.aJtt,.tI5 . ~S1f.

NASA TN D-1790

TECHNICAL NOTE

0-1790

HEAT-TRANSFER AND PRESSURE DISTRIBUTIONS AT MACH NUMBERS

OF 6.0 AND 9.6 OVER TWO REENTRY CONFIGURATIONS FOR

THE FIVE-STAGE SCOUT VEHICLE

By Paul F. Holloway and Jame s C. Dunavant

Langley Research 'Center Langley Station, Hampton, Va.

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

WASHINGTON May 1963

https://ntrs.nasa.gov/search.jsp?R=19630006418 2020-07-02T17:18:42+00:00Z

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

TECHNICAL NOI'E D-1790

HEAT-TRANSFER AND PRESSURE DISTRIBUTIONS Nr MACH NUMBERS

OF 6.0 AND 9.6 OVER TWO REENTRY CONFIGURATIONS FOR

THE FIVE-STAGE SCOUT VEHICLE

By Paul F. Holloway and James C. Dunavant

SUMMARY

Heat-transfer and pressure distributions have been, obtained for a five-stage Scout reentry configuration (both with and without a calorimeter nose cap) over an angle-of-attack range of 00 to 300 in the Langley 20-inch Mach 6 tunnel and 00 to 250 in the Mach number 9.6 nozzle of the Langley II-inch hypersonic tunnel. The. results indicate that modified Newtonian theory gives a good prediction of the trends of the pressure-distribution data. At an angle of attack of 00

, Lees' theory is shown to predict reasonably well the heat-transfer distribution over the blunt nose. Variation of Reynolds number and Mach number has virtually no effect on the heat-transfer distributions through the angle-of-attack range. The velocity-gradient correlation of Boison and Curtiss ~s shown to give a good prediction of the stagnation heating level at an angle of attack of 00 •

INTRODUCTION

As a part of the National Aeronautics and Space Administration's space research program, an investigation of the aerothermodynamic effects of reentry into the earth's atmosphere at hypervelocity speeds is currently being conducted with a five-stage Scout vehicle. During a flight test, the first two stages of the Scout configuration boost the vehicle out of the atmosphere on the ascent phase of the flight, and the three remaining stages accelerate the vehicle on the descent phase of the flight to a maximum reentry velocity of approximately 30,000 feet per second. A detailed description of an early version of the Scout vehicle is given in reference 1.

In support of this project, the pressure and heat-transfer distributions of an early version of the reentry nose cone, both with and without the calorimeter nose cap, have been obtained at Mach numbers of 6.0 and 9.6. The objectives of these wind-tunnel experiments were to determine the maximum heat-transfer rate and the heat-transfer distributions at various angles of attack to implement the design of the spacecraft heat protection and also to provide a comparison for the flight heating results.

The purpose of this report is to present the heat-transfer and pressure distributions over a fifth-stage Scout reentry spacecraft, both with and without the calorimeter nose cap. The angle~of-attack range was 00 to 300 at a Mach number of 6.0 and 00 to 250 at a Mach number of 9.6. The Reynolds number of the heat-transfer tests at 00 angle of attack has been varied over a wide range so that the effect of this parameter on the laminar heat transfer at hypersonic speeds may be ascertained. Also, by comparison of the heat-transfer data for Mach numbers of 6.0 and 9.6, the combined effect of Reynolds number and Mach number on laminar heat transfer at ,angle of attack is observed. Finally, the effect of model size and test techniques on the accuracy of heat-transfer tests is discussed.

D

h

K

M

p

r

!hemi

Rr

s

2

SYMBOLS

specific heat of model construction material

pressure coefficient,

diameter of curvature of spherical segment nose, 2rhemi

time derivative of measured wall temperature

heat-transfer coefficient, defined in equation (1)

thermal conductivity of air

Mach number

Prandtl number

pressure

dynamic pressure

nose reference length (distance from cone center line to intersection of corner radius and cone, see fig. 1)

radius of curvature of spherical segment nose

radius of cross section at a longitudinal station

free-stream Reynolds number based on r

surface distance measured from 00 angle-of-attack stagnation point

t

T

u

x

a.

p

T

local skin thickness

temperature

velocity

distance along longitudinal axis measured from stagnation point at 00 angle of attack

angle of attack referenced to axis of cone

velocity gradient, du/ds

cone half- angle

angle between free-stream velocity vector and a vector normal to body surface

viscosity

density

time

meridian angle defined in figure 1

Subscripts:

aw adiabatic wall conditions

I local conditions

s stagnation point

t total

w wall

00 free-stream conditions

cr free-stream conditions immediately behind normal shock

CONFIGURATION CONCEPTS

In some instances, shapes of reentry bodies may be dictated by requirements other than aerodynamic. Such was the case in the design of the two reentry nose cones investigated in this report. The design of the fifth stage of a Scout vehicle designated to reenter the earth's atmosphere at speeds greater than orbital velocity was dictated by the mission requirements and the dimensional

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limitations set by the booster stages. The maximum length and diameter of the reentry stage were limited by booster and heat-shield considerations, and the minimum diameter was necessarily greater than that of the 17-inch spherical rocket motor of the fifth (payload) stage. Nose bluntness was dictated by the mission requirement that the reentry vehicle have a low maximum heating rate. Thus, the reentry shape evolved, a shape that was essentially a short, very blunt, low-angle cone. A slight modification to this shape was produced when a thin-shell, metal calorimeter was placed over the forward half of the flight reentry capsule, and the cone angle over that portion of the nose cone was thereby reduced. More details of the experiment are given in reference 2.

The final configurations of the flight reentry spacecraft differ from the configurations of this investigation in that the cylindrical afterbody was eliminated from the flight configuration, and the forward portion of the config­uration after the calorimeter nose cap was jettisoned had a short cylindrical section aft of the nose-corner radius. Some longitudinal aerodynamic character­istics of several versions of the reentry configurations with and without the calorimeter nose cap are presented in references 2 and 3.

APPARATUS AND METHODS

Models

Sketches of the exterior shape of the models are shown in figure 1. Fig-ure lea) shows configuration I, the shape without the calorimeter nose cap and figure l(b), configuration II, the shape with the calorimeter nose cap. A total of eight models, two pressure models and two heat-transfer models of each of the two configurations were made (the larger of. each twice the size of the smaller). The designations (a) and (b) will be used to refer to the large and small models, respectively. The smaller models were designed for test in the smaller M = 9.6 tunnel and the larger models, for the M = 6.0 . tunnel. The pressure models were machined from solid stock and the relativ,ely thick (about 1/8 inch) walls were instrumented with 0.040 I.D. pressure orifices at locations given in table I. The heat-transfer models were made from Inconel sheet, spun in two pieces, and welded together at a point behind the corner radius. Thermocouples were not installed near the weld joint. The thickness of the models was 0.050 inch on the blunt nose and tapered to a 0.030-inch thickness on the conical portion, the taper ueginning on the corner radius. (Local measured skin thicknesses are listed in table L) Thermocouples were silver-soldered in holes drilled in the skin at locations also given in table I. Instrumentation was located in the windward half of the model (for positive angles of attack). Leeward data were obtained by testing the model at negative angles of attack.

Wind Tunnels

The tests were conducted at M = 6 in the Langley 20-inch hyperSOnic tunnel 'and in the M = 9.6 nozzle of the Langley II-inch hypersonic tunnel. The 20-inchhypersonic tunnel is of the intermittent type operating from a stored air

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supply at stagnation pressures from 19 to 38 atmospheres and a maximum stagnation temperature of 6000 F. The tunnel exhausts to the atmosphere through a diffuser augmented by an air-ejector. A more detailed description of the tunnel is given in reference 4. Most of the tests were made at a stagnation pressure of 24 atmos­pheres and temperatures of 5000 F to 6000 F, for which the corresponding Reynolds number per inch is approximately 0.48 X 106; however, for an angle of attack of 00 , both the small and large models were tested over a range of pressure and tem-

perature to obtain uni~ Reynolds number (per inch) of 0.33 X 106 to 0.70 X 106 .

The M = 9.6 nozzle of the II-inch hypersonic tunnel is a three-dimensional rectangular, contoured nozzle having a calibrated Mach number of 9.6 at the nom­inal test conditions of 46 atmospheres total pressure and 12000 F total tempera-.ture. The Reynolds number per inch for these test conditions is 0.10 X 106 . Only the small models were tested in this facility. The tunnel is also of an inter­mittent blowdown type but exhausts to a vacuum sphere and has a running time of 1 to 2 minutes. A more detailed description of the nozzle and some calibration data can be found in reference 5.

Methods

Pressures.- The model static pressures in the M = 6 tests were recorded by photographing a multiple-tube mercury manometer board. Tunnel stagnation pressure was measured on a calibrated Bourdon gage. This method leads to inac~racies of any measured pressure of less than ±2 percent.

For the M = 9.6 tests, the pressures were recorded on six-cell aneroid, recording type pressure instruments. Pressures were read 60 seconds after the flow was initiated to insure that the pressure in the cell was fully stabilized. Inaccuraqy for any measured pressure in the M = 9.6 tests is less than ±2~ percent.

Heat transfer.- Aerodynamic heating was measured by the transient calorimetry technique by which the rate of heat storage in the model skin is measured. The models initially at room temperature were suddenly exposed to the airstream and the rate of temperature rise of the skin was measured as soon as possible while the model was in a nearly isothermal condition; hence conduction between the sur­face elements was a minimum. In the M = 6 facility, exposure of the model was accomplished by quickly injecting the model into the stream from a sheltered posi­tion beyond the tunnel wall. With injection, there was a cavity in the tun-nel wall which led to interference effects on some of the leeward data. These effects are discussed subsequently. There was no cavity in the tunnel wall during the pressure tests. Injection was accomplished in less than 0.25 second.

In the M = 9.6 tests, the model was installed in the test section prior to start of the airflow. The heating rates were measured as soon as flow condi­tions (set~ling chamber pressure and temperature) stabilized and while the model was still at nearly isothermal conditions. Approximately 2 seconds were required to ~chieve these conditions.

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DATA REDUCTION

The Mach number used in the reduction of the pressure and heat-transfer data for both the M = 6 and M = 9.6 tests was based on previou~ tunnel calibrations.

At M = 6, the thermocouple outputs were recorded on a BeCkman 210 high­speed digital data recording system. The output voltage of each thermocouple was sampled at a rate of 40 times per second, converted to a binary digital sys­tem, and recorded on magnetic tape. The temperature-time data were fitted to a second-degree curve by the method of least squares, and the time derivative of temperature was computed on a card programed computer.

At M = 9.6, the thermocouple outputs were continuously recorded on four 18-channel D'Arsonval type galvanometers. The time derivative of temperature was then determined graphically from the temperature-time curve.

The measured local heat-transfer coefficient was calculated from the fol­lowing relation:

where Tw = Measured wall temperature and Taw is given by

Taw = T + VNpr(Ts - T) (2)

where NPr was assumed to be 0.69 and the temperature T was calculated from an isentropic expansion of the flow from the stagnation-point pressure and tem­perature behind a normal shock wave to the measured local static wall pressure. The heat-transfer coefficients were not corrected for any lateral conduction of heat in the model skin at either M = 6.0 or M = 9.6.

For the M = 6.0 data, the heat-transfer coefficients were computed for a time interval of approximately 0.1 to 0.6 second after the model was injected into the airflow. The coefficients presented herein were calculated from a time derivative of temperature and a measured wall temperature determined at approxi­mately 0.10 second after the model is in position in the tunnel. The maximum surface temperature increase was 250 F at the time at which the coefficients were calculated. The low skin-temperature increase together with the thin skin thickness minimized the conduction error. The exact skin thickness was measured at each thermocouple location. The repeatability of M = 6.0 heat-transfer data is generally within ±5 percent. The maximum inaccuracy of the data in most regions is believed to be ±lO percent. However, in the corner regions where the conduction is large, the data may be in error by as much as ±20 percent.

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For the M = 9.6 tests, by using a quick-starting technique and allowing the flow to stabilize before calculating the heat-transfer coefficients, the maximum surface temperature increase was 1600 F. This relatively large increase compared with that of the M = 6 tests led to a higher conduction error, par­ticularly in the corner region. The repeatability of the M = 9.6 heat-transfer data is generally within ±10 percent. The inaccuracies of the data on many regions are also believed to be within ±10 percent. However, in regions where the conduction is large, it will be shown later that the data may be in error by 50 percent or more.

RESULTS AND DISCUSSION

Schlieren Photographs

Figure 2 presents schlieren photographs of the two configurations at various angles of attack in both the Mach 6.0 20-inch hypersonic tunnel and the Mach 9.6 nozzle of the ll-inch hypersonic tunnel. These photographs were obtained during the pressure tests.

An effect of the different corner radii of the two configurations on the local flow in the corner region is evident in the comparison of figure 2(a) with figure 2(b) for the M = 6 tests. On configuration I(a) (fig: 2(a))" there is a small separation region apparent for ~ ~ 100 as evidenced by the presence of a reattachment shock. The strength of the reattachment shock decreases as ~

is increased from 00 to 100 . Separation above an angle of attack of 100 is not apparent on the windward side of the model; however, in figure 2(b) the small separated region at the corner of configuration II(a), which has a much smaller corner radius, remains through the a,ngle-of - attack range . Moreover, the sepa­rated region on configuration II(a) is much larger than that of configura-tion I(a) since reattachment occurs farther back along the cone surface. This separation is not observed in the schlieren photographs at M = 9.6 (fig. 2(c)).

No boundary-layer transition is evident from an examination of the schlieren photographs.

Pressure Distribution

In figure,s 3 and 4, the pressure distributions over configuration I, the reentry cone without the calorimeter nose cap, are presented for an angle-of­attack range of 00 to 300 at a Mach number of 6.0 (fig. 3) and 00 to 250 at a Mach number of 9.6 (fig. 4). Figures 5 and 6 present the pressure distributions over configuration II, the reentry cone with the calorimete~ nose cap, ~or an angle-of-attack range of 00 to 300 at a Mach number of 6.0 (fig. 5) and 00 to 250

at a Mach number of 9.6 (fig. 6). In each figure, the experimental pressure distribution is compared with the distribution predicted by modified Newtonian theory for which

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Cp = C cos2n p,s 'I

where Cp,s is the maximum (stagnation) pressure coefficient behind a normal

shock wave at the calibrated free-stream Mach number.

General agreement of theory and experiment.- Comparison of the experimental pressure distributions with modified Newtonian theory in figures 3 to 6 shows good agreement of the trends of the data and in most cases of the local pressure. However, it should be pointed out that the deviation of theory and experiment in the corner region will lead to large differences in the experimental and theoret­ical velocity gradients for this region.

Spherical nose.- In figures 3 to 6 it can be seen that the pressures in the stagnation region are very well predicted for each configuration over the angle­of-attack range. However, as the subsonic flow over the spherical nose accel­erates from the stagnation point toward the corner, there is an experimental pressure deviation below the Newtonian predictions for both configurations that increases as the distance from the stagnation point is increased for a ~ 15°. For a> 150 , the location of the maximum pressure does not move as far windward as Newtonian theory predicts so that the measured pressures over the spherical nose are higher than the theoretical pressures.

Corner region.- The surface pressure data on the corner radius of the con­figurations tested are always lower than modified Newtonian theory. (See figs. 3 to 6.) In figure 5, the effects of the separation region and reattach­ment shock on configuration II (discussed in the flow-field analysis section) are apparent on the pressures around the corner and just aft of the corner-cone junction. The overexpansion and subsequent separation of the flow on the corner creates a low pressure on the corner and extreme forward tip of the cone compared with the pressures rearward on the cone. This apparent separation phenomenon is not observed at M = 9.6 in figure 6. The pressure distributions of configura­tion I in figures 3 and 4 also do not indicate separation. As noted in the flow­field analysis section, if separation did occur on configuration I, it affected a much smaller region than on configuration II. Since there is less instrumenta­tion in the corner region of configuration I than in that of configuration II, a small reg~on of separation might not be seen in the pressure distributions.

Conical-cylindrical afterbody.- Comparison of experimental pressure distri­butions with the Newtonian theory in figures 3 to 6 shows that the pressures over the forward portion of the cone of each configuration are always larger than the Newtonian pressure. These high experimental pressures are induced by the strong bow shock wave caused by the very blunt nose shape of each configura­tion. In fact, the leeward surface pressures are larger than the Newtonian values over the entire cone surface, even when portions of the cone are in the Newtonian "shadow" region. On the windward side of the cones, the nose blunt­ness effects quickly die out with increasing values of sir for a ~ 150 so that the surface pressures over the rearward portion of the cone are either about equal or slightly less than the Newtonian pressures. For a > 150 , the windward surface pressures on the conical surface are higher than the theory over most of the cone, the deviation increasing as the angle of attack increases.

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The surface pressures on the cylindrical afterbody (figs. 3 to 6) are generally in agreement with the theory through the angle-of-attack range and the Mach numbers investigated for each configuration. The experimerital pressures on the cylindrical afterbody of configuration I are approximately the same as those on configuration II; thus, the nose-shape variation has no effect in this region.

Heat-Transfer Distributions

The heat-transfer distributions at both Mach number 6.0 and 9.6 are pre­sented in figures 7 and 8 for configurations I and II, respectively, over the angle-of-attack range. The measured heat-transfer coefficients referenced to the theoretical stagnation heating rate of a hemisphere hs,hemi determined from a modification of the theory of Sibulkin are presented in dimensionless ratio form. (See ref. 6.) The heating rate at the stagnation point is given by

where the velocity gradient parameter at the stagnation point ~D was deter­(J

( 4)

mined from modified Newtonian theory. The diameter of curvature of the spherical segment nose D was used for both configurations to calculate hs hemi. ,

It should be noted that at M = 6.0, the heat-transfer distribution over the leeward portion of the cone was not obtained for angles of attack above 100 •

Also, at M =6.0, the data for the rearward portion of the leeward cone region of configuration I(a) at ~ = 100 and configuration II(a) at ~ = 50 and 100

have been omitted because of interference effects caused by a disturbance ema­nating from the tunnel wall cavity (for model injection) and intersecting the conical portion of the model.

Finally, a model imperfection just aft of the 5.70 and 90 cone intersection on model II(a) was created during thermocouple installation. This protrusion on the ¢ = 900 ray apparently affected the heating rates rearward of the cone intersection at certain angles of attack and Reynolds numbers by tripping the boundary layer and thereby causing transition and higher heating rates.

Data accurac as affected b model size and test techni ues.- The distribu­tions of heating on the nose of both models I a and II a at M = 6.0 (figs. 7 and 8) show the maximum heating to be near the corner radius even at ~ ~ 00 •

The comparable tests at M = 9.6 using the smaller models do not show a similar rise in the heating toward the edge radius even though theory gives a rise which is essentially independent of Mach number. Furthermore, just downstream of the corner radius, the heating distribution at M = 6.0 usually shows a small mini­mum (note particularly the windward ray) which is also not seen in the M = 9.6' distributions. These differences in results are attributed to the effects of

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model size and testing techni~ues. The smaller size of the models tested at M = 9.6 and the larger time re~uired to stabilize flow conditions lead to larger conduction errors. This is of particular importance in the region of small radius (hence, large heat-transfer gradients which result in lar~e skin-temperature gradients) such as the corners and adjacent regions of configurations I and II. The shorter surface lengths and longer times involved in the M = 9.6 tests permitted the heat absorbed by the skin at the forward part of the corner to be conducted to the lower temperature skin just downstream of the corner. Hence, neither the peak heating ratio ahead of the corner nor the minimum just down­stream of the corner seen in the M ~ 6 tests was detected in measuring the heat stored in the skin during the M = 9.6 tests. The increase in heating in the corner region above the stagnation level at ~ = 00 is of importance and veri­fies the predictions of the theories of Lees (ref. 7), Beckwith and Cohen (ref. 8), and others for a body of this type.

Discussion.- The distributions presented in figures 7 and 8 indicate that the heat transfer in the region of the forward-most corner is always greater than the measured values at the ~ = 0 geometric vertex regardless of angle-of-attack range for M = 6. The increase may be as much as 65 percent, depending on the angle of attack. This increase is also indicated for the M = 9.6 distribu­tions at angles of attack ~100.

For configuration I(a) at M = 6.0, a slight separation region on the cone just beyond the corner at 00 angle of attack (fig. 7(a)) is indicate~ by the relatively low heating measured. For ~ > 00 (figs. 7(b) to 7(f)), there is insufficient instrumentation to determine whether the flow separates over the windward ray. A separation region is not apparent in the M = 9.6 data for configuration I(b).

In figure 8(a), a possible separation region is indicated for configura­tion II(a) at M = 6.0. For ~> 00 (figs. 8(b) to 8(f)), the separation region is always evident on the horizontal ray. On the windward ray, the flow also might be separated up to an angle of attack of 150 • However, the heat­transfer coefficient is higher than might be expected for a separated region. Perhaps this is due to conduction. No comparable decrease in heating is seen in the corner region for the M = 9.6 data of configuration II(b).

It is felt that the increase in heating rates over the rearward portion of the cone on the windward ray for the M = 6.0 distributions (on configura­tion II(a) in fig. 8) is the effect of boundary-layer transition which may have been influenced by the protrusion in the model skin previously mentioned. At ~ = 300 , natural transition may also have occurred. (See fig. 7(f).)

An unusual trend is apparent in the distributions for configurations I(b) and II(b) at M = 9.6. (See figs. 7(e) and 8(e).) The heat transfer over the leeward" (vertical) ray begins to increase for a sir value of approximately -3.0 and increases by 500 percent. This unusual phenomenon is unexplained.

For comparison, the heating rates on the windward ray of the cone are pre­dicted by a simple crossflow theory where the cone is considered to be a cyl­inder swept to an angle e~ual to the local surface inclination and having a

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radius equal to the local cone radius. A prediction of the heating rate can be obtained from the results of references 9 and 10 as

__ h __ O. 57rhemi (90 !;::.) - ---- cos - a - u

hs,hemi 0.76 rl

This prediction is presented for the highest angles of attack, 250 and 300, in figures 7(e), 7(f), 8(e), and 8(f). The agreement of this theory with the meas­ured heating was extremely good, a result that was not expected when the short length of the body was considered.

Mach number and Reynolds number effects at a = 00 ._ Figure 9 presents the measured heat-transfer distribution referenced to the theoretical stagnation heating on a hemisphere over the full Reynolds number range at an angle of attack of 00 • The increase of Mach number from 6.0 to 9.6 and the concurrent decrease in Reynolds number by a factor of 10 did not significantly affect the laminar heat transfer for the configurations tested through an angle-of-attack range of 0° to 250 • There is seen to be little uniform effect on the heating rates except in the corner region and over the rear portion of the cone. The devia­tion of heating rates in the corner region between models r(a) and reb) or models rI(a) and II(b) at M = 6 indicate that the conduction error in a region of small radius is mainly due to the smaller model size since the t~sting tech­nique is the same for all t.he M = 6.0 tests. Evid.ence is given that the con­duction error may also wash out the apparent separation effect since the smaller models do not indicate the effects of separation on heat transfer as do the larger models. However, it should be noted that this may very well be an effect of Reynolds number variation since there are no schlieren photographs available to confirm the existence of separation on the smaller models.

The variation in heating over the rearward portion of the cone may be due to a Reynolds number effect connected with the beginning of transition. Cer­tainly this is evident for configuration II (fig. 9(b), Rr = 0.93 X 106 ) where for the maximum Reynolds number the boundary layer is definitely transitional.

In reference 11 the effect of having a spherical segment nose rather than a full hemisphere is shown to increase the stagnation velocity gradient and hence stagnation-point heating above the hemisphere stagnation level. In fig­ure 10, the measured stagnation heating rates are compared with the heat-transfer rate by using the velocity gradient of Boison and Curtiss in reference 11. Using the Boison-Curtiss correlation as a reference, the variation of the measured stagnation-heating-rate ratio is generally less than 10 percent for each config­uration over the Reynolds number range (in which Rr is varied by a factor of 12 times the minimum value obtained in the M = 9.6 tests). Since this variation agrees with the accuracy which was expected in these tests, it may be said that Reynolds number had no effect on the stagnation-heating parameter h/hs,hemi.

Comparison of heat-transfer distribution with theory at a = 00 ._ Figure 11 presents the measured heat-transfer distributions referenced to the measured stagnation heating rate for the full range of Reynolds number. Also presented

11

in figure 11 are the theoretical heat-transfer distributions as predicted by the theory of Lees (ref. 7).

Calculations have been based on a Newtonian pressure distribution for M = 6.0 and M = 9.6 and measured pressure distribution for M = 6.0. The theory agrees reasonably well with the experimental values over the nose surface up to the corner region for each configuration, as might be expected. The rise in heating on the corner region is less than that predicted by theory. However, even for the M = 6 tests there must be some conduction error in the corner region so that it is reasonable to postulate that more accurate experimental data (less conduction) should give better agreement than the results indicate. Over the conical-cylindrical afterbody, the theory gives a reasonable prediction in the trends of the data. However, the theory neglects the effects of pressure gradient on boundary-layer profiles and is expected to be higher than experiment. Hence, the theory is applied here only to give qualitative results and is not intended to give an accurate prediction in this region. The theoretical distri­bution based on the measured pressure distribution gives the best estimate of the trends of the data. It is evident that a change in pressure distribution (that is, between measured and Newtonian values) can give a marked variation in theoretical predictions. Note that the theory (using a Newtonian pressure dis­tribution) predicts a much larger increase in heating at the 5.70 to 90 cone junction of configuration II (fig. ll(b)) than was found in any of the tests.

SUMMARY OF RESULTS

Heat-transfer and pressure distributions have been obtained for two reentry configurations for the five-stage Scout vehicle over an angle-of-attack range of 00 to 300 in the Langley 20-inch Mach 6 tunnel and 00 to 250 in the Mach number 9.6 nozzle of the Langley ll-inch hypersonic tunnel. Analysis of the experimental data and comparison with theory have yielded the following results:

1. Modified Newtonian theory gives a good prediction of the trends of the pressure distributions through the angle-of-attack range.

2. Heat-transfer data accuracy is strongly affected by model size and testing techniques particularly in a region of large gradients of heat transfer (corner region). Experimentally, the heat-transfer peak on the corner can be lost because of conduction.

3. The heating rates on the corner region of the configurations tested are greater than the measured values at the a = 0 geometric vertex regardless of angle of attack within the range investigated for M = 6. The increase may be as much as 65 percent depending upon the angle of attack.

4. The increase of Mach number of 6.0 to 9.6 and the concurrent decrease in Reynolds number by a factor of 10 did not significantly affect the laminar heat­transfer distribution over the configurations tested through an angle-of-attack range of 00 to 250 .

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5. At the high angles of attack (250 and 300 ) for which the low-angle con­ical afterbody may be approximated by a cylinder, crossflow theory was found to predict the measured heating level on the windward ray extremely well.

6. The variation of free-stream Reynolds number by a factor of 12 had no significant effect on either the stagnation heating ratio or the heat-transfer distributions at an angle of attack of 00 where the flow was laminar.

7. The correlation of Boison and Curtiss for the effect of having a spher­ical segment nose rather than a full hemispherical nose on the stagnation heating agrees well with the experimental results.

8. The theory of Lees agrees reasonably well with the experimental results over the nose of the configurations. On the conical afterbody, Lees theory is higher than the measured heating rates.

Langley Research Center, National Aeronautics and Space Administration,

Langley Station, Hampton, Va.; February 18, 1963.

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REFERENCES

1. Mayhue, Robert J., Compiler: NASA Scout ST-l Flight-Test Results and Analyses, Launch Operations, and Test Vehicle Description. NASA TN D-1240, 1962.

2. Johnston, Patrick J.: Longitudinal Aerodynamic Characteristics of Several Fifth-Stage Scout .Reentry Vehicles From Mach Number 0.60 to 24.4 Including Some Reynolds Number Effects on Stability at Hypersonic Speeds. NASA TN D-163S, 1963.

3. Harrison, Edwin F.: Static Stability Tests in the Langley 24-Inch Hypersonic Arc Tunnel on a Blunted Cone at a Mach Number of 20. NASA TN D-150S, 1962.

4. Sterrett, James R., and Emery, James C. I: Extension of Boundary-Layer­

Separation Criteria to a Mach Number of 6.5 by Utilizing Flat Plates With Forward-Facing Steps. NASA TN D-61S, 1960.

5. Bertram, Mitchel H.: Boundary-Layer Displacement Effects in Air at Mach Numbers of 6.S and 9.6. NASA TR R-22, 1959. (Supersedes NACA TN 4133.)

6. Crawford, Davis H., and McCauley, William D.: Investigation of the Laminar Aerodynamic Heat-Transfer Characteristics of a Hemisphere-Cylinder in the Langley ll-Inch Hypersonic Tunnel at a Mach Number of 6.S. NACA Rep. 1323, 1957. (Supersedes NACA TN 3706.)

7. Lees, Lester: Laminar Heat Transfer Over Blunt-Nosed Bodies at Hypersonic Flight Speeds. Jet Propulsion, yolo 26, no. 4, Apr. 1956, pp. 259-269, 274.

S. Beckwith, Ivan E., and Cohen, Nathaniel B.: Application of Similar Solutions to Calculation of Laminar Heat Transfer on Bodies With Yaw and Large Pres­sure Gradient in High-Speed Flow. NASA TN D-625, 1961.

9. Truitt, Robert Wesley: Fundamentals of Ae:r:odynam1c Heating. The Ronald Press Co., c.1960.

10. Feller, William V.: Investigation of Equilibrium Temperatures and Average Laminar Heat-Transfer Coefficients for the Front Half of Swept Circular Cylinders at a Mach Number of 6.9. NACA RM L55FOSa, 1955.

11. Boison, J. Christopher, and CurtiSS, Howard A.: An Experimental Investiga­tion of Blunt Body Stagnation Point VelOCity Gradient. ARS J0ur., vol. 29, no. 2, Feb. 1959, pp. 130-135.

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TABLE L- COORDINATES OF THERMOCOUPLES AND PRESSURE ORIFICES

(a) Configuration I

Skin thickness, t, in. , Instrumentation for -Coordinates

for heat-transfer model - Heat-transfer model Pressure model

sir x/r ¢, deg rl/r I(a) I(b) r( a) reb) rea) I(b)

0 ----- --- 0 0.049 0.053 X X X X .229 ----- 0 .229 .050 .053 X X X .229 ----- 90 .229 .050 .053 X X X X .459 ----- 0 .456 .050 .050 X X X .459 ----- 90 .456 .050 .050 X X X X .687 ----- 0 .681 .049 .046 X X X X .687 ----- 30 .681 .049 .046 X X X X .687 ----- 60 .681 .050 .047 X X X X .687 ----- 90 .681 .049 .046 X X X X .847 ----- 90 .839 .043 .043 X X

.928 ----- 90 ·920 X X ·950 ----- 0 ·940 .045 .040 X X X X

1.071 0.246 90 ----- .031 .028 X X 1.166 .341 0 ----- .031 .028 X X X 1.166 ·341 30 ----- .032 .026 X X X 1.166 ·341 60 ----- .032 .026 X X X 1.166 .341 90 ----- .036 .039 X X X X 1.304 .478 0 ----- .029 .031 X X X 1.434 .627 0 ---...;- X X 1.434 .627 90 ----- X X

1.705 .873 90 ----- .030 .030 X X 1·914 1.104 0 ----- .031 .030 X X X X 1.914 1.104 165 ----- .030 .031 X X 1.914 1.104 150 ----- .030 .031 X X X X 1.914 1.104 135 ----- .030 .031 X X 1·914 1.104 120 ----- .030 .031 X X X X 1.914 1.104 105 ----- .030 .031 X X 1·914 1.104 90 ----- .030 .031 X X X X 2.424 1.582 90 ----- .029 .031 X X 2·908 2.060 0 ----- .029 .031 X X X

2.908 2.060 90 ----- .029 .031 X X X X 3.336 2.480 90 ----- .030 .031 X X 3·881 2.882 0 ----- .032 .034 X X X X 3.881 2.882 165 ----- .032 .032 X X 3.881 2.882 150 -_ .... _- .032 .032 X X X X 3.881 2.882 135 ----- .032 .032 X X 3 .• 881 2.882 120 ----- .032 .032 X X X X 3·881 2.882 105 ----- .032 .031 X X 3.881 2.882 90 ----- .030 .031 X X X X

4.373 3·503 90 ----- .030 .031 X X X 4.866 3.983 0 ----- .032 .032 X X X 4.866 3·983 90 ----- .032 .034 X X X X 5·225 4.487 90 ----- .032 .032 X X X X 6.078 5.186 0 ----- .029 .032 X X X 6.078 5·186 90 ----- .030 .033 X X X X

15

TABLE I.- COORDINATES OF THERMOCOUPLES AND PRESSURE ORIFICES - Concluded

(b) Configuration II

Skin thickness, t, in., Instrumentation for -Coordinates

for heat-transfer model -Heat-transfer model Pressure model

s/r x/r ¢, deg rZ/r II(a) II(b) II(a) II(b) II(a) iI(b)

0 0 --- 0 0.051 0.051 X X X X ·311 ----- 0 ·310 .051 .051 X X X X ·311 ----- 45 ·310 .051 .051 X X X ·311 ----- 90 ·310 .051 .051 X X X X .468 ----- 90 .465 X .623 ----- 0 .618 .051 .050 X X X X .623 ----- 30 .618 .051 .050 X X X .623 ----- 60 .618 .051 .050 X X X .62;' ----- 90 .618 .051 .050 X X X X ·778 ----- 90 ·770 X

·933 ----- 0 ·919 .046 .047 X X X X ·933 ----- 15 ·919 .046 .047 X X X .933 ----- 30 ·919 .046 .047 X X X X ·933 ----- 45 ·919 .046 .047 X X X ·933 ----- 60 ·919 .046 .047 X X X X ·933 ----- 75 ·919 .046 .047 X X X ·933 ----- 90 ·919 .046 .046 X X X X ·977 ----- 0 .962 .043 .037 X X ·977 ----- 90 ·962 .043 .042 X X

1.075 .248 90 ----.- .032 .032 X X X X

1.328 ·501 0 ----- .030 .033 X X X X 1.328 ·501 90 ----- .029 .033 X X X X 1.583 . ·749 0 ----- .032 .031 X X X 1.583 .749 90 ----- .032 .033 X X X X 2.085 1. 25\ 0 ----- .032 .030 X X X X 2.085 1.251 165 ----- .032 .032 X X X 2.085 1.251 150 ----- .032 .032 X X X X 2.085 1.251 135 ----- .032 .032 X X X 2.085 1.251 120 ----- .032 .032 X X X X 2.085 1.251 105 ----- .032 .031 X X X

2.085 1.251 90 ----- .032 .031 X ,x X X 2·580 1. 750 90 ----- .034 .032 X X X 3.080 2.242 0 ----- .033 .032 X X X 3·080 2.242 90 ----- .033 .. 032 X X X X 3.594 2·753 90 ----- .032 .033 X X X 4.089 3. 235 0 .03'1 .032 X X X X 4.089 3·235 165 ----- .030 .032 X X X 4.089 3·235 150 ----- .030 .033 X X X X 4~089 3·235 135 ----- .030 .032 X X X 4.089 3·235 120 ----- .030 .032 X X X X

4.089 3·235 105 ----- .030 .033 X X X 4.089 3·235 90 ----- .030 .033 X X X X 4.612 ~.740 90 ----- .032 .032 X X X 5·112 4.248 90 ----- .031 .032 X X X 5.581 4.706 0 ----- .030 .029 X X X X 5·581 4·706 90 ----- .031 .030 X X X X

16

I-' ......:j

I

I f---ll

,l __ 3.zsr

=====-=-=+_1 11 -s

<t rt)

+5

I· 4.84r

(a) Model 1.

5.7 0

.10r

2..98 r

(b) Model II.

9.00

~, ,J .\. -I 1.07r

I I

o rt)

~, ,J 2.39r I. .1 .98r

Figure 1.- Model dimensions.

MODEL I(a) r= 1.466 inches

MODEL reb) r =.7 33 inc hes

MODEL II (a) r= 1.614 inches

MODEL II(b) r=.807 inches.

18

(a) Configuration I at M = 6.0 .

Figure 2 .- Schlieren photographs.

L-63-55

a = 10°

a= 20°

(b) Configuration II at M = 6.0. L-63-56

Figure 2.- Continued.

19

Configuration I Configura tion n

(c) M = 9.6. L-63-57

Figure 2 .- Concluded.

20

I\) I-'

Cp

1.5

1.0

.5

a = 2.5°' 0

1.5

1.0

.5

a = 0° 0

-.5 -7

Cylinder

-6 -5

Cone

-4 -3 -2

Corner radius

-1

Sphericai nose

o

sir

1

Corner radius

(a) ~ = 0° and 2.5°.

2

Figure 3.- Pressure distribution over blunt-nose cone at M

0, deg

o gO (vertical meridian) D 0 (horizontal meridian) <> 30 ,6 eo

Modified Newtonian theory (vertical meridian) Modified Newtonian theory (horizontai meridian)

Cone Cylinder

4 5 6

6.0. Configuration I(a).

7

£\) £\)

Cp

1.5

1.0

;5

a = 7.50 : 0

1.5

1.0

.5

a = 50 0

-.5 -7

Cylinder Cone

-6 -5 -4 -3 -2

Corner' radius

-1

(b) a

Spherical nose

o

sir

1

5° and '7·5°·

Figure 3.- continued.

Corner radius

2

¢, deg

o 90 (vertical meridian) o 0 (horizontal meridian) o 30 .t:::. 60

---- Modified Newtonian theory (vertical meridian) - - - - Modified Newtonian theory (horizontal meridian,

Cone Cylinder '," I

3 4 6 7

r\) \.>I

1.5

1.0

.5

a = 12.50 0

Cp

1.5

1.0

.5

a = 100 o

-.5 -7

Cylinder Cone

-6 -5 -4 -3 -2

Corner JJ I JJJJJj~corner

radius Spherical _ radius . nose

-1 0

sir

( c) a = 10° and 12.5°.

Figure 3.- Continued.

2

¢, deg

o 90 (vertical meridian) o 0 (horizontal meridian) o 30 6 60

Modified Newtonian theory (vertical meridian) Modified Newtonian theory (horizontal meridian)

Cone

4 5 6

Cylinder I'

7

r\)

+=-

Cp

1.5

1.0

.5

a = 20°' o

1.5

1.0

.5

a = 15°' 0

-.5 -7

Cylinder Cone

-6 -5 -4 -3 -2

Corner radius

-1

Spherical nose

o 1

sIr

(d)' a = 15° and 200.

Figure 3.- Continued.

Corner radius

2

~, deg

o 90 (vertical meridian) o 0 (horizontal meridian) <> 30 ('" 60

Modified Newtonian theory (vertical meridian) - Modified Newtonian theory (horizontal meridian)

,Cone}" i-Cylinder

3 4 5 6 7

I\) \J1

Cp

+j

1.5 It

1.0

.5

a = 300 o Iii i I I i I 1-' II iii I Iii i I I I iii I I I I I I I I I I I i I I I iii iii I I I iii iii iii iii Iii i I 1I1111111111i,1I

1.5

1.0

.5

a = 250 0

-.5 -7 -6 -5 -4 -3 -2

, , I,

Spherical. nose

iE! -1 0 1

sir

(e) a = 25° and 30°.

Figure 3.- Concluded.

Corner radius

2

¢, deg

o 90 (vertical meridian) D 0 (horizontal meridian) <> 30 /',. 60

Modified Newtonian theory (vertical. meridian) Modified Newtonian theory (horizontal. meridian) ,

TrP

11'1

+ 'fl'IH'

II 11'1

, : CTI -tt±±±:

I'

, ,

" ,

~llllll ! I' '~"I i II~II i i 1;)11 P~I, 1IIIi II i i 111:1 \ Iii j",fj t ,1 . , , , II

I j

T--4 Cone Cylinder

3 4 5 6 7

~

Cp

1.5

1.0

.5

a ~ 50 0

1.5

1.0

.5

a ~ 00 0

III!IIIII'I 5mlr~l;elrl HII fllllllllllllillione

-. -7-6 -5 -4 -3 -2

Corner radius

-1

Spherical .ose

o

sir

~, deg

o 90 (vertical meridian) o 0 (horizontal meridian) () 30 !:::. 60

Modified Newtonian theory (vertical meridian) Modified Newtonian theory (horizontal meridian)

If t-~

corner~ radius ~~ Cone

2 3

C:(linde~

4 5 6

(a) a = 0° and 5°.

Figure 4.- Pressure distribution over blunt-nose cone at M 9.6. Configuration I (b) .

7

I\) --J

Cp

,1.5

1.0

5

a = 250 ; 0

,1.5

1.0

.5

a = 150 0

-.5 -7

Cylinder Cone

-6 -5 -4 -3 -2

Corner I I I I I ~ I I I ±t:t:Et±t Corner radius SpherlCal radius

nose

-1 0 1

sir

(b) ~ = 15° and 25°.

Figure 4.- Concluded.

1'J,deg

o jlO (vertical meridian) o 0 (horizontal meridian) <> 30 .6. 60

---- Modified Newtonian theory (vertical meridian) - - - ~ Mooified Newtonian_theory (horizontal meridian)

i 1"""1~

Cone Cylinder

2 4 5 6. 7

I\) ():)

Cp

1.5

1.0

. 5

a = 2.5°· 0 r

; .

o [}

o !::,. t::,.. .D o

¢, deg

90 (vertical meridian) o (horizontal meridian)

15 30 45 60 75 Modified Newtonian theory (vertical meridian) Modified Newtonian theory (horizontal meridian)

"1111111111111111111111111111111111111111111111111111!1111111111111111!lilllllllllllll!111111111111 1.0

.5

a = 0°. 0

Cylinder

-.5 -7 -6

9° cone 5.7° cone-·

-5 -4 -3 -2

Corner radius

-1 o

sir

(a) ~ = 0° and 2.5°.

Figure 5.- Pressure distribution over blunt-nose cone at M

'H gO cone

2 3 4 5 6 7

6.0. Configuration II(a).

£\) \D

Cp

1.5

1.0

.5

a ~ 7.50 0

L5

1.0

.5

a; 50 ,0

-.5 -7

Cylinder

-6 -5

o o <> t:, D. D o

'¢, deg

90 (vertical meridian) o (horizontal meridian)

15 30 45 60 75 Modified Newtonian theory (vertical meridian) Modified Newtonian theory (horizontal meridian)

corner_s h . al_corner p erlC di go cone 111111111'5.70 conectiradius" nose "ra us::: 5.70 cone:;t:j:::j::j::j gO cone

I, , Itll i! 11111 , Cylinder'

-4 -3 -2 -1 0 1 2 3 4 5 6 7

sir

(b) n = 5° and 7.5°.

Figure 5.- Continued.

\.),I o

Cp

1-.5

1.0

.5

a = 12.50 0

1.5

1.0

.5

• .,00 0 _II!~IIIIIIIIIIII ilill :lllllllllill;1111 c,+: T ••• 90

lr.... :}'"o~

I I 111I t II II II I I 1II1I II II 1111 -'~7 -6 -5 -4 -3 -2

Corner

g <> fd b­tl o

~l

~, deg

90 (vertical meridian) o (horizontal meridian)

15 30 45 60 75 Modified Newtonian theory (vertical meridian) Modified NeWtonian theory (horizontal meridian)

radiust. 5.70 cone~ I !~ Ii! ! 190

cone II1I ! ! II ~III \CYlind~r_ .~." f" .... ·1 \: 1;:.;:.1:::;: ·1"111 II i 1?0 -1 o 1 2 3 4 5 6 7

sir

(c) ~ = 10° and 12.5°.

Figure 5.- Continued.

\.)J ~

Cp

1.5

1.0

.5

a = 20° 0

1.5

1.0

.5

"" ,," , I~~~~ilili III1I11 ~~~i 111111111~:~ ~i·= -.5

-7 -6 -5 -4 -3 -2 -1

Spherical nose

o

sir

!1!, deg

o 90 (vertical meridian) o 0 (horizontal meridian) <> 15 t::,. 30 t>. 45 D 60 o 75

Modified Newtonian theory (vertical meridian) Modified Newtonian theory (hoiizontal meridian)

Corneril-n radius = 5.70 cone

2

go cone

4

Cylinder

5 6

(d) a = 15° and 20°.

Figure 5.- Contin~ed.

7

\)j £\)

Cp

1.5

1.0

.5

a = 300 0

1.5

1.0

.5

a = 250 o

-.5 -7 -.6

o o <> 6. b­e:::, o

~, deg

90 (vertical meridian) o (horizontal meridian)

15 30 45 60 75 Modified Newtonian iheory (vertical meridian) Modified Newtonian iheory (horizontal meridian)

-5 -4 -3 -2

Corner nose radius +5.7° cone ~ gO cone Cylinder

-1 o 1 2 3 4 6 7

sir

(e) ~ = 25° and 30°.

Figure 5.- Concluded.

\>I \>I

Cp

1.5

1.0

.5

a ~ 50 0

1.5

1.0

.5

a ~ 00 0

-.5 -7

Cylinder

-6

90 cone

-{i _4 '_3

Corner radius

5.70 COne .

-2 -1

Spherical nose

o

sir

1

(a) a ~ 0° and 5°,

Corner radius

Figure 6.- Pressure distribution over blunt-nose cone at M

0, deg

o 90 (vertical meridian) o 0 (horizontal meridian) o 30 b. 60

-----ModiiiedNewtonian theory (vertical meridian) Modified Newtonian theory (horizontal meridian)

5.70 cone go cone Cylinder

2 4 5, 6

9.6. Configuration II(b).

7

\.>I +"

Cp

1.5

1.0

.5

a ~ 25° 0

1.5

1.0

.5

a ~ 15° 0

-.5_'7

Cylinder 9° cone

-6 -,5 -4 -3

Corner 5.7° cone' radius

-2 -1

(b) a.

Spherical nose

o

sir

1

15° and 25°.

Figure 6.- Concluded.

dorner'

¢, deg

o 90 (vertical meridian) o 0 (horizontal meridian) o 30 6. 60

---- Modified"Newtonian theory (vertical meridian) Modified Newtonian theory (horizontal meridian)

radius f-j 5.7° cone 9° cone Cylinder

2 3 4 5 6 7

\).I \J1

2

2

oo°ct

sir

la) CL = OQ.

Figure 7.- Heat-transfer distribution over blunt-nose cone at M 6.0 and M 9.6. Configuration I.

\)oj 0'\

hZ lis,hemi

hZ hs,hemi

lO­

S

6

10-

lO­S

1

l

L

i 10--S

c

Cylinder

-7 -6

p B n b 0

IA

0

I~ 0 n 0

~

Cone

-5 -4 -3

IS(';!CC O©§o

n In ~ 0

-U

'" go

M 9.6 Configuration I{b)

R:E8§< §8~<t

:)

if'\ '-"

8 Q: ...D c-' u

~o 'Cb

M= 6.0

Configuration Ha)

Corner Spherical Corner radius nose radius

-2 -1

sir

(b) a. = 5°.

Figure 7.- Continued.

¢. deg

0 90 (vertical meridian) 0 o (horizontal meridian) = 0 15 -D, 30 -b. 45 -

D 60 -

G. 75 -

r 0 t 0 a 0 0 0

Fi n

,-,.

~ 0 ~

C ~

~ 0 0 0 0

Cone Cylinder

\)l -.:]

hZ hs,hemi

hZ hs,herni

¢, deg

1---0 90 (vertical meridian) ~LJ 0 (horizontal meridian)

10-

10-

1

8

10-

t---<> 15 t---b. 30 f---G. 45

D 60 f---O 75

l

~

I

i 10--8 -7

0 IV

C

Cylinder

-6 -5

B B h

~

Cone

-4 -3

q~(;) b§§~O

r\ 10 Q) o U~

h ~. v 0 0

c., 00 0 0

~n L

B g [J

r

M- 9.6

Configuration I(b)

0

lAne> IAP~ [

LD I"'" ~!' ~

0 ( 0 ~ 0 b ... - 0 0

~ IU [

0 ~ p 0

0 .r.

M 6.0

R5 Configuration I(a)

Corner Spherical corne:r r.radius nose radius Cone Cylinder

-2 -1 - -

sir

(C) a. = 10°.

Figure 7.- Continued.

\>I co

hZ hS,hemj

hZ lis,hemi

B

6

lO­

B

6

10-

2

1·

6

1O~ 8

4

l

1

..

~. deg

o 90 (vertical meridian.) ~ 0 0 (horizontal meridian) ~ 015 f-- L 30 I---'--- 1:0. 45

l\ 60 f--- 0 75

0

L

o 0 '--

Cylinder

'10-2 . -8 -7 -6 -5

0 P iA

-

" IT n 0 0 U

-

Cone

-4 -3

!;!l;'\ t:J ( p(;l~r -

In 0 n r n'"" 0 5 () ~ 0 U 0

is ~ !',. p

f5 Q 8 ~

R

-- - -

M 9.6

Configuration I(b)

f;)O ~Rr:-;[

I'" u" ~

;, ~ ~ 0 ~ 0 u

<.. D. p r L

IU 0 0

0 0

M= 6.0 F1 Configuration 1(al

Corner Spherical Corner l{radiUS nose radius Cone Cylinder -, , - I--. -,- , -

-2 -1 o sir

(d) 0.=15°.

Figure 7.- Continued.

\.)l \0

hZ hs,hemi

__ hZ_

hs,hemi

1

8

.---

1== === ---

0 0 () 6. to. D a

¢, deg

90 (vertical meridian) o (horizontal meridian)

15 30 45 60 75

_ - - - Crossflow theory

10-8

10-

1

8

10-

1

1

2L.-. 10--8

Q

c

'7

Cylinder

-7 -6 -5

~~.

0 0

~ .~

0 0 o u

Cone

-4 -3

~O ~ nAr=1 ,RB n

--p U -

0 Co ~

Rbi~ UR t. 0

A <; r:

(> D '" v 0 CI

0

--'

r':> M= 9.6 - f----

nU Configuration I(b) ,

Q,U

,i;LO~i2..

13 ~ v v

..&... r:,

Z'. 6. .L'-, -

0 wu

0 0

M- 6.0

Configuration I(a)

Co~er Spherical Corner

IIradius nose ..1 radius Cone Cylinder ~

-2 -1 o

sir

(e) a. = 25°.

Figure 7.- Continued.

g

hZ hs,hemi

10 8

6

4-

2

1

B

6

4

2

lO- l

B

6

4

2

l 10--8 -7 -6

0, deg

0 90 (vertical meridian) 0 o (horizontal meridian)

0 15 f':.. 30 t, 45 D 60 0 75

- - - - Crossflow theory·

f--I -,-

-5 -4 -3 -2

." U 0

~ PD P

AO~r '-.

L

~ U l'c

w

<v -- --~-~---.-- - i 0 I 0 [ oj I

0 0 - - ... _---

--- -_ •• _-1- •• ______

I .- ..j

M= 6.0 I Configuration I(a) ,

I I !

I

-_ .. ---t-Corner -I

I Spherical nose -radius Cone Cylinder I I--- .. _-- --+-r-- I

-1 o 2 3 4 5 6 7

sir

(f) a, = 300 •

Figure 7.- Concluded.

+:­.....

_h_l_ hs,hemi.

2

o 0

8

6

t-' 4

0

2 0 0

0

10-1 ;-.---. - .. t---

8

6 r-----r----- M = 9.6

Configuration IICb) 4

2

Corner Spherical nOse J radius 5.7° cone .J.

10-2 o

I' I 2

0

0 o c

pO v p 0 0

p 0 0 0

0 0

... ~ .• <;'2-.. _ r-.

J M= 6.0

Configuration lIla) l I

Spherical Corner

gO cone Cylinder. nose l[radiUS 5.70 cone gO cone Cylinder

1 I 1 -' \' 1 I I J - --L--4 6 o 2 4 7

sir sir

(a) a. = 0°. Figure 8. - Heat-transfer distribution over blunt-nose cone at M 6.0 and M 9.6. Configuration II.

-i=="" (\)

hZ hs,hemi

hZ hs,hemi

2-¢, deg

o 90 (vertical meridian) -~.O 0 (horizontal meridian) 1---0 15

1-

a

4·

lO­

a

4

2'

10-

:

:

1

,

1-

,-

1.0 -1

10 -2

1===6 30 G. 45

I---e:, 60 1---0 75

-8 --'I

v

Cylinder

-6 -5

b

~ " 0

gO Cone

-4 -3

@~ t:1 r:J ~

L

0

.R R O @I

~ ~O bJ P ~ e 'J

- 0 0 0 P

M= 9.6 Configuration II(b)

~ ---- - - -- ....... - '-----

a~R b@§~

PO U

!!i. ()

~ :J u

~ 0 0 0 0 0

(

M= 6.0 Configuration II(a)

Corner Spherical Corner 5.70 cone radius nose radius 5.70 cone "gO cone Cylinder

-2 -1 a 2 4

sir

(b) a, = 5°.

Figure 8.- Continued.

+" \.>I

hI

hs,hemi

hI

hs,hemi

0 --0 -0 I­

S

4

lO­S

10-

4

lO­S

i

c----;---t; '-'---b,. -f-,--D _~O

1

2

1

9l, deg

00 (vertical meridian) o (horizontal meridian)

15 30 45 60 75

v U

()

Cylinder 2 ___

10--8

.L -7 -6 -5

( E

~

gO cone

.L -4 -3

p ~ ( ~~i

_n 0

n n ~ 0 :;) ____ 0 '" !" ,

( Q g B' P ~ D 0

M 9.6 Configuration U(b)

~BI§( ~~~

n fT"T

" [ U ~ 0 0

0

~ 0 "S!JC i'X :J 0

0

M 6.0

Configuration Il(a)

Corner Spherical Corner 5.70 cone radius nose r- radius i5.70 cone gO cone Cylinder

.l I T I I -2 -1 2 4

sIr

(C) a = lOo.

Figure 8.- Continued.

+""" +"""

hZ

hs,hemi

hZ hsJhemi

4

10-

4

10-

1

8

4

10-1

8

1

2

¢, deg

o 90 (vertical meridian) t:=== 0 0 (horizontal meridian) 1-----0 15 I----- t; 30 I----- i::>. 45 I--D 60 1--0 75

,-,

v

Cylinder

10-2 -8 -7

I -6 -5

b E

0 C o ("

gO cone I

-4 -3

@~ t:1 ~ ~

0 C ~ r- () h 0 0

roB Bn ~ v t:i 0

L P ~

P <; P 0 ,

" M= 9.6

Configuration ll(b)

-- -- -'-

§'~I

r-. - n IU l? 0 () ~ ~

:~ u

'" 0

P ti

M= 6.0 Configuration II(a)

Corner Spherical Corner

5.70 cone radius nose radius , 5.7° cone gO cone Cylinder

I I' I I I -2 -1 0 2

sir

(d) a. = 15°.

FigUre 8.- Continued.

-!="" VI

hZ hs,hemj

nz hs,hemi

4

10-8

l

t= ~ I--!---------I--

0 0 0 £:, L-L 0

---

~-

10-

4-

1

8

10-8

i

l

i 10--8 -7

11l, deg

90 (vertical meridian) o {horizontal meridian}

15 30 45 60 75 Crossflow theory

§ E

0

t- ()

~I a 0

Cylinder gO cone

-6 -5 -4 -3

- - - - -

F.I F; M ~ ~ !" --{ R

~

H fJ [ b, D- O

DRe

RQ 6 ">

EI -=- u g C> 0 P :J

'lS M 9.6

n ( Configuration II(b) I

p

Q § ~ v

- ~ I=" -'J

!:-' ~

~ p-

o P

M= 6.0 , Configuration Il(a)

Co~ner Spherical Corner

5.70 cone ra us nose [,radius 5.70 cone gO cone Cylinder

-2 -1 o sir

(el a, = 25°,.

Figure 8.- Continued.

+"" 0'\

hi

hs,hemi

10 8

6

4-

2

8

6

4

2

10 8

6

4

2

10

-

1

2

-8 -7 ~

-6

¢, deg

0 gO (vertical meridian) 0 o (horizontal meridian)

0 15 6 30 b. 45 D 60 0 75

- -- Crossflow theory

--4 -3 -5 -2

.r-. 0

~§ ~ p ~ 0

r-.

' -

P f> ~

e,

0 p 0 ~ I

M= 6.0

,Configuration II(a)

, ,

Spherical Corner nose Ir-radius 5.70 cone gO cone Cylinder

\ -1 o 2 3 4

~ ~

6 7

sir

(f) a. = 30°.

Figure 8.- Concluded.

8

6

4

2

1

8

6

4

2

8

6

4

2

. i' I

i··r"

'I ,

: Sphe:"" "a1

-nose

IIII

: '.

,I.I I

, I

' .. J

:one

.'·,:Jc~,:t:,:,l· .: .

o Rr = 0.88 x 10~, Model I (a); M = 6.0 ORr = .54 X 10

6, Model I (a)i M = 6.0

ORr = .48 x 1061 Model I (b); M = 6.0 ~Rr = .,1 X 106' Model I,-{-b)i M = 6.0

Rr = ·07 x 10 , Model I (b)i M = 9.6

.:: t

, , ,.

I' 'i; , ,

I'

.. i

'., I-. :'1.'; " !

, "

, , !;"

, ,

linder' -w

10-2~ww~~~~~~~~~~~~~~ww~~~~~~~~ww~~~~~~~ o 1 2 4 5 6 7

sir

(a) Configuration I.

Figure 9.- Reynolds number effect on heat-transfer distribution at an angle of attack of 0°.

10 , ke,. I,,",:, h~'I~'"" """ :'?" I -IT II ,1"" I I" L"',;d,;,',',/';,', ' , ,,</'c,;,roc.'cf ,','<, ?i' ;'~,:,' ",'+"',"" ' o Rr = 0.9) x 10~, Model II

• ,.'! , :',::"'''' I:';:' :","',' (a)j M = 6.0 :c: ~~: ",c;~'C:=':'

c¢ CFf:: o ~ = .59 x 106

, Model II (a); M = 6.0 l <> Rr = .49 x 10" Model II (b); M = 6.0

:::-~' [f~ £::, Rr = .26 X l~, Model II (b)j M = 6.0 :

Rr = .08 x 10 , Model II (b); M - 9·6 hi:

11:!:(Ttl:i .,:1:1 :; ii': jiFlllj:f11 iii::!:!::: :±~:i:i tiC! +11 ,":-f:'

: e'l I,,~:' ,:,< 'i: I.i"l' ,', ",,,,,,e"",'" ",~,~"

8

6

4

,I, ::ii ii: i'"" ,"'ii', i:' ,':1:' i,ii if: 'rA:cE~cf0[ ~:' fT;Ii:!~ •• <,!:.,. :m~,: ±=

",'

'e,:: ' I : I i: j:i < jj{ ,-:,1, ,;-I

::'\ +:';-!: , 3:I:l +' 2 ,

I I I I

"

i , II

I i I

1 I I I ""~ Ii;''''-'' I ,-::'

8 1,1= I

I'S]' I , .' ,

hL 6 I;;,;,',

i I

p~ I : I

hs ,hem1 r. ,

: " , , ,

" 4 " ,

, i , , , ; ,

: ! I

, ,

i "

I , , , , , , 2 I I ,

" I ,

1 i 1

I I 1

"

8 ;

6 I

4 : i

I I

, ,

2 i ,

:;" :~' t~

:~;;~ rradius 5 • .,0 cone '00 cone: 14 ... ",,, ... ,

o 1 2 ) 4 5 6 7

sir

(b) Configuration II.

Figure 9.- Concluded.

48

+=­\0

hs

hs,hemi

Configuration II

hs

4

3

2

1

o

2

hs,hemi 1

Configuration I

o ~m

';-~~ -.~ -:--;::s:~,

- -.--7"--'- f -~

o .1 .2

+------.!.--i.­" I

i:

·3 .4 ·5 .6

Free-stream Reynolds number, Rr

o Model I (a), M = 6.0 o Model I (b), M = 6.0 II Model I (b), M= 9.6 <) Model II (a), M = 6.0 6, Model II (b), M = 6.0 A Model II (b), M = 9.6

1 +.-,

Correlation from reference 11

-.l

·7 .8 ·9 1.0 x 106

Figure 10.- Reynolds number effect Gn stagnation heating at an angle of attack of 0°.

2

1.0

8

6

4

hL/hs

2

8

6

4

2

I I I

(0. o M = 6.0, Rr = 0.88 x 106, Model I (a) o M = 6.0, Rr = .54 X 106, Model I (a)

I ' \ <> M = 6.0, Rr = .48 X ldS, Model I (b) __ DJ~ !:::,. M = 6.oA Rr = .31 X l~, Model I (b) -;VR_ • M = 9.6, Rr = .07 X 106, Model I (b) -

w: M = 6.0, Lees'theory based on Newtonian --c pressure distribution ---- M = 6.0, Lees t~eory based on measured

pressure distribution I ---M = 9.6, Lees theory based on Newtonian

pressure distribution II 1--, -- -- -....... lA. -....

(---- .............. -....... - -- ....... ~ ----,,;:;;--- -- -"'" - --

Ii r-j --... -0 ---- ----I-----==--~ 0 (}

, I',

A \ ~ ~ ~. I\. - ----

• A ~ 1---.- t---

;:.., ,., (=) I '-" '-' • I

A-

I v

: • w

" '-'

Spherical Corner nose r-radius

Cylinder I , I Cone L I I I

1 2 3 4 5 6 7

sir

(a) Configuration I.

Figure 11.- Comparison of heat-transfer distributions for ~ = 00 over the full Reynolds number range investigated with distributions predicted by the theory of Lees.

2 I I 6 I

o M = 6.0, Rr = 0.93 X 10 , Model II (a) OM., 6.0, Rr = .59 X 106, Model II (a)

,- <> M = 6.0, Rr = .49 X 1'1, Model II (b)

£ "t ~ M = 6.0, Rr = .26 X 106' Model II (b) - e M = 9.6, Rr = .oS X 10 , Model II (b) .1:1' ~ ~ M = 6.0, Leed theory based on Newtonian .....,. • {

pressure distribution - --- M = 6.0, Lee~ theory based on measured

t..

1.0

8

pressure distribution --.- M = 9.6, Lee~ theory based on 'Newtonian 6

~ pressure distribution

4 I I

I / ---- - 0 -:t.lt -- ---- -- .... ---- -- .0 r-I-- '-.

1/ --------til r---_--~ "--

bS-!-- ---~ ~ ~ p '------

~ @ ~

2

• 8 I-

6 V

r-.-=,--'---.

~ 4

r "

2

Spherical Corner I [radiUS 5.70 cone I 90 cone Cylinder nose

1 I I I I

1 2 3 4 5 6 7

sir

'(b) Configuration II.

Figure 11.- Concluded.

NASA-Langley, 1963 L-3368 51