Collie, Katherine A

Piland, Robert O

English

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- NACA 
RESEARCH MEMORANDUM 
AERODYNAMIC HEATING O F  ROCKET-POWERED RESEARCH 
VEHICLES AT HYPERSONIC SPEEDS 
By  Robert  0. Piland and Kather ine A. Collie 
Langley Aeronautical  Labora tory  
Langley Field,  Va. 
I 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
II' J ~ Y  19, 1955 
I Declassified September 1, 1959 
https://ntrs.nasa.gov/search.jsp?R=19930088771 2020-06-17T08:45:05+00:00Z
NACA RM L57ElOc 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH WORANDIIM 
AERODYNAMIC HEATING OF ROCKET-POWERED RESEKRCH 
VEHICLES AT HYPERSONIC SPEEDS 
By Robert 0. Piland and Katherine A. Collie 
SUMMARY 
Skin temperature measurements have been obtained on two free-f l ight  
research vehicles; measurements were made up t o  Mach numbers of 5.4 i n  
one case and up t o  10.4 i n  the other. On the f i r s t  model (M = 5.4), 
temperature measurements were obtained a t  nine s tat ions (between 6.5 
and 29 inches) along the 1 5 O  conical nose. The temperature d is t r ibut ion  
along the nose, a t  a Mach number of 5.4, indicated t rans i t ion  t o  have 
occurred a t  a Reynolds nunher of about 9 X 106. The theories of Van 
Driest for  laminas and turbulent flow proved adequate f o r  predicting 
the skin temperature h is tor ies  fo r  laminar and turbulent s ta t ions,  
respectively. 
A skin temperature his tory was obtained a t  a single s t a t ion  
(26.6 inches from tip) on a fineness-ratio-? K&& nose shape t o  a 
Mach number of 10.4. The peak skin temperature of 1 , 2 0 0 ~  F occurred 
shortly a f t e r  the time of peak Mach number and calculations indicate 
tha,t at t h i s  time the convective heat t ransfer  t o  the model was being 
balanced by radiative heat t ransfer  away from the model. The theory of 
Van Driest was used i n  calculating a skin temperature and appeared 
adequate up t o  a Mach n u d e r  of about 6. Above t h i s  value, there i s  
some divergence between calculated and measured data, possibly due t o  the 
existence of other than turbulent flow over the measuring s tat ion.  
It i s  well known tha t  the problem of aerodynamic heating becomes 
more severe with increasing Mach number. In order t o  determine jus t  
what heating conditions w i l l  be encountered a t  higher Mach numbers, 
research vehicles are being f l i g h t  tested, a t  present, t o  Mach numbers 
as high as 10.4. 
The purpose of t h i s  paper is  t o  present and discuss skin temperature 
measurements from two f l igh t  t e s t s .  Temperature measurements were obtained 
t o  a Mach number of 5.4 on the f i r s t  f l i gh t  and t o  a Mach number of 10.4 
on the second f l ight .  
NACA RM L55ElOe 
Mach number 
Reynolds number 
s t a t i c  temperature, degrees Rankine 
skin temperature, degrees Fahrenheit 
length on which Reynolds number calculation i s  based 
RESULTS AND DISCUSSION 
Figure 1 presents the var iat ion of s t a t i c  temperature, Mach number, 
a l t i tude ,  and Reynolds nuniber per foot with f l i gh t  time for  a three- 
stage vehicle which had a l 5 O  t o t a l  angle conical nose. 
F i w r e  2 presents the variation of skin temperature with time meas- 
ured a t  a distance of 29 inches from the nose t ip .  The nose cone was 
constructed of 0.027-inch-thick Inconel. The maximum measured tempera- 
tu re  is  seen t o  be 1 , 9 0 0 ~  F, a t  which time (23.7 seconds ) the vehicle 
i s  a t  a Mach n u d e r  of 3.4. 
A temperature his tory has been calculated using the turbulent theory 
of Van Driest (ref.  1) applying a conversion factor t o  make it applicable 
t o  the cone (ref .  2). Turbulent flow was assumed t o  ex i s t  from the nose 
t ip ,  md the recovery factor  used i n  calculating the adiabatic wal l  
temperature was assumed equal t o  the cube root of the Prandtl number Pr 1/3 
based on skin temperature (ref .  2). The r a t i o  of specific heats y was 
taken equal t o  1.4. The calculo.ted temperature his tory i s  indicated by 
the symbols i n  figure 2 and is seen t o  be i n  good agreement with the  
measured temperature history. 
In  addition t o  the temperatures measured a t  s ta t ion  29, measurements 
were also made a t  e ight  other s ta t ions along the same 150 cone. Figure 3 
presents the var iat ion of skin temperature along the nose cone a t  a par- 
t i c u l a r  time during the f l ight .  The f l i g h t  time considered i s  23.7 sec- 
onds a t  which time the Mach number is 5.4 and the Reynolds number per 
foot i s  9.8 X 106. The measured temperature i s  indicated by the so l id  
l ine  and the t.icks on t h i s  curve and along the center l ine  of the cone 
indicate the s tat ions a t  which the temperature measurements were made. 
NACA RM L55ElOc 
The shape of the measured curve indicates regions of laminar, t ransi-  
t ional ,  and a t  l eas t  the ear ly  stages of turbulent flow. The abrupt 
change i n  slope of the temperature curve indicates tha t  t rans i t ion  from 
laminar flow began a t  a Reynolds number of about 9 X 106. It i s  in ter -  
es t ing t o  note tha t ,  i n  a f l i g h t  time of only 23.7 seconds, a difference 
of YO0 F exis t s  between laminar-station and turbulent-station tempera- 
ture  measurements . 
The lower broken l ine  i n  figure 3 indicates temperatures calculated 
using the laminar theory of V a n  Driest ( re f .  3 ) )  applying a conversion 
factor t o  make it applicable t o  the cone (ref .  2). The recovery fac tor  
was assumed equal t o  the square root of the Prandtl nuplber (ref.  2). 
Fa i r  agreement is  seen t o  ex i s t  between the measured and calculated da ta  
a t  the s ta t ions where the laminar flow is believed t o  ex is t .  
The upper broken l ine  i n  figure 3 indicates temperatures calculated 
using the turbulent theory of Van Driest as described previously. Toward 
the a f t  end of the cone (s ta t ion  29)) where it i s  believed turbulent flow 
exis ts ,  the agreement between the measured and calculated data  i s  quite 
good. 
1,et us now consider temperature measurements t o  a Mach number of 10.4 
obtained during the f l i g h t  t e s t  of a four-stage vehicle, which is des- 
cribed i n  reference 4. 
Figure 4 presents the var iat ion of s t a t i c  temperature, Mach number, 
a l t i tude,  and Reynolds number a t  the temperature measuring s tat ion,  f o r  
the t e s t  vehicle. The nose shape on which the temperature measurements 
were made was a fineness-ratio-? K&& nose shape whose t i p  was modi- 
f ied  t o  form a l 5 O  t o t a l  angle wedge. The nose skin was of 0.032-inch- 
thick Inconel. A sketch of the nose shape i s  presented i n  figure 5 and 
the temperature measuring s t a t ion  i s  shown t o  be 26.6 .inches a f t  of the 
nose t i p .  The variat ion of skin temperature with time is also presented 
i n  figure 3 and i s  indicated by the sol id  l ine.  A peak temperature 
of 1,2000 F is  seen t o  have been reached just  a f t e r  peak Mach nuuiber. 
A t  t h i s  time, the heating potent ia l  i s  s t i l l  considerably higher than 
the skin temperature, and considering only convective heat t ransfer  a 
peak temperature would not have been expected u n t i l  a l a t e r  f l i g h t  time. 
The explanation f o r  the peak a t  t h i s  time is  tha t  the convective heat 
t r m s f e r  t o  the skin i s  being balanced by the radiative heat t ransfer  
away from the skin. Keeping t h i s  i n  mind, consider the temperature his- 
tory calculated using the turbulent theory of Van Driest ( ref .  1) f o r  a 
cone end including radiative heat transfer.  The calculated values (square 
symbols) agree quite well  t o  about 28 seconds, a t i m e  which corresponds 
t o  a Mach number of about 6. A t  t h i s  point the measured and calculated 
values diverge and a t  the time of peak measured temperature the calculated 
c w ~ e  i s  s t i l l  r i s i n g  rapidly,  However, i f  t r a m i t i o n  t o  laminay flow 
i s  a r b i t r a r i l y  assumed t o  occur a t  a Reynolds n u ~ b e r  of 5 X 106 and 
laminar heat- t ransfer  coeff ic ients  are used, the  points indicated by 
the  so l i d  square symbols are obtained. The inference i s  t h a t  laminar 
flow must have actual ly  exis ted on the  model a t  t h i s  time of peak temper- 
~ t u r e .  Calculations are a lso  presented considering f l a t -p l a t e  theory; 
these a re  indicated by the c i rcu la r  symbgls. A t  the  higher temperatures, 
the  agreement between the  f l a t -p la te  calcula t ions  and the  measured da ta  
i: b e t t e r  than t he  calculations fo r  the  cone; t h i s  i s  probably fo r tu i tous  
and may be due t o  the  existence of t ransi t iona, l  r a ther  than turbulent  
flow over the  measuring s t a t i o n  fo r  some time in t e rva l  p r i o r  t o  time of 
peak temperature. The primary purpose of making the  f l a t -p l a t e  calcu- 
l a t ions  was t o  determine the  difference i n  estimated maximum temperature 
occasioned by using the  simpler f l a t -p l a t e  procedure. The di f ference 
i s  seen t o  be about 15 percent. 
CONCLUDING REMARKS 
The following observations are made from the  da ta  presented here: 
1. Free-f l i g h t  skin-temperature measurements have indicated 
t r a n s i t i o n  t o  occur on a conical  nose a t  a Reynolds number of 
about 9 X 106 a t  a Mach number of 5.4. 
2. The theor ies  of Van Driest  f o r  turbulent  and laminar heat- 
t r ans f e r  coeff ic ients  appear adequate fo r  calcula t ing skin  temperature 
h i s t o r i e s  t o  Mach numbers on the order of 6. 
3. Radiative heat  t r ans f e r  a t  Mach numbers of 10 can play an 
important r o l e  i n  the  determination of peak skin  temperature. 
Langley Aeronautical Laboratory, 
National Advisory Committee f o r  Aeronautics, 
Lmgley Field,  Va., April  26, 1955. 
NACA Rf4 Lrj5El.G~ 
I, Van Driest ,  E. R.: The Turbulent Boundary Layer f o r  Compressible 
Fluids on a F l a t  Pla te  With Heat Tm.nsfer. Rep, No, A2-997, 
North American Aviation, Inc., Jan. 27, 1950, 
2. Eckert, Ernst R. G.: Survey on Heat Transfer a t  High Speeds. WADC 
Tech, Rep. 54-70, Wright A i r  Dev. Center, U. S. Air Force, 
Apr. 1954. 
3. Van Driest ,  E. R.: Investigation of Laminar Boundary Layer i n  Com- 
press ible  Fluids Using the  Crocco Method. NACA TN 2597, 1952. 
4. Piland, Robert 0.: Performance Measurements From a Rocket-Powered 
Exploratory Research Missile Flown t o  a Mach Number of 10.4, 
NACA RM ~54L29a, 1955. 
NACA RM L55E10c 
FLIGHT CONDITIONS FOR MODEL TESTED 
UP TO M=5.4 
8.5" DIAM. loo 
' ! f $ = 3 3 a - $ . 6  DAM. 
M ALT., FT 
6 -  - SOX lo3 
------- - _ _  
-----__ 
4 - - 40 
- 2 0  
0 5 
0 10 15 2 0  25 3 0  
TIME, SEC 
Figure 1 
MEASURED AND CALCULATED TEMPERATURE HISTORY 
TO A MACH NUMBER OF 5.4 
P-29'~-j r TEMR MEAS. 
- ME AS. 
0 TURB.,CONE 
(VAN DRIEST THEORY) 
TIME, SEC 
Figure 2 
MEASURED AND CALCULATED TEMPERATURE DISTRIBUTION 
FLIGHT TIME, 23.7 SEC; M =59; R/FT=9.8% lo6 
VAN DRIEST THEORY 
----_ 
------ - ----- TURB., CONE 
/
I I I I 1 
0 10 20 3 0  40 
STATION, IN. 
Figure 3 
NACA FM L55ELOc 
FLIGHT CONDITIONS FOR MODEL TESTED UP TO M-10.4 
r6a' ~liivl;,, ''9 
19.3" DlAM. ALT., FT Rz -26.6" 
-, 120x103 60x106 
Figure 4 
MEASURED AND CALCULATED TEMPERATURE HISTORY TO A MACH NUMBER 
OF 10.4 
TIME, SEC 
Figure 5 
NACA - Langley Field, Va. 


Aerodynamic heating of rocket-powered research vehicles at hypersonic speeds

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