E I
Z
"0-' RADIUS TNGSE HEM
-58-
•) I• I • ----
I - l .- ---
S• ,j•BýTu•MoPRATURE
- - MEASURE BDISTRIB1UT0N
OF'TEMPERATURE TEP
THIGH
= I
Ld.
LUJ
TLOW ""-
STFOG - . .- . .
0
DISTANCE
FIGURE 4.3. MEASUREMENT OF Low TEMPERATURE SURFACE,ADJACENT TO HIGH TEMPERATURE ZONE
-59-
2400-
2300-
' SYSTEM2200 - -- SSATURATI ON
2100
S2000
S1900 -ui(n,o 1800 -
z 1700 -
>- 1600 -
1500o•0
j14 001400 BACKGROUND TEMPERATURE, K
ci" 1300~Liw 2700
1200- -1800 200S~1600
1100 -I0 0.125 0.25 0.375 0.5 0.675D. SYSTEM 82, GEN II, 300 NSEC EXPOSURE, 0.3
MAGNIFICATION
FIGURE 4.4. MINIMUM MEASURABLE TEMPERATURE VS DISTANCE FROMHIGH TEMPERATURE BACKGROUND
-60-
of 2700 0K, the exposure flare effect results in a reading of '
1400 0 K. The luminous source for such potential error was the
hot heatshield as seen in some of the ICC photographs for
various shots in Figures 4.5 and 4.6. The use of a resin-rich
heatshield has alleviated this luminosity source as evidenced
in Figure 4.6 shot 4963 ICC photographs on which the new heat-
shield was used. Other modifications to eliminate or reduce
the effects of extraneous illumination included honeycomb lightabsorbers inside the rail sections in the vicinity of the focal
plane, mechanical light blockers, rails painted flat black, and
optical blockers in the ICC's. The 4th item (Bleed through) wasdetermined not to be a problcm and the 5th item, although may
be a source of error in the presence of an erosion environment
or TCNT models, should not influence the current test results.
The effectiveness of these modifications is demonstrated in
Figures 4.7 and 4.8 which are plots of the mean temperature
distributions at the 5 ICC stations for shots 4909 and 4963.p The corresponding ICC photographs in Figure 4.6, particularlyat ICC's 4 and 11, show the effectiveness of the resin rich
heatshield used in 4963, but not 4909 where a significantamount of luminesence is observed on the standard heatshield.
Results in Figure 4.7 for ICC4 which show a flat temperaturedistribution is probably due to the exposure flare from the hot
heatshield. The corresponding temperature distribution inFigure 4.8 is consistent with those obtained at the other ICC's
and agree well with calculations. These results indicate that
the exposure flare problem along with other extraneous illumination
problems have been substantially reduced and probably does not
degrade the data. The exposure flare may still affect the
stagnation region temperature when a turbulent temperature
distribution occurs which results in a higher temperature ring
surrounding the stagnation point on the nosetip. However,
estimates of the effect from studies such as that shown in
Figure 4.4 indicate that the majority of the current resultsshould not be influenced by this.
-61-
LIn
LL.U
0% CI0
C.14
44,. CA
160
coa
1 - I
-62-D
Q w)
42N C. 0
00
PPF.ILA.
ýA <ýV V) 14
co LA LU
v 3C
r:L..
-63-
-oaum
1. STRTIO'N 4 30M Z
IC STRTI !,•11 ,72M wIC STRT1DN 19 12'M 0IC STPTION 2- 2M X
.2 -C 5RTION '41 2b55!
F 24w
I
2.4
rjj-..-
- 2.0
"........... .
020 40 0 so
BODY :INGLE f OLG)
FIGURE 4,, BRIGHTNESS TEMPERATURE DISTRIBUTIONON 1/4" RADIUS TUNGSTEN HEMISPHERE
-64-
, .
3.6
SHOT 4963 DRTRIC STPTION L 30M ZIC STRTION 11 72MIC STRTION 19 121M 0I IC STPTION 29 182M x
3.2 IC STRTION 4i 255M a
* o 2.8
CD
r 2.4
I D
2Lij
0- XC00 a0
~20 9
- ,,ARON HEAPEIA OEIW
z -77 0a
L-)0
x U
1. 2 LI I
nl 20 Li0 60 60
FIGRE4.8 MANBODY RNGLE (DEG)FIGUE 48. EANBRIGHTNESS TEMPERATURE DISTRIBUTION
AROUND HEMISPHERICAL NoSETIP, W.
-77
4.4.2 Other Considerations
Other problem areas that have been considered were
(1) model tilt which was rectified by using a longer wheelbase
model, (2) potential pressure buildup in the blast tank prior to
model passage through the quick opening valve which was deter-
mined to be minimal and can be accounted for, (3) rail. vibrations
due to the launch transients were determined to be preceded by
the model and therefore should not affect the transition results
and (4) range debris parti.cularly near the blast tank exit where
the jet blast may stir up loose particles of dust that would
alter the model surface and possibly change the transition be-
havior. More effective prelaunch cleanup procedures and use of
equal blast tank and range pressures to minimize the stirring
effects will help to reduce the debris present.
-66-
5.0 ANALYSIS
5.1 Data Format
The brightness temperature data on the entire nosetipis available in 3-degree increments, both along the body(represented by the angle 0 or wetted distance 8) and around
the nosetip (represented by the azimuthal angle *). Therefore,there are 120 temperature profiles along the nosetip at
each image converter camera location. If the freesteam andsurface conditions were such that complete axisymmetry
exists, then these 120 profiles would be identical. However,due to the nonuniformity in the surface roughness elements, aswell as the possibility that an angle of attack may exist, axi-
symmetric conditions are not generally obtained. The degree ofdeviation from the symmetric conditions in the data is indica-Ftive of how uniform the surface conditions are. A detailedlook at the individual temperature profiles can shed somelight on whether or not a preferred orientation of transition
location exists. Also, the potential existence of gouges ccn
be identified via hot spots. This task, however, requiresdetailed analysis of the large bank of data generated at eachICC for each shot (a maximum of 600 profiles for each shot).
The brightness temperature data were averaged in the
direction (i.e., over the 120 profiles) to give mean temperatureprofiles along the nosetip at each image converter camera loca-tion. Superposing the mean temperature profiles for all the
ICC stations gives the nosetip mean temperature-time historydown-range. Such information was plotted for each shot (Figure5-1 depicts an example).
In addition to the mean temperature profile3, the maxi-mum, minimum and standard deviations were obtained along the
nosetip, as illustrated in Figure (5.2). The magnij de of thestandard deviation represents the degree of departure from axial
-67-
H:* 'T 3 D 7 i
T P TM I n IT P, T , V. N I'-3 1 F3P 2M
3" 2 , TflTiDN 4. T21 I:-. F C 5jTA~Tt frN 2!'3 •f"2."
i - 2i q NT. 1-.
Li i- I
I-4
F I .. .-. MT
FIS
I- .. IC-' L 4. .
z .-1 ::
t r z :- .. r. -
I4 . :' G - *
FIUE5.,NSTI EN EPR TUR PROI LES •:t:1
FOR VARIOUS DO)WNRANGE LOCATIONS,, SHOT p4963
7,!
.L.
751
DEVAT.O -i S 4
2.j
*' .r-I
4..
,.. . ---
BOYIN- € iL_!
FIUE52-EN AIUMNMMADSADR
"LEIAIN • IJS S] 46
symmetry in the measured temperatures, and hence the nonuni-
formity in surface roughness spatial distributon.
5.2 Modeling and Data Reduction
The objective of the testing and data analysis was to
investigate
1) transition-front location/offset,
2) stagnation region heating augmentation, and
3) transitional and turbulent heating augmentation,
under typical flight environments. Since state-of-the-art pre-
dictive models are mainly based on wind tunnel data generated
under idealized conditions, a secondary objective of the analysis
was to test the extrapolation of these predictive methodologies
to flight environments. These environments are characterized
by both high freestream Reynolds numbers, and stagnation pressures.
The problem of concern is a very difficult one; the
phenomena involved are highly coupled. Transition from a lami-
nar to turbulent flow is triggered by some roughness height
charact•_istic of the particular 'surface. In the transitional
and turbulent flow regimes, roughness further augments the heat
transfer to the surface. Depending on the location of transition,
the stagnation region temperature is markedly affected and may
indicate a stagnation region - transition related heat transfer
augmentation. Furthermore, the surface roughness of real
materials is such that a statistical distribution exists, and
most probably a single roughness height cannot adequately
represent roughness effects on transiton and heating augmentation.
Finally, the roughness height that may correlate with transiton
data may not be the same for the heat transfer augmentation.
The steps undertaken in the data analysis are summarized
as follows:
1. Derive the heat transfer distribution around the
nosetip utilizing a two-dimensional transient heat conduction
code. The code used was the SAI CAPER-2D heat conduction code.
The approach was to utilize the mean temperature-time history
around the nosetip to predict the corresponding heat flux. The
nosetip is then acting like a thick wall calorimeter where bothindepth and longitudinal conduction fluxes are accounted for.
Transition
2. Infer the location of the transiton region front from
the shape of the heat flux around the nosetip. The transition
front was taken to be either the minimum heat flux point or the
location of the intercept of the two tangents to the heat flux
distribution curve.
3. Compare between the data inferred transition frontlocation and that calculated using state-of-the-art transitionmodels, e.g., Anderson (Ref. 14), Bishop (Ref. 15), and Dirling
locaion ndtAtn acuaeduigettro-teatornsto
(Ref. 16) which are built in a boundary layer predictive codelike the ABRES Shape Change Code (ASCC).
4. Infer the statistical distribution of the location ofthe transition front from the individual surface temperature
profiles along the nosetip. The temperature-derived transition
location was at first taken to be the minimum temperature
point, i.e., the point where dT/de = Zero. Due to possible
ambiguities, the transition locations were determined manuallyto be at the intersection of tangents to the laminar and
turbulent portions of the temperature profile. A comparison
between the heat flux and the most probable surface temperaturederived transiton locations was then performed.
Laminar Region Heating Augmentation
5. Extract the laminar region heating augmentation by
direct comparison between the data derived stagnation point
heat transfer and that computed using a laminar theory, e.g.,
Fay-Riddel as given by ASCC.
6. Verify the extrapolation of state-of-the-art
laminar rough wall heatinc augmentation methodologies to
flight environments by dicect comparison with the data derivedaugmentation levels. Phinneys' and the PANT rough wall laminar
augmentation correlations were used in the anlysis.
Transitional/Turbulent Region Heating Augmentation
7. Extract the transitional/turbulent region heatingaugmentation by direct comparison between the data derived
heat transfer distribution around the nosetip and that computed
by ASCC for the smooth wall conditions. Since the augmentation
level is a strong function of the location of the transitionfront, the heat transfer coefficients were based on the same
location of transition. That is, the location of the transi-tion front was set equal to data derived values.
8. Verify extrapolating state-of-the-art transitional,/turbulent heat tra.nsfer augmentation as given by ASCC metholo-logy, to flight environments, by comparing the data derived
and ASCC computed rough wall heat transfer distributions around
the nosetip. Tvo situations were considered here for the
ASCC calculations. In the first the transition locations wereset equal to the data derived values, and in the second situa-tion the transition was computed using the PANT transition
correlation.
A different approach was also taken in the data analysiswhere the nosetip measured mean surface temperatures were com-
pared with those predicted using the ABRES shape change code(ASCC). The degree of agreement between the two temperatureprofiles at all image converter camera locations determines the
capability of the code and its built-in predictive models tosimulate the transition location and heat transfer distribution.
Deviations between the measured and predicted temperature pro-
files can be due to inadequacies in the transition and/or heat
transfer models, or due to incorrect representation of the
surface roughness. This data reduction approach is difficult
due to the strong coupling between transition, heat transfer
and surface roughness characterization.
The advantage of the heat flux approach over the surfacetemperature one is that the derived heat flux is independent of
any boundary layer theory one may use, and represents the net
heat -flux the model is subjected to. The approach, however, is
sensitive to the nature of the temperature-time history and
a maximum number of data points along the trajectory are needed
for the approach to be reliable. Unfortunately only a maximum
of five image converter cameras are available, and for a few ofI the shots the data at all locations were either not availableor unreliable.
one-dimensional transient heat conduction calculations
were first made to determine the sensitivity of the heat trans-
fer to the time-interpolation of temperature data between the
ICC stations. In Figure 5.3, the wall temperature history was
calculated for an assumed cold wall heating input. Then dis-
crete computed temperature data at times approximating ICC
station locations were determined with T(t 0) assumed to be at
room temperature. Interpolation schemes were tested to determine
best fit to the calculated points in between the data points.From to to the first ICC (4), a quadratic itnterpolation fitsthe data very well as cne might anticipate on theoretical
grounds since AT t½ for short times. In between the other
points, i.e., ICC locations, a 3 point Lagrangian interpolation
appears quite good. A comparison of the derived heating rate,
inputing the wall temperature in this discrete-interpolatedmanner, with the initial.ly input heating rate is shown in
Figure 5.4. The results are seen to be in some error forIsmall t, but in good agreement elsewhere.
-73-
40004000 o o o A
• 3000
iL••computed from Caper-ID, Q cw Input
2000S•+ a 5 Point Lagrangian Interpolation
+ 0 3 Point Local Lagrangian
+ + Quadratic Interpolation
3000
4Pt (ms)
FIGURE 5.3, WALL TEMPERATURE INTERPOLATION
-74-
30
Q0 Input
o o Q Calculated from Tw input"at IC Locations with
20 0InterpolationU0D 00
N d
104
0
'-75
I I
•C 1 0 9
9Q
9i
0l0 20 30 .O50 60 7
FIGURE 5,4, HEAT TRANSFER RATE CALCULATII0NS
-7/5-
Similar to the one dimensional case, two-dimensional
analyses were performed, utilizing the SAI CAPER-2D heat con-
duction code, in order to assess the effect of the temperature-
time interpolation scheme between the ICC locations on the
derived heat flux. The following was performed for a .25'
tungsten model with a range pressure P /P = 570/570 torrs, a
launch velocity, U1 = 15,825 ft/sec and a ballistic coefficient11= 10 lbf/ft 2
f
1. For both laminar and turbulent situations the con-
tinuous temperature-tilde histories of the model surface tempera-ture were generated. Figure 5.5 shows the laminar and turbulent
heat fluxes input to the model surface at all ICC locations.
Figure 5.6 depicts the corresponding thermal response of themodel surface.
2. Generate the model surface heat fluxes which cor-
respond to the discretized surface temperature at the ICClocations. Between launch and the first ICC a quadratic in-
terpolation was used, i.e., AT ý, t . Downstream of the first
ICC a 3-point lagrangian interpolation along the trajectory
was used.
3. Compare between the input (step 1) and the generated
(step 2) heat fluxes.
Similar to the 1D analysis, the comparison was found to
be within 1% for both the laminar and trubulent cases, asexpected. The heat flix is nearly one dimensional during the
initial heat-up period of the model surface (the conduction
depth is small, and heat is flowing mainly normal to the
surface).
5.3 Codes Utilized in the Analysis
5.3.1 The ABIES Shape Change Code (ASCC)
This code represents state-of-the-art in nosetip shapechange calculations and utilizes boundary layer computations,
-76-
fv% 0
cc0
a ac
rot ba
5-
C4
0 00 0 4
C14 Xl4 .L'1
-77-'
m a0
JT -7coa~
0~~ UN 0 U
OL~~~ ~ ~ ~ A. -.~~d3 ovin
LALL
.00 LW
0 z
C)
0m 0o
.f~y-~
(c-OLN.) nmuw nu0
-78-
based on the integral technique. The aerothermal environment
,is obtained via predicting both the inviscid (pressure distri-bution and bow shock shape) and viscous (momentum and energy
transport) flowfields. The inviscid flowfield solution estab-
lishes the boundary layer edge conditions necessary tc performthe viscous calculations, accounting for entropy swallowing
effects. The viscous solution provides wall fluxes (heattransfer and shear) as well as boundary layer characteristics
(momentum thickness, shape factor, enthalpy thickness, etc.).
The laminar, transitional and turbulent boundary layer regimes
are also treated. The surface temperature, indepth temperature
profiles and shape change are obtained from solutions to thecoupled energy and mass balance equations at the surface, and
the indepth two-dimensional transient conduction equations.
The code can also be run in the aeroheating mode where the heat
flux distribution is generated for an input surface temperaturef profile around the nosetip.
The boundary layer parameters and wall fluxes are obtained
by solving the integral forms of the axial momentum and thermalenergy equations. The basic heat transfer and wall shear laws
are those of an incompressible flow over a flat plate. Influence
factors are then applied to these laws to account for the
effects of acceleration, blowing, variable properties, Machnumber, and roughness. Details of the code can be found in
Reference 17.
5.3.2 The SAI CAPER-2D Code
The CAPER-2D code is a two-dimensional transient code
which solves the axisymmetric heat conduction equation accounting
for variable properties and shape change. The numericaltechnique utilized is a fully implicit finite difference scheme,
where the governing differential equation together with the
surface relations are quasi-linearized and cast in their finitedifference analogues. The resulting nonlinear algebraic equations
-79-
are solved using the alternating-direction-implicit (ADI)
method. The solution procedure incorporates surfade and internal
regions in order to minimize the overall computational costs.
The~re are two options for the code which are:
1. Heat Flux Boundary Condition: In this option the
heat flux around the nosetip is input as function of time, for
example utilizing an aeroheating boundary layer code, and the
ou.tput is the temperature-time history of the model, i.e., the
thermal response.
2. Surface Temperature Boundary Condition: Here the
temperature-time history of the nosetip surface is input to the
code and the output is the corresponding indepth heat conduction
f lux.
Details of the code are presented in Reference 18.
5.3.3 Matrix of the Analyzed Shots
Table 5.1 lists the shots analyzed. These shots were
chosen based on availability and quality of temperature data.The model surface characterization, and the ICC locations atIwhich the brightness temperature data was available are also
given in the Table. The table indicates that some temperature
smoothing was performed for a few shots to eliminate the effects
of the hot heatshield. Also, the data at some ICC locations
which were not consistent with the rest of the data due to ICC
tube saturation were eliminated. The launch and trajectory
information are given in Tables 5.2 and 5.3 respectively.
-80-
4) -4 4)A -4 4 ).
.4 4 w- 0 41 .4 44 41 00 4) U
( )-.4 44 4 0.1'..' 41 W 41 0
4)~~ V . o
- ) i444 .4 Q) .) .4)t~ ~ -1 LI ) ~ - io-4 -.4 4 4
'-4 i -).0~ ~ ~ ~' 1-44 > 44) U -)04 -4'0 --L4) m' o(a
w. 4)4) 1)4 4) M) M4)0I4)' -4 V'0 .0ý i0' m -V 0 4.410 '. M)t>>1 to to4 %a4 41 v 0. m ).,=4104 4u 4)4))4)cL Q.. .-4 z a -44 0 t0*- M)4 .41.e. 1)4 4. .0 m Q4) 4 410 0
m to 1 -.~44.40 Al 0'A ~ 41 c (U -1.(JI = -41 J w .4 10 IA c UCTI . c m 'a 43 ) 10.1.4 m14 41 to.~ W ". o4 Q.4.4' 0 4) 4)4
(A d 4 rj 4) 4).4 4) (3) = .4) 01.4 4) ?%~.14) 66.4 M~- .0) 0.4 (14UU ) 04in> -4w ).4 > 4 >I d - - W.' '0 0 44)Q w 0. A.' 0. 41 w 04)0)4)4IA Aj4 ZLA 24-4 41'0 V $40 = 4C1Q1.44 a4441. 01 ' wi u' '0 Q0 u
__1__ ATW N41 6 414.4 00))0 040 44)4)4) Q-0L a) 4'a 'a r4. Uo' . ' 00 ý 0 6
0) w .4 4j to4 M.. 61 a ...) o4 -h ..C 4w . c =4 -4 a A q c4 ý4 q. -4
Cl) '.Q 9) % W % W% 0ý -14 4
W 414 0 0 0 M 1C 0% 4 0 0) 0 0 % U
m4 .. 4 14 .-4 - 4 1.4 1-1 Ch
C)4) N4 NA NA NA 0N
'-1 .011 1 0 c
'0 Q 4*ý .J0 4 0 0''0' '0 %m '0 0 4)r
40).0114 'WN. 4 -4 -4 .4 .4 ý-w0 a0 x0 . 0 .0..
41
Q) 4) Q a 3) 0 00AA.0 41 N1 4' 4. -W 4030
14'-1 4 -
-4
II I N N N N
41 0 ~ lLA I. Ln - , w' LA aO 140 0n co 0% m% 0% 0 0 0 0 0
V0 4' V. 4. -w.LA 4 LM Lnt)
4.) m% 1. r, c 4 f en N %D44 W N M% %0 * o a Nq
ca 0% a; 4' r4 4 a; 4' 4' H4H- - ON m% 0h 0% r- I- r% C% in er-
r. ' 0 0 N %V 0 0 0o G)' 0 F- in IA N~ N in in M r.Hw- H( U ") N H- r- q' 0y% C1 0 on 0 (A
*r4 4.) W0 %0 w. in wD in %a in %D %0 inS 4-4 r- r- H_ H- H- HA H4 PH H- r- H
00
uj4. 0 0 Ch co 0 H- IV 0 ON0 0Q 54 f r- V. 4m C) 0 IV~ 0 0. 4 0
04 in H 0 - - N N H H- H- (- c
:34-)~ 4'(' 0 0 0% 00 0 - r 0 0 Na- E-' ( - 0 H% in 0% V 0 1- q* 0
8 9
EW -. r4 Dof l ý rN 4 v: co N NV en qr 1 (In m e IV fn M MLA LM En f4 LM 0% UM n U) Un U
I-z0 %
0I
4t-.)i '. 0 n 0 I
141C4*
en -W82-G _ 0w a
4. CO %0**r M P4 - c*n %o %
TABLE 5,3, TRAJECTORY INFORMATION OF ANALYZED SHOTS
Freestream Stagnation StagnationShot Freestream,Shot rsr Re No. Enthalpy PressureNo. Mach No.
_ _(10-7/ft) (BTU/Ibm) (atm)
24 13.85 6.98 4991 191.3
119 12.48 6.3 4055 155.8
29 11.65 5.88 3533 135.7
4963 4 13.78 6.95 4941 189.4
11 13.23 6.67 4552 174.9
119 12 .60 6.35 4128
151.6
29 11.89 6.0 3681 141.5
4953 4 14.06 1.85 5144 51.7
11 13.9) 1.83 5036 50.5
19 13.69 1.8 4879 49.0
29 13.55 1.78 4776 48.0
4974 4 13.66 2.4 4857 64.8
11 13.k'.6 2.35 4716 63.0
19 11.25 2.32 4570 61.0 J29 13.00 2.27 4396 58.7
4951 4 14.26 2.53 5294 71.5
13. 14.06 2.49 5143 69.4
19 13.83 2.45 4975 6.
L29 13.54_ 2.40 47-13 64.5
-83-
- ..-
TABLE 5.3. TRAJECTORY INFORMATION OF ANALYZED SHOTS (CONT'D)
Shot ?reestream Freestream Stagnation StagnationNo. iCC Mach No. Re No. Enthalpy PressureN(10- 7 /ftN (BTU/lbm) (atm)
5018 4 13.76 2.44 4926 66.4
11 13.50 2.39 4744 64.1
19 13.21 2.34 4541 61.4
29 12.86 2.28 4303 58.1
4871 4 14.11 1.80 5177 50.4
11 13.91 1.78 5034 49.0
29 13.50 1.72 4743 46.2
41 13.18 1.68 4522 44.1
4880 4 13.80 1.22 4956 33.5
11 13.68 1.21 4869 33.0
19 13.54 1.20 4767 32.3
29 13.36 1.19 4644 31.5
4954 4 14.27 0.89 5300 25.2
11 14.20 0.88 5248 24.9
19 14.12 0.88 5187 24.6
29 14.01 0.87 5110 24.3
5068 4 13.85 1.82 4991 50.0
11 13.59 1.79 4802 48.3
19 13.28 1.75 4592 46.1
29 12.92 1.70 4345 43.7
-84-
. ... .. ... . .. . .......... - -" "-• -.... • • • T• •
TABLE 5,3, TRAJECTORY INFORMATION OF ANALYZED SHOTS (CoNT'D)
Shot Freestream Freestream Stagnation StagnationNo. I Mach . Re No. Enthalpy Pressure
o(0 7/ft) (BTU/lbm) (atm)
5069 4 13.73 3.65 4906 99.2
11 13.38 3.55 4656 94.3
19 12.98 3.45 4383 8R.8
29 12.50 3.32 4066 82.3
8i
II
-85-
6.0 TRANSITION RESULTS
In this section the transition analysis results are
presented. Transition-front locations were extracted from both
the data derived heat flux and the temperatuze data distributions.Comparisons between the data inferred transition locations and
those predicted by state-of-the-art nosetip transition modelswere also performed. A brief discription of the transition cor-
relations used in the analysis is given below.
6.1 Nosetip Transition.Correlations
6.1.1 Anderson - PANT Correlation (Ref. 14)
The PANT correlation represents state-of-the-art in
SI predicting laminar to turbulent boundary layer transition. The
model, which was based on wind tunnel testing, is given by thefollowing relations
/k 0.7
Re8 A(at sonic point) = 255 for transition onset0 7 (6.1)
Re 0 ,tr (Fe = 215 for transition location
where ki is the intrinsic roughness, and T is a perturbation para-meter given by
6.1.2 Bishop's Transition Correlation (Ref. 15)
The Bishop transition criteria basically divides the
noretip into two regions: a forward region where the curvature of
the concave streamlines (in the shock layer) influences transi- jtion, and a backward regirn where streamline curvature effects
-86-
on transition are small. This dividing point on the nosetip is
approximately 200 off the stagnation point. In the flow regime
close to the stagnation point transition location is given by
1.23 M 1.96
152 (TTe) 1.23 M 6 (6.2)Re
For e > 200 the transition location is given by
5.6 = 1 + 4.5( Me (6.3)
Where K/D is the dimensionless roughness required to produceI transition at a point on the nosetip where (Ki/D K/D).
6.1.3 Dirling's Transition Correlation (Ref. 16)
Dirling's approach is very similar to the PANT. Thedisturbance parameter, however, assumes a different form. Themodel is based on defining an effective roughness height, K,
which accounts for the nosetip bluntness, at which the transition
parameter is evaluated as follows
K 1 + 350 1 and
(6.4)
-- U- 1.60K K w
The quantities p- and U- are obtained from the Pohlhansen
fourth degree polynominial fit to the compressible laminarboundary profiles.
-87-
6.2 Data Inferred Transition Location
The nosetip transition-front location was determined in
two ways:
1. From the mean temperature data-derived haat fluxdistributions the transition point was determined to be the point
[ of either the minimum heat flux (dq/dS - ziro) or the intercept
of the tangents to the laminar and transitional legs of the heatflux profile around the nosetip. The latter determination required
some engineering judgement, but for the most part was unambiguous.t Transition statistics may be oltained, in principle, from identical
firings of many models of the same material.
2. Statistical transition information may be obtained
from the data of a single flight if it is assumed that the bright-
ness temperature distribution is directly related to the heat
transfer distribution. The brightness temperature data on the entirnosetip is available in 3-degree increments, both along the body
(represented by the angle 8 or S/RN) and around the nosetip (rep-epresented by the azimuthal angle). There are 120 temperatureprofiles along the nosetip at each image converter camera location.
Initially from the brightness temperature data thetransition front location was determined to be the point of
minimum temperature, i.e., dT/dS = zero. At each ICC location
the minima were obtained numerically for each of the 120 nosetiptemperature distributions (each consists of 26 data points). Thedata points were first smoothed locally by applying a first-
degree least squares approximation to three contiguous points
(Ref. 19) sequentially until a new set of smoothed data points
are obtained for the temperature distribution. The first derivativeI
of the smoothed data was then calculated numerically and the zero
slope point was determined. Historgrams of the transition frontlocation were then constructed. However, the results obtained
numerically in this manner were greatly influenced by the numericaldata smoothing and were judged to be inaccur&te. Hence, thetemperature deduced transition data were all hand processed at
each ICC location as follows:
-88-
77 _Z1
I *ar 4" Go CIAI '(' 0 1 n SI I S* aI maf 111 f 0f
0 1-4
414c c
.4 W4
Sr mke in &nU'ef p-Cý a in Ln cM c q 1* InN NCO OCOO Omml
tAI
z le%4
W" E4'! irs l a 1 45so is o 0
1- 4 -- W W W4 -IMUV4 s4M m
-~-4
3w~ *. ad .1P..41 45I CCD 4SC C C C C C C )
00%@M00 -4M OW-I -4-. - 4,_w~ .4 - CS-S- a... t40 -§S
14 U44 . 4ý 4" p jo 1( glg~~I !_ _III__ _IA__a__I_ _4M_ _r-__ _ _ cC
U) uCL-
W~ ~ ~ ~o I E q- 5 a G6fl1. $4 6 w*P 6C wO w~ WS' w46 5 0 A w
-4- "44-4 4 "4- ) -) S
0. v
41
0) 02 " N 4"0 4 N 04 N -40 v- "440.4N
I I I j OW< ~p . 4 4 4 N N NS~ ~
a% aAU' 0
W1 N4 f" m 4 4 go -4 C @ +cc 5 0% 0% 0% We co 4D %
V~ q. 4r 4n i*4 iS n an
-89-
(a) For each temperature profile, the transitionfront location was determined to be the pointof intercept of the laminar and turbulent legsof the temperature profile.
(b) Historgrams showing the percentage of occur-rence of transition at a given location on thenosetip were constructed.
(c) Transition front density functions werethen constructed from the histograms. Theprobability density curves determine the width and
movement of the transition front for each nosetipmaterial.
Table 6.1 shows a comparison between the transition loca-
tions inferred from the temperature-transition histograms andthose inferred from the data derived heat flux distributions.Although, in principle, the heat flux-inferred transition loca-
tion represents a more realistic location of transition, theS~derived heat flux distributions were based on mean temperature
profiles at each ICC. As a result, any preferred orientation inthe transition location was smoothed out in the averaging process.Furthermore, the results show an apparent heating augmentation atthe stagnation point (see section 7) which gives a rather flatheat transfer distribution and cause.s difficulties in defining a
transition point. Hence, the range of heat flux inferred transitiollocations are shown for some of the shots in Table 6.1. Also
shown in Table 6.1 are PANT/ASCC predicted transition locationsusing different roughness and well temperatures to show these
effects on transition according to PANT. Inferring the transitionlocation from the temperature data can be in error due to thelateral conduction smoothing effect of the surface temperature,and also due to the transient response of the model. That is, iftransition offset took place the heat flux distribution becomeslaminar while the surface temperature may still indicate atransitional/turbulent behavior. In this case, the temperaturedata may erroneously indicate a forward moving of the transition-
-90-
- .
T-
front. Tn any case, only temperature data are available to determine
the transition statistics.
Detailed transition statistics have been generated for the
Track G shots that had valid data. The specific shots analyzed
were:
Sht#Material Surface Conditions
4882 W Pre-roughened
4974 ATJ-S Pre-ablated4951 994-2 Pre-ablated
5018 994-2 Pre-ablated
4880 CC-223 Pre-ablated
5068 CC-223 Pre-ablated
5069 CC-223 Pre-ablated
For each shot, the following was determined
(1.) Circumferential histograms of the transition frontlocation as shown in Figures 6.1-6.7.
(2) Circumferential cumulative distribution of the trans
ition front location (showing the transition frontwidth) as shown in Figures 6.8-6.14.
(3) Mean location of the transition front.
(4) Standard deviations transition front.
The transition location statistics are summarized in Table6.2 along with the previously discussed heat transfer derived
transition location and PANT predicted values.
-91-
6.3 Comparisons with Transition Correlations
The data derived from the track G shots as well as datafrom earlier free flights in Range G, are compared with the corre-lations obtained in the PANT program, those derived by Bishop, andby Dirling. These correlations were summarized in Section 6.1.Table 6.3 is a summary of the calculated local flow propertiesrequired to compute the correlation parameters.
6.3.1 Comparisons with the PANT Correlation
Two equivalent means of comparison were used. The fli'stwas to calculate the implied roughness according to PANT using thetemperature data derived transition locations. The second methodwas to calculate transition locations using the characterizedroughnesses in the PANT criteriL and compare this with the ballisticsrange data.
Figures 6.15 to 6.17 show comparisons between the data infe-rred"and PANT computed transition locations. In displaying the resultsthe sand grain roughness type materials (W, ATJ-S, 994-2) wereseparated from the weave-type materials (CC-223). This was done toeliminate any material type effects on transition. For the sake. ofcompleteness some of the free flight results are displayed. Figure6.15 shows that using the Krms as the characteristic roughnessheight of the material in PANT does not correlate the data well.Some data points lie to the left of the laminar boundary as in-dicated in the figure for shot 4963, where the flow was completelylaminar over the supersmooth nosetip. Also, the PANT correlationwas based on peak to valley roughness heights (which was almost 4times the Krms). Figures 6.16 and 6.17 show the comparisons basedon Kmean. It can be seen there that most of the data lie below thePANT line, i.e., transition occurs upsteam of predicted values andvery close to the stagnation point. It was curious for the smoothtungstan free flight data that either the flow remained laminar onthe supersmooth models, or, when transition occurred, it was alwaysclose to the stagnation point.
-92-
Table 6.5 lists the magnitudes of the inferred roughnessheights that insures agreement between data and PANT predictions.For CC-223 the inferred roughnesses were found to be of the orderof the macroroughness. For preablated 994-2 the inferred roughnessheights were only 1.6% higher than the average values for thetrack shots, which indicated good agreement with the PANT correlation.
6.3.2 Comparisons with Dirling's and Bishop's
Transition Correlations
Table 6.5 shows a comparison between the data inferredtransition location and predictions by the PANT, Oirling's andBishop's transition correlations. The comparisons cover a rangeof stagnation pressure from 40 to 197 atmospherees, freestreamReynolds number from 18 to 70 x 106 /ft and mean roughness heightsfrom 0.25 to 0.57 mils. The table lists the transition locationsas predicted by the three predictive models and those inferedfrom the mean temperature distribution around each nosetip. Ingeneral, the Dirling correlation predicted the most forwardlocation of transition based on K mean. Also, all the modelspredicted backward movement of the transition location as well astransition offset, i.e., relaminarization of the flow on thenosetip. The data however, indicated an almost fully turbulentflc'd for the flight conditions analyzed.
Shot 5068 was selected to investigate the effect of thecharacteristic roughness hMight, used in conjunction with thepredictive correlations, on the transition-front location and itsagreement with the data. The results showed excellent agreementbetween the data and the PANT prediction, if the macroroughnessheight (K - 2.5 mils) was used instead-of the microroughness (K -
.25 mils). Hcwever, the use of the ma-roroughness as the transi-tion characteristics roughness height may not be justifiable.Comparisons of the data with the Dirling and Bishop correlationsare shown in Figures 6.18-6.20.
6.3.3 Comparisons with Reda-Raper Transition Data
Recently (Reference 20) transition experiments wereconducted on large preablated CMT graphite nosetips in the AEDC
-93-
Track G facility. The effects of nosetip radius and freestream
static pressures on the transition-front movements were investi-
gated. The data indicated that nosetip radius has no effect on
the transition location. The mean transition front location wasobserved to progress forward with the freestream static pressure.
The correlated relation between the mean transition-front location
and range static pressure was given as
S .0371-P -1.4127R (6.5)
t
Figure 6.21 is a plot of this correlation with the current temp-
erature-derived data, while Figure 6.22 shows the heat flux derived
data. The current data do not show agreement with the Reda-Rapercorrelation.
Other correlations with k/6, Re and Re. are shown in Figures6.23 and 6.24 and do not show much promise in correlating the current
I data.
It is felt that a larger sample of high quality data on a variety
of well characterized materials as was intended in the current progiram,
before its untimely demise, is necessary before any firm conclusions
concerning the veracity of the existing tiansition correlations. The
large CMT models flown by Reda and Raper were certainly in the rightdirection.
-94-
. .. .. - - . - . . . .---------. - ---. - -
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-95-
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-109-
- N1( w D r- a e Or~N COD N% - W .4% - 0 - Lm4 M -Wr -
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103
,-0. /- PANT CORRELATION
77
'" = 215
00 '
04731 W
10 470 994-2,P SHOT 4963 (W, SS, RN: 2"
S4 94-2, PR
.11.0 10. 10
CORRELATION FOR K,(FREI 4774 9---- - - -7- 0 994--2 - -R .- ,--- ',-.-.----SHOT 496 (F REE FLIGHT )
F IT
103TRACK SHOTS - CPEN SYMBOL, R% = .4"
FREEFLI,:-0,' SHOTS - CLOSED SYMBOL, RN .25"
RI
PANT CORRELATION
102 7R" e. pe 215
zLu
I: " * 4731 I!, PQ RI' 4721 994-2,1S
A 4725 14,P R:' : 1010 4 t774 994-2, NSi! 1:, 700 99L:-2,PR
S4784 994-2,PR R[KP '7
,, :• d 5013 9 94-2, .PA ri __
&, ' 4974' ATJ-S,PA
"<> '4953 ATJ-S,PA
1 .1 1'0 10. 100.9p 3
DiSTURBANCE PARAMETER (K i
FIGURE 6,16 COMPARISON BETWEEN DATA AND PANTFOR W, ATJ-S AND 994-2 AND KMEAN
-112-
2• 2. " :."- **" -
10'
102 PANT CORRELATION
Re,(K!x), 215
-~ I
0ze
Lu
cm
I--)r
FALA4 CORR ATIN 2
H ALF SHADD52
"1... I Pe) I
d 4. loPAc.A 5068 22T-CCRPA,)
U 4954 223-CCPA
ICC STATION R C23 ANDK
OP-N13-
F LAGGED 19 rl* ý SHOT 4963 (WS, RN .25")
HALF SHADED 29
Z FIGURE 6,17 COMPARISON BETWEEN DATA AND PANT
FOR CC-223 AND K MEAN
'113
I -_73-
I:1
0o (U 0., 0 WD 0 %a 0 IAN LM7 N. r- 0- N rý E
-W r 0 '0 0'7 74 0- M-. E F- Lm4 0% 7r % -4 e - - - = r n C 4 >
F.C 4w %Q U) 0 "W _n w %a a% r __ 0
O n z
LU~~A 04 W7%0 1%. '0 '094 - 7
U.1 >1.
r- A cc AL L N-F r- r- Fo flFn-rc w=z xU 0) r4 Nrc9 N; LA C4 C4 LA NO4)
a%7q 0 m L 0 F- I "r C4 m f4 en
(n q L0 n w D % -r % % a0APZ r- -- -
U.I- W( M w 10 r- LA) '.0 r-4 '.0 coF % 0 cc cc0 0(Nrn Z It 74
v Ln -w m aq i W "r" q w
M.. 0 * L .id '.0 L JA W 4 LA m A 0 a 0W F- N93 Uo 74 0 V- N0 U
LU. E-4 $4 w4 14 -F4 ).4 -P4 -44 P4 k -P4 P
Eq!'~~~ý 1-414' ~ 4~n 7
744
NJ N 10
41 C4 E- A l
-114-
in i fnN wr- w CoU% .- 0f m N kmW M 0 rri W w m ..f
"~4 -4 - .4-
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N -V CC C4 00 0MM 0 a 0M 0% tM rf% LLM LANN N NI ý4 O f c 4 co NNM 1, 0 4.4e4mfnN % nNq M
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-15
Pf
I-
K[I O5 • "• Ft'Aw 160"•o
SHOT Ic. 5068 0 4o 4380 11
45069 C 19A 4974"501804951c 4882
1 0 . .. . ....0.1 1.0 10.0
PK UKKp, ue
FIGURE 6.18, COMPARISONS BETWEEN MEAN TRANSITION FRONT DATA
AND DIRLING'S CORRELATION FOR K=KMEAN
-116-
Awý "
10-
[PKUKK] 160
oW IR
k Ck
1'~ I.
Ii 4SHOT IC
S5068 04c4880 ol i
S5069 19
S4974 29.25018' 495104882
100.1 1, 0.0
PK UKPe Ue 9
FIGURE 6.19 COMPARISON BETWEEN MEAN TRANSITION FRONT DATA AND
DIRLING'S CORRELATION FOR K - KMEAN/(1+ 3 5 0 KMEAN/RN)
-117-
F SHOT Ir 5068 4
c4880 11S5069 i1
4974Q 2904 5018
1; 4951~.- 4882
I; 12.97 K~ -0.51
10104O0- 10-2
K2RN
FIGURE 6.'20 COMPARISON BETWEEN MEAN TRANSITION FRONTDATA AND BISHOP'S CORRELATION
-118-
1.0
REDA-RAPER DATA
(S/TR,~T ME - .0371 Poo.±2
0
10.1
I-I
E.-4
IC
0 994 Q 4
o 223 11i'>ATJ ) 19
W Q 2 9
0.01 - I I ,.1 RANGE PRESSTURE Pa, (atm) 1.0
"FIGURE 6.21 MEAN TRANSITION FRONT DATA VS FREESTREAM PRESSURE
-119-i .&~
1.0
REDA-RAPER DATA
ýSVTR, MA .0371 :.42
0.<1
0 994 a a
O 223 ( 19< ATJ O 29
0.01 -0.1 RANGE PRESSURE P,,(atm) 1.0
FIGURE 6.22 MEAN HEAT FLUX INFERRED TRANSITION FRONT DATA
VS FREESTREAM PRESSURE
-120-- -
r
.. 0
I, Lo
IR
100
S-ic
c 4
ATJ-S9 0 94-2 13 9
CC-223 29
1 . . . 1 0 1 0 2.r
1, 10
(K10)
*1FIGURE 6,23. MEAN TRANSITION FRONT DATA VS (KI0)
-121-
,,•1
0
rim
102 IC
S0o ATJ-S •1So 994 o 19
c, CC-223 q 2
102
Re 45
<0
4410
1 10 102 103
II
'eRNX 10-
FIGURE 6.24 'IEAN TRANSITION FRONT DATA VS FREESTREAM REYNOLDS NUMBER
-122-
7.0 HEAT TRANSFER RESULTS
The object of this section is to investigate the effects
of surface roughness on the stagnation region and transitional/
turbulent heat transfer augmentation. By direct comparisons
between the data-inferred heat transfer to the model surface,
and the smooth wall calculated values using the ASCC, augmenta-
tion levels can be extracted. Also, the various heat transferaugmentation predictive methodologies can be verified.
7.1 Stagnation Point Heat Transfer Augmentation
7.1.1 Augmentation Factors
The heat transfer augmentation levels were extractedas follows:
1. Derive the model surface heat transfer usinq the
CAPER-2D heat conduction code from the mean surface tempera-
ture-time histories down range.
2. Calculate the smooth wall stagnation point heat
transfer using ASCC. The ASCC prediction should be in agreement
with the Fay-Riddel theory.
3. Compute the heat transfer augmentation factors by
taking the ratio of the data-inferred stagnation point flux
to the smooth wall value.
4. Verify laminar augmentation methodologies by comput-
ing augmentation factors for the test conditions and compare
with augmentation levels derived in step 3.
Table 7.1 lists the data derived augmentation factors J
at the stagnation point for all the shots analyzed. There is
a good agreement between the theory and data for Shot 4963,
where the model surface was super smooth. The data indicated
augmentation factors around unity as expected. For the pre-roughened and pre-ablated models, augmentation due to roughness
-123-
, .. .. _ -. . . .. . .. ,...... ',.._....-,-_ ------ _-__.._.__._..._..-_..- -.. - .. .
TABLE 7.1. STAGNATION POINT HEAT TRANSFER AUGMENTATION
Stagnation Point Heat TransferAugmentation Factors =
Shot # Material ICC Qdata/Qcalculated
ASCC Generated Wall TemperatureWall Temperature Data
4882 W 19 1.2229 - 1.17
4963 W 11 .93 .9619 1.03 1.0329 1.03 .98
4953 ATJ-S 11 .99 1.0219 1.17 1.1529 1.2 1.17
4974 ATJ-S 11 1.04 1.0619 .88 .9529 .92 .97
4951 994-2 11 .92 .9919 1.34 1.2729 1.31 1.27
5018 994-2 11 .87 .9319 .91 .9329 1.22 1.12
4871 CC-223 11 1.05 1.0629 1.18 1.11
4880 CC-223 11 1.25 1.1719 .99 1.0229 .83 .92
4954 CC-223 11 1.02 1.0219 .77 .9429 .62 .79
5068 CC-223 11 - 1.1019 - 1.0529 - 1.17
5069 CC-223 11 - 1.1719 - 1.2529 1.28
-124-
was observed. For the flight conditions tested augmentation
factors as high as 2.0 were observed (Figure 7.1). To elimi-
nate the effects of wall temperature on the augmentation fac-
tors, the calculated stagnation flux was based on the measured
wall tempera'tuze. Listed in the table also are augmentation
factors where the ASCC computed fluxes were based on the ASCC
generated surface temperatures using the transient option of
the code.
7.1.2 Comparisons with the PANT and Phinney's LaminarAugmentation Correlations
Comparisons were made between the measured stagnation
point augmentation levels and those predicted usin4 the PANTand Phinneys correlations, in order to verify extrapolation of
the predictive methodologies to the ballistics range flight
environments. The correlations have the following forms:
PANT Laminar Augmentation
K£ =1 for 2 < 50
= 1.307 in t + 23.09 -*606 - 6.269 for t > 50
2
where 4- ( 2u y k
Phinney's Heating Augmentation
1+ (P2 2RN),2.2
tk k(k 0017 , for < 2.41f =
k.004 , for 3 > 2.41
Figure 7.1 shows a comparison between the data and
predictions of the PANT correlation for all the materials. The
materials were divided into sand grain and weave type materials,
-125-
~ - - -
1!1
U-
.r. -. k -f
39 bt.wO .u
z- 0)
o00e0, IN,
O-3 - wz g
0*"PA aa
0 0 as Q
w
I-z
I~ Xnld a aioIn/i CMV3N viva
- -126-.; '
in order to eliminate effects of the nature of the surface on
the results. It can be seen that for the flight regime, augmen-
tation factors up to 200% were observed, while the PANT cor-
relation did not predict any augmentation. The comparison be-
tween the data and predictions using Phinney's correlation
depicted in Figure 7.2, indicates that the correlation over-
predicted the data for all the materials tested.
No attempt will be made here to correlate the observed
augmentation factors. However, the data will be presented in
terms of all the pertinent correlating parameters Rem, Re 2 , K/e,
K/6* as presented in Figures 7.3 to 7.10.
7.2 Transitional/Turbulent Heat Transfer Augmentation
In this section the results of the heat transfer augmentation
around the nosetip are presented for all the analyzed shots.
First, the heat transfer augmentation factors in the ballistics
range environments were derived by direct comparison of the
derived heat transfer data to the calculated smooth wall values.
Secondly, state-of-the-art turbulent heating augmentationmethodology, as given by the ASC code, was compared with the
data, in order to validate the extrapolation of the predictive
model to flight environments.
The heat transfer around the nosetip is affected by the
following parameters.
1. Transition-front location: transition from laminar
to turbulent flow is triggered by some roughness height char-
acteristic of the particular surface.
2. Roughness effects: heat transfer is augmented bysome roughness height, which may be different from the height
which correlates with transition data.
3. Real materials: surface roughness of real materi-
als is such that a statistical distribution exists, and
-127-
J .2 __ _ _ _ _
Cý
ow C/)
W- 41
ww ZLL
00060
'z
0.011 17-
z zo 0Oit H
- -n
-00~ L.)U.L
- WA"-~---- -- ~ .0
LAL
LIn
00
-z
a) z
U3zw
0 uia) -
Z Cl0 >
LA 'UIx
C,4N
~~C 1-;~6
-129-
U L7
* LU
0
CC.LL.
LLsco
WEI-
I-IwIItLU
LLL
c-.i ,-
m -134-
- - -~ -44
04
LU
0) L
I-
0 0
00
.,4 >
4 ý
r-4 ,.4 N V 0
C' C' 0
u CN (.0
u -Z
u L
0 .N
-1i1(pqenLIo D/ ---- -, -. -
LU
.0
U,.
1 0
00U
0 /,C o4
U)
oc
oo
UL
U P- 4
LID 0
-132-
0
I.-
z
(D04
LU
1. NCNC
b i-Ir4C'J * N4
-pq~ole) I-'/C", 1
-13
LL
CN
-134-
10L
I LL-
C),
Il-
W04cl -
V ý4
u CL- La -ff) zCN ý-w
N:F l'
C4*1
1-C)noVO N~/(e4,p i
-t 135-
4'-
4,
<Ij
0--
LU
0--
r0I-
'0 co
-136
most probably a single roughness height cannot adequately represent
roughness effects on transition and heating augmentation.
The heat transfer in the transitional region is mainly
dependent upon the magnitude of the intermittency factor, width
of transitional region, and roughness effecti. In the fully
turbulent region the heat transfer depends on roughness effectsin the transitional (K+ < 70) and fully rough regimes (K+ > 70).
In order to properly derive the magnitudes of the heat
transfer augmentation factors around the nosetip the effect ofthe transition-front location had to be accounted for by fixing
the transition location at each ICC location. Therefore, to beconsistant with the derived heat transfer data the transition
location should be set equal to the experimentally derived location.Predicting the transition location by some methodology, e.g., thePANT correlation, which may not be consistant with the observed
transition location, can greatly bias the au~gmeuLtation fnctors.
This effect will be seen later in this section.
L ~A suumary of the steps undertaken to derive tha heat transfer
augmentation factors around the nosetip is as follows:
1) Derive the heat transfer distributions fromthe mean temperature-history data downrangeusing CAPER-2D.
2) Infer the transition-front location from theheat flux data at each ICC location.
3) Predict the smooth wall heat transfer distributionaround the nosetip at each ICC locationusing ASCC-aeroheating option (after set-ting the transition locations equal to the dataderived values).
4) Derive the augmentation factors by simply taking the ratio of the data derived heat fluxto the computed smooth wall values at each ofthe body points on the nosetip, and at each ICClocations.
-137-
7Y.
-- - -~ - -
The steps undertaken to verify the ASCC rough wall aug-
mentation predictive methodology were as follows:
1) Use the ASCC code in the aeroheating calculationmode at each ICC location using the mean roughness.
2) Set the transition location to the data-inferredvalues.
3) Compute the rough wall heat transfer around thenosetip using the ASCC methodology.
4) Compare the data derived heat flux to the ASCCcomputwý rough-wall heat transfer and/or comparethe ASCC computed rough wall augmentation factorto the data derived values.
In the ASCC the turbulent heat transfer enhancement
due to roughness is treated via roughness augmentation coef-
ficients given by:
ir h1 = 1 + 0. 3 f (!) g (x)h,t,r
wheref 1 + 0.09 t\ + 0.53 (1 - e
g(x) = x + 1.5 (1 - e-x) for x > 0
= 0 for x 4 0
+ / k+k e Vw0 fts/2
kt is the local turbulent surface roughness and Cf , is thesmooth wall turbulent friction coefficient. In the transitional
region, the heat transfer is computed via an intermittancy factor
as given by the ASCC methodology.
-138-
7.2.1 Data Derived Rough Wall Heat Transfer
Augmentation Factors
Figures 7.11 through 7.32 display comparisons betweenthe data derived and smooth wall computed heat transfer distributionsas well as, the heat transfer augmentation factors around thenosetip. As an example, Figure 7.11 shows plots of the data-derived and computed smooth wall heat transfer versus the body
angle, for ICC 19 and 29 of the tungsten shot 4882, For both ICClocations the transition-front location was set to 79, as derived
from the heat transfer data, and indicated in the figure. Figure7.12, as an example, shows heat transfer augmentation factors asfunction of the body angle on the nosetip (ration of the data-derived to the computed smooth wall heat transfer coefficient).
The super smooth tungsten shot 4963 is of particularinterest here. This shot's data essentially calibrates not onlythe ballistics range data, but also verifies the data reductionscheme utilized in the analysis (i.e., the derivation of the heat
transfer distribution from the mean-temperature-time histories atall ICC locations). For this shot and at all ICC's the surfacetemperature distribucions indicated fully laminar flow on the
nosetip, as expected since the roughness height, Kr-,,, was < .01 mils.The flow parameters were srch that natural transition was not
observed. Figure 7.13 and 7.14 indicate very good agreementbetween the data derived, and computed smooth wall heat transfer
coefficients.
Table 7.2 lists the heat transfer augmentation factors
(heat transfer data/computed smooth wall heat transfer) at thetransition-front location and at 40* off the stagnation point.The details of the heat transfer coefficients ratios are given inFigures 7.11 to 7.32. It can be seen in the table that a maximumof 1.4 was observed for the rough wall augmentation
-139-
SORTA I NOIE %UX
Q WT4e2 !29 ýC=Ii
3C. C.I ~Cr. 4C.0 30. Z C -
-- --- C .-• TE-.L'
"*1
FIGURE 7,11. DATA-INFERRED, AND PREDICTED SMOOTH
'IALL HEAT TRANSFER DISTRIBUTION AROUND
NOSETIP FOR SHOT 4882 (DATA-INFERRED
TRANSITION-FRONT LOCATION)
-140-
[ TZJZ~. sE - Z1~2~~ -
r-7'
~. SHOT 4i8e2 :C:S 'RNz7 ~ (.O
=• " . • -u 2m. M0 " 3 '4..0 :. ^G 0 30. i0•. .50cr ;N~ :OE.:-
SHO 4M8 'C-9 '9!04=7 "t~ ý2
•-w
u
FIGURE 7,12, RATIO OF DATA-INFERRED, TO PREDICTED SMOOTH
WALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIP
FOR SHOT 4882 (DATA-INFERRED TRANSITION-FRONT
LOCATION)
-141-
- U-- - -|- - -! ;. . . .. . . . . .. . . .. . ... . .. .. .... , •Ž J > 7.. . . -• - . . .. _, . .. ..._. ..... . L_ -, .. , -• • ... -. . . ... ... --.. ,- .. .. ---:. . .-
ini"Z
2- oc 5 :
FIUR 7.13 -AAIFRE N RDITDSOT
"-1
14-4
- SM• - 'SS" " --
13-- • "
FIGURE 7.13, DATA-INFERRED, AND PREDICTED SMOOTH j
WALL HEAT TRANSFER DISTRIBUTION AROUND
"I ~NOSETIP FOR SHOT 4963 (DATA-INFERRED
, ~TRANSITION-FRONT LOCATION)'
-142-i--
- -.. ~- -~--
-- SKIT ~4,363 !tI! -.R!WRf 5?OT 49563 'CIS LW
IqN~r- IX
.5 .8 .
-- I
• .
{- I
F/OR § SHOT 41963 (DTAINERE IRNIIO-RN
--- 43
-tIII U • I
FIUE71. AI FDAAIFRE, T RDCTDSOT
FO SHOT' 4963 (DAT"INERRE..RNSITON-RON
LOCATION
•.143
• _. SN " ,9SS :,. i ,t " 25QN:5 OE ; ,. .o " ! c g ... ..
appyr8 toI~ 4 .,, gin~A ;9ftl*O 'LUX
.~~~~~~. .... .l' lpp~. ,.' .
r3
FIUR 7.5 AAIFREADPEITDSOT1 -I
F]
114
-~ ----------
kiL~~D~
SMiT 49g53 !C', ?"W=5 CS WT~ 49S5! IC1g TRWziS aDr c.-I
.- ......... .' . . .........
f
v;;- mk ze S':.c 0 S
I i W T' 4953 TftqN=S C.r k.
Fi
-- j ..
•- .5..
-5-
FIGURE 7.16. RATIO OF DATA-INFERRED, TO PREDICTED SMOOTHWALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIP
FOR SHOT 4953 (DATA-INFERRED TRANSITION-FRONT
LOCATION)
--145-
+ ~ *-' .:
- -
II
I %1
3 -_ 3 C
FIGURE 7,17. DATA-INFERRED, AND PREDICTED SMOOTH
WALL HEAT TRANSFER DISTRIBUTION AROUND
NOSETIP FOR SHOT 4974 (DATA-INFERRED
TRANSITION-FRONT LOCATION)
-146-
ý= ...... ..... ..... .1
SH.. ' 4974 :C21 TRANul' DEC; Kx o SmT '497% IC1g TR"N3S DEO km '
E.5*, ...
-8
^0 cc * s:.o'
ZOd- " R C S6.0C" C J-
ISHOT 0974 !C29 TRQNzlo 0!0 kL-.oX
II
5 55
FIGURE 7.18. RATIO OF DATA-INFERRED, TO PREDICTED SMOOTHWALL HEAT TRANSFER COEFFICIENTS AROUND NoSETIPFOR SHOT 4974 (DATA-INFERRED TRANSITION-FRONTLOCATION)
-147
------ 4
,,-- -"-
'U"IIu
21
_ea am o a.-Ii
00.00 50.00 sc. 0
:=FIGURE 7%19. DATA-INFERREDj AND PREDICTED SMOOTHSWALL HEAT TRANSFER DISTRIBUTION AROUND
NOSETIP FOR SHOT 4951 (DATA-INFERRED
TRANSITION-FRONT LOCATION)
B]
S~-148- -
' d
v" 'K
2]
17O %1951.: :11 Ttc s ~t5 OE QMOT 495: IC 19 TRA~N25 C! kz.:Z
=•
:. . . . S3
S-. -. 46C s.: s.c z:c2mc 3.O 4.Z
SHOT 49S: 1C 29 TRqN=5 OEG .C'
=II
FIGURE 7,20. RATIO OF DATA-INFERRED, TO PREDICTED SMOOTH
WALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIP
FOR SHOT 4951 (DATA-INFERRED TRANSITION-FRONT
LOCATION)
-149-
%J-- .~ -~- ---
. SHCT 5018 ICIl TRqN:10 DEG K=.TC; H_ 5 01 6 s C1E R19 'Q4=lC JE K=.00
- CNPUTED -L.I - c .-ur,W RE. .rrq[, .. X 3472 :h~rtC[ 'ý;;
; Z 1
A-rL: .=•EG) 3C"'INLEi0
,iC""D -L61
= "1
FIGURE 7.21. DAT'A-INFERRED, AND PREDICTED SMOOTHWALL HEAT TRANSFER DISTRIBUTION AROUNDNOSETIP FOR SHOT 5018 (DATA-INFERRED
TRANSITION-FRONT LOCATION)
-150-
I-I
!
V t i" i T-ýg I=k 1 - s 3! S
II
I SH 5018 1C29 •l4=IC G .
:. 2 3C. cc 4Z.o C.0 5:0Z 50.ZC
FIGURE 7.22. RATIO OF DATA-INFERRED, TO PREDICTED SMOOTHWALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIP
FOR SHOT 5018 (DATA-INFERRED TRANSITION-FRONT
LOCATI ON)
-151-
S. ýC-7 4871, 1Ci j 44
CI"U 'LUN
I :1
-- r *•ý Mf "u0[ --" i
"A .• L ,j
I-
[2]
FIGURE 7.23. DATA-INFERRED, AND PREDICTED SMOOTH
WALL HEAT TRANSFER DISTRIBUTION AROUND
NOSETIP FOR SHOT 4871 (DATA-INIFERRED
TRANSITION-FRONT LOCATION)
-152-
r
m Sir'"T '487 Tit4N%4 :EC 0
J.' 3^0 0 .n s. SO
.. tSHOT 48': :C29 TRqN:5 DEG k.=.0
a l n
0C !6,00 20 - 30C 46.0.C.O cx bo. -c
FIGURE 7.24, RATIO OF DATA-INFERRED, TO PREDICTED SMOOTH
WALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIP
FOR SHOT 4871 (DATA-INFERRED TRANSITION-FRONT
LOCATION)
-163-
"= - Z--2 - -:~L .-
- -- JCO .-
Li i
I
F FIGURE 7 25. DATA-INFERRED, AND PREDICTED SMOOTH
WALL HEAT TRANSFER DISTRIBUTION AROUND
NOSETIP FOR SHOT 4880 (DATA-INFERRED
TRANSITION-FRONT LOCATION)
-154-
..- ' _a
SHCT 488C k.:.SC C I1 TRe,- DE r,.00
... 3 -- NC..
* **g_*_t* U*=
K.4
SwC =eC '12_ T =-,!'E
FOR SHOT 4880 (DATA-INFERRED TRANSITION-FRONT
LOCATION)
-155-
--T
2C~i~2TRNIIO-RN LOCATION)
x156- .,,- -
1:
.1
FIGURE 7 27. DATA-INFERRED, ANL• PREDICTED SMOOTHWALL HEAT TRANSFER DISTRIBUTION AROUNDNOSETIP FOR SHOT 149514 (DATA-INFERRED
TRANSITION-FRONT LOCATION)
-156-
S -
•4c-.. ca.t- 3 - •
SHý 42S34 TUC~ TAN&S
,... . .. . . . . . . . ... . .. .. ...
FIGURE 7.28. RATIO OF DATA-INFERRED, TO PREDICTED SMOOTH
WALL HEAT TRANSFER COEFFICIENTS AROUND NoSETIP
FOR SHOT 4954 (DATA-INFERRED TRANSITION-FRONTLOCATION)
- ,..- ,
SH| ,, SWO !am'. -M
[7
-4 1
Z 7
i.1
ZZ 2C. 30: sze 'R4l c.
-" -.
FIGURE 7.29. DATA-INFERRED, AND PREDICTED SMOOTH
WALL HEAT TRANSFER DISTRIBUTION AROUNDNoSETIP FOR SHOT 5068 (DATA-INFERREDTRANSITION-FRONT LOCATION)
"-158-
.- ... . .u .
P m ..... mm.. ..... . . .. "
* mm
-- . . . . . . . ..
FIGURE 7.30, RATIO OF DATA-INFERRED, TO PREDICTED SMOOTHWALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIPFOR SHOT 5068 (DATA-INFERRED TRANSITION-FRONTLOCATION)
-159-
81 I I
* - - - -.. .. - -- -
~ .4
smo -M C1Ia E MTS9 T- R~ýD5K-C
-5-i
0",!
SK-OS !CS -r2 OEI k- 0
[A
ji"l
34TSC SOGSý 1C26LU :3DE
• -- 4 M T • : 9 'il q = ' • 3 K .,,
FIGURE 7.31, DATA-INFERRED, AND PREDICTED SMOOTHWALL HEAT TRANSFER DISTRIBUTION AROUND
NOSETIP FOR SHOT 5069 (DATA-INFERRED
TRANSITION-FRONT LOCATION)
-160-
- - - ~ - -::•?I llZ7 . C .7
se. s ots 506 t C1 cqkaL, on :* SpqeT 5060 IC19 ?TWO7 XG ka.0011
5--
.• .. .- .... ....
-=.. . .- .. . . ...........= ;; ;:E
r. *.** s**
FIGURE 7,32. RATIO OF DATA-INFERRED, TO PREDICTED SMOOTHWALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIP
FOR SHOT 5069 (DATA-INFERRED TRANSITION-FRONT
LOCATION)
-161-
•i ... . .- ! .... .I ," . .. ... _ . .. .,. ._ • . ,.: . . .... . . ..... . .- •- -. - 7.." . - --, • ,. . .- - -- -.. .• ' .. . . ...• • • " ,... .. . - • l
TABLE 7,2. HEAT TRANSFER AUGMENTATION FACTORS
. - Heat Transfer Augnmentation Factor
Shot # Material ICC (data/QC-s) * j(0dta C-S)transition 40'
4882 W 19 1.23 1.0529 1.32 1.16
4963 W 11 Laminar Laminar19 Laminar Laminar29 Laminar Laminar
4953 ATJ-S 11 1.06 1.0719 1.16 1.1229 1.20 1.20
4974 ATJ-S 11 1.10 .9319 .95 .9S 29 .98 1.0
4951 994-2 11 1.11 1.0419 1.28 1.2
29 1.42 1.45018 994-2 11 .96 .9
19 1.0 .98529 1.10 .9
4871 CC-223 11 1.13 1.0319 .913 1.91
4880 CC-223 11 1.120 .1419 1.04 .9529 .99 .98
4954 CC-223 11 1.03 1.0419 .94 .9129 .81 .8
S5068 CC-2,23 11 1.1i0 .9519 1.05 .9829 1. 17 1. 1
S5069 CC-223 11 1. 20 1. 0
19 1.25 1.029 1.28 1.1• 0 data = data inferred heat transfer
QC-S= computed smooth wall heat transfer
-162-
. i l i '.>- / - ',F . .. . . . . . . ,-
factors (or ratios), but most of the data indicated small heattransfer augmentation ratios t 1.20). This was expected sincefor the surfice roughness and flight conditions tested, the
+roughness Reynolds number, k , were mainly in the transitionally
rough regime (k+ < 70), as indicated by Table 7.3. The calcula-+tions indicated that not only were k values small, but that
the flow was mainly transitional i.e., the boundary layer wasnot fully turbulent. (The definition of the roughness regime
here is not exact, since it is only applied to fully turbulent
flows and the flow here was transitional.)
7.2.2 Comparisons with the ASCC Turbulent Heating
Augmentation Methodology
In this section a comparison is made between the data
inferred heat distribution around the nosetip, and the roughwall computed values using the rough wall augmentation metho--
dology of the ASC code. The purpose here was to verify e xtrapo-lating the ASCC analytical rough wall augmentation model to the
ballistics range flight environment.
Figures 7.33 to 7.42 show the distributions of the heattransfer coefficient ratios around the nosetip. These are the
ratios of data inferred to the ASCC computed rough waZZ heat
transfer. A value of unity to the heat transfer coefficient
ratio represent an exact agreement between the data and theory.In computing the heat transfer coefficient around the nosetip
at each ICC location the transition-front location was set
equal to the data inferred value, and the surface temperaturedistribution was set equal to the measured data. It was found
that, in general, the computed rough wall heat transfer coef-
ficients were higher than the data-inferred values. The calcula-tions were at the most 25% higher than the da,1 a.
-163-
TABLE 7.3, MAXIMUM ROUGHNESS REYNOLDS NUMBER
ON THE NOSETIP (COMPUTED BY ASCC)
Shot # k ICC k+mean
4882 .292 19 7329 57
4953 .55 11 7119 5529 44
4974 .55 11 9719 8429 63
4951 .57 11 9019 6329 48
5018 .57 11 liE19 10029 80
4871 .25 11 3019 20
4880 .25 11 23A19 2229 19
4954 .25 11 2019 2029 20
5068 .25 11 3319 2729 20
5069 .25 31 4219 34
S_ .. l 29 27
tijoto i - RoughnfoD noynoldo 1'3r'bor
Table 7.4 lists a detailed comparison between the data
and ASCC predictions at two points on the nosetip. These are
the transition and the 400 point. The details at all the
points on the nosetip can be found in Figures 7.33 to 7.42. At
the 400 point on the nosetip the following is listed.
1. Qdata/Ascc fixed transition/rough wall: transi-tion-front locations we:e set equal to the data inferred
values, and the rough wall turbulent augmentation was included
in the heat transfer calculations. This ratio indicates the
degree of agreement between data and rough wall theory. It can
be seen in the table that for most of the CC-223 shots the
agreements are within 10%. For the graphite shots the agree-
ments are within 20%. That is, for all the flight conditions
tested, the roughness effects on enhancing the heat transfer
were small, as expected for the calculated values of kt.I
2. Qdata/QASCC fixed transition/smooth wall: again
the locations of transition in the calculations were set equalto the data inferred values but QASCC were smooth wall values.
This ratio represents one of the following: (a) heat transfer
dugmentation values due to roughness, (b) degree of agreement
between data and laminar theory if roughness effects are small.The heat transfer ratios indicated that roughness effects were
small, and fairly good agreement between data and laminar
theory was found.
3. 0 data/AsQc PANT transition/rough wall: QASCC
was based on a transition location predicted by the PANT tran-
sition criteria. The heat transfer coefficient ratios become
large as indicated in Table 7.4 when the theory predicts lami-
nar flow or transition offset, while the data infer a transi-
tional/ turbulent flow. This was the case for the CC-223 Shots
4871, 4880, 4954 ind 5068 (ICC's 29, 41). Figure 7.43 shows
ratios of data derived t( jCC calculated rough wall heat trans-Fri- "naficiarta. The tr ... ition front location was predicted
TABLE 7.4. RATIO OF DATA DERIVED TO COMPUTEDHEAT TRANSFER RATES
"aet Transfer Coefflicients Natio
(0dats"OAlAW) (Qftt/ONM )
Tranition 40 DegreesShot I material ICC Comamets
PANT• Fixed PABIT TramnLi.ionto TransitiinTransition Transition Smooth Nog
mg mall mail wall
4382 W 19 "15 1.23 1.0 1.05 .9 * AS= (PAINT) predicts tranmition bw29 1.66 1.32 1.95 1.16 1.0 8-108.
a Data derived transitiom is 7.
4963 W 12 Laminer Laminar Laminar Laminar Laminar 0 Lasminar-good aemnt. between data and19 Lami.ar Laminar Laminar Lamina Lamlnar theory.29 Laminae Z, oinar Laftnar Laminar Lmnar
4953 ATJ*-8 11 1.17 1.06 .9 1.07 .92 . AS= (PANT) predicts tramaition at19 1.35 1.16 .97 1.12 .97 approximately 10-190.29 1.45 1.20 .08 1.2 1.07 e Data ladicate transition at approxiatel
3-51.* Turbulent agreement betwee data and
theory for *8CC with fixed traneitionfairly good.
4974 ATJ-8 11 1.1 1.10 1.0 .93 .8 9 AS=C (PANT) i tZ transition at19 .9 .95 .9 .9 .0 awff Wimetely 12-19%29 1.0 .93 .1 1.0 .9 a Data Indioate transitiom 5-17'.
for smooth wall case (fixed transition).
kough well calculations o*erpraidotsdate.
4951 994-2 11 1.15 1.11 .9 1.0 .93 9 ASCC jPANT) predicts transition betwee19 1.36 1.29 1.05 1.2 1.2 10-15.29 1.5 1.42 1.2 1.4 1.4 a Data derived transition is S3.
Good data theory agreement at IC 11(fixed transition) for smooth wall.
Sot)O smooth and rough wall under-predicted at IC 19 and IC 29.
5018 994-2 11 .95 .96 .75 .9 .12 a AS )PART predicts transition between
19 .96 .96 .7.5 .O 10-13 aa alo the data.29 1.1 1.10 .00 .9 .85 . Dot" smooth mad rough wall calculations
(fixed transition) overpredscts Gate.
4871 CC-223 11 Large 1.13 Large 1.03 .95 9 AWC (PANT) predicts lmainar flow29 Large 1.13 Large 1.11 1.08 through entire flight.
a Data indicates transition at approxi-mately 8 - ,'1.
9 Good ag.q.emnt between data and theoryfor •8CC with fixed transition (bothsmooth and rough wall).
4000 CC-223 11 Large 1.20 Large 1.14 1.1 a A8CC (PANT) indicates laminar flow18 Large 1.04 Large .95 .9 through entire flight.29 Large a9 Lrge .98 .97 a Data indicates turbulent flow.
a good agrMeeant between data and theoryfor fine transition.
4954 CC-223 11 Larpe 1.03 Large 1.04 1.0 a *8CC (PANT) indicates laminar flow19 large .94 Large .91 .9 through the entire flight.29 Large .81 Large .8 .77 0 Good agreement betwe data ad fixed
trastion UMC for IC 11. Theory over-predicts (smooth and rough) at IC 19-29.
5068 CC-223 11 1.45 1,10 1.05 .99 .9 a 8CC IPANT) Widiealtg transition at19 1.67 1.05 1.1 ,o .9 approximately 23-30'.29 Large 1.17 Large 1.1 1.07 1 Data Indicte transition at approxi-
mately 3-5'
SPART t:ansition criteria predicts trmnsi-tion offset at ICC 29.
O Good agreement between data and theoryfor AeCC with fixed ttansition andsmooth wall.
sof9 CC-323 I .1.59 1.20 1.0 1.0 .9 a AACC (PANR) predtos trans itin att9 1.96 1.25 1.05 10 .I oppronimately17-2)',29 1.9 1.21 I.s 1,1 1.0: 0 Date inticate traniLtion at approxi-
mately 4".G veod turbulent agreement between data"wd theory for *8CC with fixed tran-sition and smooth wall,
*Transition- front location was predicted by PANT correlation
tTransition- front location was equal to the data-derived value
.-. . .
FIGUR 7.3 RAI OF DAAIFERD ToA RDCE OG
- S0 -
FO SHO 488 (FU-NFRE TASITONF
*. . . . .
-= I
-- .
S S
FIGURE 7,33. IRATIO OF DATA-INFERRED., TO ASCC PREDICTED ROUGHWALL HEAT TRANSFER CCE.FFICIENTS AROUND NOSETIPFOR SHOT 4882 (FLux-INFERRED FRANSITION-FRONT
. . .. -.- .• ...
. . . . . . . . . . . . . . . . . . . . . ......
= - ,,a
--. .._
FIGURE 7.34, RATIO OF DATA-INFERRED, To ASCC PREDICTED ROUGHWALL HEAT TRANSFER COEFFICIENTS AROUND NOsETIP
FOR SHOT 14953 (FLUX-INFERRED TRANSITION-FRONTLOCATION)
. ...... . ........ . . . .. . .
. ..- ......
= C=
FIGURE 7-35. RATIO OF DATA-INFERRED, To ASCC PREDICTED ROUGHWALL HEAT TRANSFER COEFFICIENTS AROUND NoSETIPFOR SHOT 4974 (FLUX-INFERRED TRANSITION-FRONT
LOCATION)
-169-
I
= =-
- -- -
FIGURE 7.36, RATIO OF DATA-INFERRED, To ASCC PREDICTED ROUGHWALL HEAT TRANSFER COEFFICIENTS AROUND NoSETIP
FOR SHOT ~4951 (FLUX-INFERRED TRANSITION-FRONTLOCATION)
I :
-~N
- * - '-Ea:
I'IURE7,7. ATI-O .,,--?IFRRD ToA PEITE OG
-171
4 . .
I1GURE 7,37, RATIO OF DATA-INFERRED, TO ASCC PREDICTED ROUGHWALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIPFOR SHOT 5018 (FLUX-INFERRED TRANSITION-FRONTL.OCATION)
-171-
I. .. .. . 483 , •,: -- 45 --
, 9•CT 7.•7" :.:2£ •c;,:5 4:2. c
FIGURE 7.38. RATIO oF DATA-INFERREDi To ASCC PREDICTE ROUGHI
WI'ALL HEAT TRANSFER COEFFICIENTS AROUND
4OTI
FOR SHOT 4871. (FLUX-INFERRED TRANSITION-FRONT
LOCATION)-172-
FGRE 7"9 AI FDT-NERD oAC RDCE OG
1-7
•CC c,--_' ,Er.--.'3 -•, : :.-. -E
NN
----
.--
FIGURE 7.39, RATIO OF DATA-INFERRED., TO ASCC PREDICTED ROUGHWALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIP
FOR SHOT 14880 (FLuX-INFERRED TRANSITION-FRONT
LOCATION)
-173-
2."', ,
. . . . .
3'* ,;5
FIGURE 7.40, RATIO OF DATA-INFERRED, To ASCC PREDICTED RoUGHWALL HEAT TRANSFER COEFFICIENTS AROUND MOSETIP
FOR SHOT 4954 (FLUX-INFERRED TRANSITION-FRONTLOCAT ION)
-174-
eN N
LiiS'Q 5:3 :cz -Rclz '
II
c ' I I
- I
FIGURE 7,41. RATIO OF DATA-INFERRED, To ASCC PREDICTED ROUGHWALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIP
FOR SHOT 5068 (FLUX-INFERRED TRANSITION-FRONT
LOCATION)
-175-
- ,--
-
.. . . ._
2-
FIGURE 7,42, RATIO OF DATA-INFERRED, To ASCC PREDICTED ROUGH
-:. WALL HEAT TRANSFER COEFFICIENTS AROUND NOSETIPS~FOR SHOT 5069 (FLUX-INFERRED TRANSITION-FRONT
LOCATION)
.. "-176-
r:
C)
0,0
a- U')zm
S -C)U)CC
LA SJ
z w z
0) 0 co(N0Lic 3m LL,
(D 0
S ~ 0C)0
U CL.
U IUU Cn
00 E GL"I Os I SI 001l SL 0 os- 0HI alluUiflJ1/H3 O1I~NN1N HIHO
C)at C)
m 0j
* 0
U- w
OO0 6- 19 0
C7) C))C-, LL. 0
by the PANT correlation. The plots show high heat transfer
r-atios in the transitional region due to inconsistency in
transition location between data and theory.
Similarly, the heat transfer ratios are listed in the
table for the case when transition is given by the PANT correlation
and transition point set equal to the data inferred values.
Detailed comments are also listed in the table.
-178-
8.0 SUMMARY AND CONCLUSIONS
Transition and heat transfer tests were performed in the
A.DC hyperballistics Track-G test facility. The objectives of
the tests were:
1) Investigate the effects of surface roughnesson nosetip transition.
2) Investigate the effects of surface roughnesson heating augmentation.
3) Verify existing simulation methodologies asto their applicability to flight environments.
'ITungsten, graphitic and carbon-carbon models were
tested. From a flight simulation point of view, the test condi-
tions were selected to simulate altitudes where transition andheating augmentation influence the shape development of nose-tips. From a testing point of view, the test environments were
selected to insure nosetip transition and to obtain surface
temperatures that are within the dynamic response range of theimage converter cameras used to record the nosetip image. Thiswas mainly to ensure more reliable data. The test freestream
conditions covered a range of stagnation pressure from approxi-
mately 20 to 195 atmospheres and freestream unit Reynoldsnumber from 17 to 70 million/ft.
The surface of the tested models were first precondi-tioned and then characterized. The tungsten models were eithersuper-polished to obtain a super-smooth surface finish (Krms1 10 pinches - 0.01 mils), or preroughened to obtain a uniform
surface roughness. The graphite and carbon-carbon models werepreablated in the Aerotherm low pressure-high enthalpy arc jet
facility, in order to create surface roughnesses similar to thosethat may develop in a laminar ablation environment prior to
nosetip transition. Representative models of each materialwere then characterized to obtain the surface roughness height
statistical distributions. Mean and median roughness heightswere then obtained from the roughness distributions.
-179-
The temperature data reduction and handling were automated.
Temperature data on the entire nosetip were obtained at 3 degree
increments along and around the nosetip. The brightness tempera-
ture data were recorded on a magnetic tape, which was then pro-
cessed by computer. For each shot and at each image converter
camera location, 120 temperature profiles along the body were
obtained. The mean, maximum, and minimum temperatures, as well
as standa-:d deviations were then obtained utilizing these 120
profiles.
Sensitivity studies were performed according to Ref. 3 to
estimate the level of uncertainties in the measured surface
temperature due to surface spectral emissivity, shock layer
radiation, surface reactions, motion blur, and gun barrel heating.
It is believed that the measured brightness and surface temperatures
are accurate to ± 100 to ± 200 degrees Kelvin.
The nosetip transition locations were inferred in three
ways.I!1. At each ICC location the 120 temperature profiles werenumerically processed and the transition was taken to be the
point of minimum temperature, i.e., the point of zero temperature
slope.
2. The 120 temperature profiles were also maniually processed,
since some engineering judgement was required to determine the
transition location, defined as the intersection of the tangents
to the laminar and transitional region of the profile.
3. Using the mean temperature profiles at all the ICC
locations, the mean heat flux distributions were inferred using a
heat conduction code. The transition location was then inferred
as the point of intercept of the tangents to the laminar and
transitional leg of this mean heat flux curve.
The first approach, though rendering a statistical distribution
of the transition front location quice conveniently by the computer,
produces results that are biased by the numerical smoothing of
the data. The second approach, though tedious, permits a more
realistic determination of the transition location. However,
inferring transition location from the temperature profiles may
-180-
lead to errors due to the 3moothing out of the nosetip, temperature
profiles because of lateral heat conduction, and, more important,
due to the transient nature of the tests in the ballistics :-ange,
the measured temperature profile shape may be completely different
from the corresponding heat transfer distribution. Human judgement
must be applied here. However, analyzing 600 temperature profiles
per shot is a very tedious task, and an automated system using
interactive graphics would be helpful.
Inferring the transition locations from the heat flux distri-
butions allows better definition of the transition-fronts.
However, utilizing the mean-temperature histories in deriving the
heat flux distribution elements obscures any statistical information
on the transition front, due to the averaging process of the
surface temperatures around the nosetip at each ICC location.
The results of the transition analyses showed that:
1. For all the tested materials transition was observedon the nosetip and no transition offset occured. Afully laminar distribution was achieved only on thesuper smooth tungsten model; otherwise, transitionoccured in the stagnation region even for nominallysmooth tungsten models.
2. The data indicated transition locations consider-ably upstream of the PANT transition correlationusing the mean roughness height, kmen
3. For certain flight-surface conditions the dataindicated transitional flow yet the PANT transi-tion correlation predicted either laminar flowor transition offset.
4. For the test conditions analyzed the PANT transit-tion correlation predictions were upstream of thosegiven by Dirling's and Bishop's transition models.
5. The inferred transition-front locations werequite scattered when compared with the Reda-Raper transition location-range pressurecorrelation.
6. The inferred roughness heights, that is, roughnessheights to produce agreement with the PANT predic-tions, were of the order of the roughness associ-ated with the matrix or weave structure for CC-223and considerably higher than the k menfor theother tested materials,.ea
Stagnation point heat transfer augmentation factorswere derived at each ICC location. The data inferred heattransfer rates were then compared with laminar theory predic-tions. To eliminate any wall temperature effects, the com-parisons were performed for the same wall temperatures as the
data. The derived augmentation factors were then compared withthe predictions of the PANT and Phinneys laminar augmentation
correlations. The stagnation point results can be summarized
as follows:
1. Stagnation point augmentation factors wereobserved in the ballistic range environments.
2. The data were scattered but a 20% excess inheat transfer over and above the laminartheory exists.
3. The PANT laminar augmentation correlationdid not predict augmentations for the flightregime considered.
4. The Phinneys augmentation correlation over-predicted the observed factors.
The data-derived heat transfer distributions around thenosetip were then compared with smooth wall predictions, (inorder to infer rough wall transitional/turbulent heating &ug-mentation levels) and rough wall predictions (to verify theASCC rough wall methodology). The comparisons were performedfor two situations, where the transition-front locations wereeither set equal to the data inferred values, or predicted bythe PANT transition criteria.
The heat transfer results off the stagnation pointshowed that:
1. Comparisons between the data derived heat fluxesand the predicted smooth waZZ values, indicated that roughnesseffects are small (= 20%). Although, the transition-front
locations in the calculations were set equal to the data in-ferred values, it is not certain whether the predicted transi-tion region width is equal to that on the model in flight.
-182-
2. Comparisons between the data derived heat fluxes
and the ASCC predicted rough waZZ values, also indicated that
roughness effects are small. The predictions showed relatively
small values of the roughness Reynolds numbers, k+, and that
the boundary layer is mainly transitional. For certain flow
situations the ASCC-rough wall uverpredicted the data by 25%.
3. In verifying the heat transfer prediction models
off the stagnation point, it is essential that the transition
front location be consistent between data and theory. For
certain test conditions the experimentally observed heat flux
showed that the flow was turbulent, while the predictions
indicated a laminar or relaminarized flow. In this case one
is taking the ratio of a turbulent flux to a laminar one, and
the ratio becomes larger, as expected and shown in Section 7.
4. Obtaining the heat transfer augmentation factors by
simply taking the ratio of the rough-wall data-inferred heat
flux to the smooth wall predicted value at the same streamwise
lccation on the nosetip, is misleading. This ratio can assume
a.•y value since the rough wall boundary layer is quite different
from that of a smooth wall, hence the boundary layer parameters
are different in the two cases. One has to correlate the
rough wall heat transfer (and not the augmentation factors)
with the rough wall boundary layer parameters e.g., k/e, k,
h~he•Based on the above discussion one can conclude the
following:
1. The use of in-flight surface temperature to verify
nosetip aerothermal gredictive methodologies requires complex
data reduction, due to the high coupling between transition,
heating augmentation and roughness effects.
2. A detailed look at the individual temperature pro-
files along the nosetip at each ICC location is useful for
defining transition front locations. An interactive
-183
graphics system is essential here for the data reduction
process. The results however can be very misleading (4 below).
3.Inferring the transition location from the data-
derived heat fl.ux eliminates any statistical behavior of the
transition front due; to the sm~oothing of the temperature
variations affected by the averaging process of the data.
4. It is essential to infer the transition-front
locations from the derived heat transfer data. Inferring transi-
tion only from the temperature data can be seriously in error
in situations where transition offset takes place. Due to the
transient nature of the model response the temperature data may
indicate a transitional/turbulent distribution but in realityF the heat flux indicates a laminar distribution. The temperature
data can therefore indicate, erroneously, a forward movement of
the transition-front. However, the temperature data can predict
transition locations correctly at the early ICC locations (ICC
4 and 11) and give information on statistical distribution of
the front.
F ~The approach then should be that transition be infe-:zed
from heat flux data and further refined and checked by inf~orma-
tion obtained from the temperature data. One can not rely only
on the temperature data. Heat flux calculations should then be
performed and compared with heat flux data for both laminar and
turbulent situations in order to insure whether or rnot transition
offset took place.
5. Transition-front locations were considerably up-
stream of the PANT, Dinling's and Bishop's predictions. Alsoc,
no relaminarization was observed. The PANT correlation pre-
dicted transition upstream of those of Dirling's and Bishop's,
respectively.
-184-
6. Due to the transient nature of the problem, predic-
tive models verifications by comparison between the measured andpredicted surface temperature at one ICC location can be verymisleading and markedly in error. Comparisons must be performedat aZl ICC locations. This approach however, is tedious due to
the strong coupling between transition, heat transfer and surfaceroughness effects.
7. While higher roughness levels might seem indicatedfrom the transition data, larger roughness hei, .s will leadto higher heat transfer rates. This may result in poorer agree-
ment between data and theory for the heat transfer. It seemslikely that roughness heights that correlate with transitiondata may not be those correlating with heating augmentation
data.
8. The hyperballistics range track facility is a uniqueground test facility to simulate a wide range of flight environ-
ments. Flow uncertainties that exist in wind tunnel and arc
jet facilities, e.g., pressure fluctuations, freestream tur-bulence, are absent. Other track related problems appear tohave been resolved.
I
i -185-
9.0 RECOMMENDATIONS
It seems clear that the ballistics range will become
an increasingly important facility in the study of aerothermalprocesses on hypervelocity vehicles and in the development of
new materials and concepts. Recommendations for utilization
and near-term improvements of the ballistics range are summarized
below.
1. Continue using the guided track mode of testing ininvestigating transition characteristics and heat transfer on
nosetips.
2. Test well and better characterized materials (W,ATJ-S, 994-2) to understand the basic transition phenomena
before investigating more complicated materials (FWPF, CC).
I 3. Reduce ballistic range data to key boundary layer
and material response parameters necessary to more carefullyassess the events and shortcomings of competing transition andrough wall heating techniques.
4. Use the developed ballistics range test techniquesto investigate transition and heat transfer of non-spherical
configurations, e.g., ellipsoids, nosetips at angle of attackand aerodynamic control surfaces.
5. Expand the materials performance characterizations
begun in this program to encompass all candidate advanced nose-tip material characterization worK, transition modeling studies,nosetip shape change modeling work, and other related experimental
programs (CALSPAN, Tunnel F).
6. Make ballistics range testing, using the technologydeveloped under these (and other) programs, an integral part ofthe pre-flight evaluation process for new materials and concepts,
to complement the testing standardly carried out in the AFFDL
50 MW facility.
-186-
7. Add high sensitivity image converter cameras at
uprange stations to improve data reduction accuracies. Reduc-tion of measured temperatures to heating rates is absolutelyessential if any reliable conclusions are to be drawn concern-ing transition and rough wall heating. Addition of at least onecamera would greatly enhance data conZidence.
8. Develop an interactive-graphics system at AEDCwhich enables the detailed investigation of the individual
temperature profiles on the nosetip at each ICC location (600total per shot). This system iv essential for investigation
of nosetip transition asymmetry and possibly gouge formation.IIft
I"
' -1.87-
r w -.-- ----. -
10.0 REFERENCES
1. Holden, M. S., "Studies of Transitional Flow, UnsteadySeparation Phenomena and Particle Induced AugmentationHeating on Ablated Nosetips," AFOSR-TR-76-1066, Oct.1975.
2. Roda, D. C., Leverance, R. A., "Boundary-Layer Transi-tion Experiments on Pre-Ablated Graphite Nosetips ina Hyperballistics Range," AIAA Paper 76-356, Presentedat the AIAA 9th Fluid and Plasma Dynamics Conference,San Diego, California, July 14-16, 1976.
3. Shih, W. C. L., Wassel, A. T., Kessler, A. J., Courtney,J. F., "Evaluation of Reentry Vehicle Nosetip Transi-tion and Heat Transfer in the Hyperballistics Range,"SAI-79-541-LA, July, 1978.
4. Norfleet, G. D., Hendrix, R. E., "Development of aHypervelocity Track Facility at AEDC," AIAA paper 77-151, Presented at AIAA 15th Aerospace Sciences Meeting,SLos Angeles, Calif., January 24-16, 1977.
5. Dyner, H. B., "Test Facilities Used for Advanced Bal-listic Reentry Systems R and D," Aerospace Report No.TOR-0077(2550-71)-2, 20 June 1977.
6. Wool, M. R., et al., "Interim Report - Passive NosetipTechnology (PANT) Program - Computer User's Manual SAANTProgram (U)," SAMSO-TR-74-86, Vol. 7, 1974.
7. Fogaroli, R. P., Brant, D. N., "Re-Evaluation of GraphiteThermo-Chemical Ablation," GE Thermodynamics Laboratory,Fundamentals Memo TRM-9151-060, October 1968.
8. Kratch, K. M., et al., "Graphite Ablation in High-PressureEnvironments," AIAA Paper No. 68-1153, December 1968.
9. Dirling, R. B., Jr., "A Method for Computing Rough WallHeat Transfer Rates on Reentry Nosetips," AIAA Paper No.73-763, AIAA 8th Thermophysics Conference, Palm Springs,California, July 1973.
10. R. B. Dirling, Jr., D. A. Eitman and J. D. Binder,"Evaluation of Post-Test Ablation Models," AFML TR-77-225, Dec. 1977.
11. Dugger, P. H., Bock, 0. H., Enis, C. P., Gilley, B. W.,"Photographic Pyrometry in an Aeroballistics Range,"Proceedings of the SPIE 16th Annual Technical Meeting,"San Francisco, California, October 1972.
-188-
--- •-• " %" " "; .- " . l• • • l• • - • ,•. . -. .. ,,,_ -
12. Norfleet, G. D,, Hendrix, R. E., Raper, R. M., Callens,E. E., Jr., "Development of an Aeroballistics Range Cap-ability for Testing Raentry Materials," JSR 12, 5, pp.302-307, 1975.
13. Reda, D. C., Leverance, R. A., Dorsey, W. G., Jr.,"Application of Electro-Optice.l Pyrometry to ReentryVehicle Nosetip Testing in a Hyperballistics Range,"6th International Congress on Instrumentation inAerospace Simulation Facilities, 22-24 September 1975.
14. Anderson, A. D., "Evaluation of Theoretical Methods forthe Predication of Nosetip Boundary L&yer Transition,"Aerotherm/Accurex TM-75-77.
15. Bishop, "Transition Induced by Distributed Roughness onBlunt Bodies in Supersonic Flows," SAMSO-TR-76-146, 29Oct. 1976.
16. Dirling, R.B. "The Effect of Transition and Boundary LayerDevelopment on Hypersonic Reentry Shape Change," MDACPaper WD 2507, May 1975.
17. Dahm, T. J., et al., "Passive Nosetip Technology Program,Vol. I: Inviscid Flow and Heat Transfer Modleling forReentry Vehicle Nosetips," Aerotherm Report -76-224,Oct. 1976.
18. Clever, R. M. and Denny, V. E., "Two-Dimensional Tran-sient Conduction and Shape Change Analysis," SAI-78-538-LA, March 1977.
19. Hildebrand, F. B., "Introduction to Numerical Analysis,"McGraw Hill, N. Y., 1956.
20. Reda, D. C. and Raper, R. M., "Measurements of Transition-Front Asymmetries on Large-Scale, Ablating Graphite Nose-tips in Hypersonic Flight," AIAA 79-0268, 17th AerospaceSciences Meeting, New Orleans, LA., Jan. 15-17, 1979.
-189-
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. - . . .. • . - • r-1 -•-