NASA TECHNICAL

MEMORANDUM

NASA TM X-2955

iX

Fl

OPERATING CHARACTERISTICS

OF THE LANGLEY MACH 10

HIGH REYNOLDS NUMBER HELIUM TUNNEL

by Ralph D. Watson, Dana J. Morris,

and Michael C. Fischer

Langley Research Center

Hampton, Va. 23665"78-191*

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION • WASHINGTON, D. C. - JUNE 1974

1. Report No.

NASA TM X-29552. Government Accession No.

4. Title and Subtitle

OPERATING CHARACTERISTICS OF THE LANGLEYMACH 10 HIGH REYNOLDS NUMBER HELIUM TUNNEL

7. Author(s)

Ralph D. Watson, Dana J. Morris, and Michael C. Fischer

9. Performing Organization Name and Address

NASA Langley Research Center

Hampton, Va. 23365

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546

3. Recipient's Catalog No.

5. Report DateJune 1974

6. Performing Organization Code

8. Performing Organization Report No.

L-925910. Work Unit No.

501-06-08-0111. Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum14. Sponsoring Agency Code

15. Supplementary Notes

16. Abstract

Operating characteristics of the Langley Mach 10 high Reynolds number helium

tunnel are presented for stagnation pressures from 138 N/cm^ to 1655 N/cm^. The

characteristics include detailed Mach number surveys in the test section from which

usable core size and regions of disturbed flow were determined, preliminary blockagetest results, and maximum run time to be expected at various stagnation pressures.

Important tunnel dimensions including details of the model mounting apparatus are given.

Measurements show the variation in average core Mach number in the test section to be

between 9.4 and 10 for the present range of test conditions. The core radius is from

23 cm to 31.5 cm, depending on stagnation pressure and axial location in the test section.

1 7. Key Words (Suggested by Author(s) )

Tunnel calibration

Hypersonic flow

19. Security dassif. (of this report)

Unclassified

18. Distribution Statement

Unclassified - Unlimited

STAR Category 1120. Security Classif. (of this page) 21. No. of Pages

Unclassified 39

22. Price*

$3.25

For sale by the National Technical Information Service, Springfield, Virginia 22151

OPERATING CHARACTERISTICS OF THE LANGLEY MACH 10

HIGH REYNOLDS NUMBER HELIUM TUNNEL

By Ralph D. Watson, Dana J. Morris, and Michael C. Fischer

Langley Research Center

SUMMARY

Operating characteristics of the Langley Mach 10 high Reynolds number helium

tunnel are presented for stagnation pressures from 138 N/cm2 to 1655 N/cm^. The

characteristics include detailed Mach number surveys in the test section from which

usable core size and regions of disturbed flow were determined, preliminary blockage

test results, and maximum run time to be expected at various stagnation pressures.

Important tunnel dimensions including details of the model mounting apparatus are given.

Measurements show the variation in average core Mach number, in the test section to be

between 9.4 and 10 for the present range of test conditions. The core radius is from

23 cm to 31.5 cm, depending on stagnation pressure and axial location in the test section.

: INTRODUCTION

The Mach 10 leg of the Langley high Reynolds number helium tunnels is a unique

facility capable of simulating flight Reynolds numbers in the hypersonic range. For

example, figure 1 shows the Mach number-Reynolds number trajectory of a space shuttle

(33.7-m-long) configuration, the cruise trajectory of a hypersonic airplane (24.4 m long),

and the simulation range of the Mach 10 tunnel. On a relatively small 25.4-cm model,

the Mach 10 section of both trajectories can be covered in this tunnel simply by varying

stagnation pressure. Models up to more than a meter in length can be tested, the size

depending on the fineness ratio, nose bluntness, and angle-of-attack range desired. A

maximum Reynolds number of 222 x 10^ can be obtained on a model 1 meter long.

The facility was designed to operate at a maximum stagnation pressure of

2755 N/cm^ and stagnation temperatures up to 617 K. (See ref. 1.) Early problems with

a gas-fired heater were so great that the tunnels could be run efficiently only when the

heater was removed from the system. Thus, at present, the flow is unheated. Although

removal of the heater has made heat-transfer tests more difficult'since the model must

be cooled to produce low ratios of wall temperature to to.tal temperature, the overall

research capability of the facility has not been degraded. For example, instrumentation

problems are reduced when operating at ambient stagnation temperature for instruments

such as skin-friction balances, hot wires, and many types of pressure transducers. Also,

unheated Mach 10 helium flow is an excellent test medium to provide data for checking

viscous-shear-flow calculation programs currently being developed. The largest den-

sity ratio across the boundary layer, and thus the most stringent test cases, are obtained

at ratios of wall temperature to total temperature near 1.

Both the Mach 10 and Mach 20 legs of the high Reynolds number helium tunnels can

operate at stagnation pressures as high as 2755 N/cm^; however, it becomes impractical

to operate the Mach 10 leg at the higher stagnation pressures because of limited vacuum

sphere and high-pressure-supply volume. Operating experience has shown that a practi-

cal upper limit is about 1655 N/cm2; accordingly, the tunnel has been calibrated to thispressure.

This report.contains detailed Mach.number distributions in the Mach 10 leg for , . ,

stagnation pressures from 138 N/cm^ to 1655 N/cm^ which covers a unit Reynolds number

range from 10.8 X 106 to 133.6 X 10G per meter. Total temperature was between 230 K

and 322 K for these runs. In addition to determining the Mach number distribution in the

test section for a range of stagnation pressures, .preliminary blockage tests were made

at one location and the maximum tunnel run time was obtained.

; . . , _ . . . : . SYMBOLS; . . . _ . . . . . . . _ . .

D test-section diameter . . . • . . - • .

M . free-stream Mach number. . , . . . . . . • ;

M average Mach number

p pressure • . .

Rj Reynolds number based on length •

r radial tunnel coordinate .

rc inviscid core radius , .

T temperature

x axial tunnel distance, measured from nozzle throat

x axial tunnel distance, measured from station 186.61

6 boundary-layer pitot thickness

6* ' boundary-layer displacement thickness

(p angular coordinate (see fig. 6)

• iSubscripts: v

:

t total value

w wall . .

'oo' free-stream value

DESCRIPTION OF FACILITY

The Langley high Reynolds number helium tunnels consist of.Mach 10 and Mach 20

legs connected to common high-pressure storage tanks with a volume.of 56.6 m3 and to

vacuum spheres with a volume of 6526 m3 . A schematic diagram of the facility is shown

in figure 2. Operating characteristics of the Mach 20 leg can be found in reference 2.

Figure 3 shows the components and important dimensions of the Mach 10 leg. Infigure 3(a) the overall tunnel dimensions are shown from the pressure control valve to

the sphere isolation valve downstream of the diffuser. The diffuser is a fixed-area designhaving a ratio of minimum area to test-section area of 0.83. Details of the test sectionare shown in figure 3(b). A cross-section view shows the four removable plates that giveaccess to the test section. The two side plates are of optical-quality glass for schlieren

studies and all plates are interchangeable.

Models can be sting mounted on the hydraulically driven, programmable arc mecha-

nism or from the floor of the tunnel. Since the test section diverges 0.683° downstream,a set of leveling wedges has been attached to the floor of the. tunnel for strut-mountingmodels in any part of the test section. Figure 4 is a photograph of the test section.

Details of the axisymmetric contoured nozzle are shown in figure 5. For a stagna-tion pressure of 2758 N/cm2 and Tw = T£ = 294.3 K, the turbulent wall displacementthickness 6* was calculated by the method of reference 3. The displacement thicknessis added to an inviscid core calculated by the method of characteristics to give the wallcoordinates. The test section diverges at an angle of 0.683° downstream of the cutoffpoint. Also shown in the figure are details of the nozzle throat and experimental valuesof 6 in the test section determined from the data of this report.

INSTRUMENTATION AND DATA REDUCTION

Pitot surveys were made at four axial locations in the test section using the tworakes shown in figure 6. The larger rake contained 32 tubes on 8 arms plus a center-linetube. Tube spacing was 10.2 cm on this rake. The smaller rake contained 20 tubes on4 arms plus a center-line tube as well as 9 tubes arranged to provide a coarse pressuredistribution within the tunnel-wall boundary layer. Tube spacing on this rake was 2.54 cm.

Pressures were measured by using strain-gage transducers calibrated to one-quarterpercent accuracy of full-scale output. Different ranges of transducers were used to pro-vide maximum accuracy. It is estimated that all pressures are measured to better than1 percent of the true value. Tunnel total temperature was measured with a bare-wireiron-constantan thermocouple in the stagnation chamber.

Data were recorded on,tape in digital form. Mach numbers were calculated frompitot and stagnation pressure by two different methods. Data from the center tubes of the

small rake (see fig. 6) as well as all data from the large rake were reduced by assuming

isentropic nozzle flow. Mach numbers were calculated from equation (15) of reference 4.

Within the tunnel-wall boundary layer, isentropic expansion from stagnation conditions

cannot be assumed. On the boundary-layer arm, the average Mach number was deter-

mined by using static pressures calculated from core data along with the measured local

pitot pressure. By assuming the static pressure to be constant through the boundary

layer, Mach numbers were calculated by use of the Rayleigh pitot formula (eq. (16) of

ref. 4). For both reductions the real-gas corrections of reference 5 were used.

RESULTS AND DISCUSSION

At Mach numbers representative of the highest and lowest stagnation pressures of

the calibration and for values of T^ representative of these stagnation conditions, the

following flow parameters of interest have been calculated by using the real-gas correction

factors of reference 5:

Stagnation conditions:

pt, N/cm2 137.9 1585.9

Tt, K 289.9 255.6

Mach number 9.4 10

Static pressure, N/cm2 -. . 0.0393 0.3604

Static temperature, K 17.2 14.0

Dynamic pressure, N/cm2 2.892 30.037

Velocity, m/sec . . ' 1708 1640

Unit Reynolds number/m 10.8 X 106 133.6 X 106

Starting transients in the high Reynolds number helium tunnels last about 1 second.

During this time the stagnation pressure changes from tunnel vacuum to the desired level

smoothly with no significant overshoot. However, the total temperature momentarily

reaches a higher level during the starting process. Sketch (a) shows a simple schematic

of the high-pressure'reservoir, pressure control valve, stagnation chamber, and nozzle.

Helium is throttled from a typical pressure of 3450 N/cm2 in the supply line to a lower

pressure in the stagnation chamber. The Joule-Thompson effect at the pressure-control

valve produces a slight increase in TJ-, about 15 K; however, a much larger effect is

produced as the stagnation chamber is filled. In reference 6 this thermodynamic process

is shown to produce a ratio of final to initial temperature equal to 1.67 for helium.

High-pressure storage / Stagnation chamber(3450 N/cm2) '

x_ Iy "\(X) ^

V /Pressure — i

control valve

A J

Nozzle

To vacuumspheres

Sketch (a)

Figure 7 shows typical total-temperature histories for two different stagnation

pressures. The data were recorded on a system containing filters to remove 60-Hz noise;

thus, the true peak in Tt is not recorded. The large overshoot is evident as well as the

rather slow return to near-ambient total temperature. During the run significant mass can

be removed from the storage bottle; as a result, an almost linear decrease in total tem-

perature with time is produced as the remaining helium expands. Despite the continual

change in total temperature during a run, no time variation in free-stream Mach number

could be detected at constant stagnation pressure. .

Mach Number Distribution

Mach number surveys-were made at x locations of 25.4, 76.2, 127, and 177.8 cm

for nominal stagnation pressures of 138, 276, 414, 827, 1241, and 1655 N/cm2. Data from

the small rake are plotted in figure 8 as radial Mach number distributions. In general,

the figure shows the flow to be symmetrical about the tunnel center line with disturbances

in the immediate vicinity of the center line. The magnitude of the disturbance varies with

axial location in the test section and with tunnel stagnation pressure. The disturbance

region extends radially about 5 cm, although in most cases, the Mach number deviation is

small beyond 1 cm. Measured pitot pressures were lower on the center line than for

adjacent tubes, and produced peaks in Mach number when reduced, isentropic flow in the

core being assumed.

Similar disturbances in a Mach 8 air tunnel (ref. 7) were traced to imperfections of

about 0.013 cm in the nozzle wall. The effect whereby wall-induced disturbances in axi-

symmetric flow increase in magnitude toward the center line (the "radial focusing effect")

is discussed in reference 8.

Center-line disturbances were stated in reference 9 to have negligible effect on the

overall forces on a test body and little effect on the pressure distribution over slender

bodies. Pressure and heat-transfer distributions on blunt bodies were significantly

affected by disturbances when tested on the center line. The same trends will undoubtedly

be found in the high Reynolds number Mach 10 tunnel and should be considered when the

models are located in the test section.

Average Mach numbers were obtained by fairing the data of figure 8 between radii

of 5 and 15 cm from the center line. Variation of the resulting average Mach number

with tunnel stagnation pressure is shown in figure 9. The average core Mach number

increases approximately proportional to the logarithm of the stagnation pressure.

Figure 10 shows the Mach number variation with distance x in the test section. At a

fixed stagnation pressure the Mach number change from the front to the rear of the test

section is less than 0.16.

Data from the large rake are shown in figure 11 along with core data and boundary-

layer arm data from the small -rake. The outer tubes of the large rake are within the

tunnel-wall boundary layer, and thus data for these tubes are omitted. Note that the lower

and upper plots of each figure are not to the same scale. Data from the lower part.of

this figure were used to find the edge of the usable test core for each run. The core

radius at each of the four survey stations is shown in figure 12 and figure 5.

Blockage Test

In order to determine.the approximate model sizes which can be tested in the tunnel,

starting tests with hemisphere cylinders of various diameters were made at ~x. about

70 cm. The floor of the test section and the diffuser sidewall were instrumented with

pressure transducers to determine whether potentially dangerous static pressures

occurred downstream of the nozzle when an unstarted tunnel condition existed.

Tests were begun with a 24-cm-diameter model; the tunnel would not start at stag-

nation pressures of 138 N/cm2 and 276 N/cm2. The model size was reduced to 21.6 cm

and blockage again occurred at these stagnation pressures. The model diameter was

reduced to 18 cm and again the flow did not start. A further reduction of the model

diameter to 12.4 cm allowed the tunnel to start at a stagnation pressure of 140 N/cm^.

Figure 13 shows the model to scale in the test section along with pressures measured on

the floor of the test section. The measured edge of the inviscid core from figure 11 is

shown as well as shock shapes calculated by the method of reference 10 for spherical-nose bodies.

It was impossible to determine whether the flow was started from schlieren observa-tions of the shock. The flow field appeared to be steady even when the flow was known to

be blocked. From figure 13, wall pressures are indicative of a Mach 5 start, and it canbe seen that calculated Mach 5 and Mach 10 shock shapes are not greatly different. Itappears that when model blockage is too high, the tunnel-wall boundary layer separatesupstream to the point where approximately Mach 5 flow is reached. When the tunnel flow

is started, the wall static pressures are in agreement with the expected level./

/

Tunnel Run Time

By using the hemisphere-cylinder/described in the preceding section, tunnel run

time was established by running until natural flow breakdown occurred because of highbackpressure in the vacuum spheres or until stagnation pressure dropped because oflimited supply pressure. Figure 14 shows the results of these tests. Also shown is thecalculated time limit at which the two vacuum spheres would reach atmospheric pressure,a condition at which automatic tunnel shutdown occurs since the spheres were not designedfor an internal pressure above atmospheric.

At a stagnation pressure of approximately 300 N/cm^, a sharp decrease in tunnelrun time is evident, probably caused by a change in flow conditions within the diffuser.Above 300 N/cm^, the tunnel run time decreases steadily with increasing stagnationpressure. At about 1200 N/cm2, the pressure contained in one supply bottle is not suffi-cient to maintain constant stagnation pressure until flow breakdown is reached. Anothersupply bottle is available for high-pressure storage; however, it is usually not fully pres-surized and is used to store impure helium for later purification. Optimum operation ofboth Mach 20 and Mach 10 tunnels is achieved by this arrangement.

CONCLUDING REMARKS

The Mach 10 leg of the Langley high Reynolds number helium tunnels has been cali-brated for stagnation pressures from 138 N/cm^ to 1655 N/cm2. Mach number distribu-tions at four positions within the test section are presented along with the results of a

flow-blockage test and the determination of tunnel run time within the range of calibration

stagnation pressures.

In summary, four general characteristics of the flow can be noted as follows:

1. Pitot surveys indicate an average core Mach number variation from 9.4 to 10 at

corresponding unit Reynolds numbers of 10.8 X 106 to 133.6 X 106 per meter.

2. The radius of the inviscid core ranges from 23 to 31.5 cm depending on stagnation

pressure and test-section location.

3. Tunnel run time varies from about 16 seconds to about 3 seconds, depending on

stagnation pressure.

4. A preliminary blockage test at one test-section location has shown that a spheri-

cally blunted, 12.4-cm-diameter body will start at a tunnel stagnation pressure of

138 N/cm^; a model 18 cm in diameter will not start at this pressure.

Langley Research Center,

National Aeronautics and Space Administration,

Hampton, Va., February 13, 1974.

REFERENCES

1. Schaefer, William T., Jr.: Characteristics of Major Active Wind Tunnels at the

Langley Research Center. NASA TM X-1130, 1965.

2. Watson, Ralph D.; and Bushnell, Dennis M.: Calibration of the Langley Mach 20

High Reynolds Number Helium Tunnel Including Diffuser Measurements. NASA

TM X-2353, 1971.

3. Persh, Jerome; and Lee, Roland: .A Method for Calculating Turbulent Boundary

Layer Development in Supersonic and Hypersonic Nozzles Including the Effects of

Heat Transfer. NAVORD Rep. 4200 (Aeroballistic Res. Rep. 320), U.S. Navy,

June 7, 1956.

4. Mueller, James N.: Equations, Tables, and Figures for Use in the Analysis of

Helium Flow at Supersonic and Hypersonic Speeds. NACA TN 4063, 1957.

5. Erickson, Wayne D.: Real-Gas Correction Factors for Hypersonic Flow Parameters

in Helium. NASA TN D-462, 1960.

6. Dodge, Barnett F.: Chemical Engineering Thermodynamics. McGraw-Hill Book

Co., Inc., 1944.

7. Fitch, C. R.: Flow Quality Improvement at Mach 8 in the VKF 50-Inch Hypersonic

Wind Tunnel B. AEDC TR-66-82, U.S. Air Force, May 1966.

8. Meyer, R. E.: The Method of Characteristics for Problems of Compressible Flow

Involving Two Independent Variables. Part II. Integration Along a Mach Line.

The Radial Focusing Effect in Axially Symmetric Flow. Quart. Jour. Mech. and

Appl. Math., vol. I, pt. 4, Dec. 1948, pp. 451-469.

9. Sivells, James C.: Aerodynamic Design and Calibration of the VKF 50-Inch Hyper-

sonic Wind Tunnels. AEDC-TDR-62-230, U.S. Air Force, Mar. 1963.

10. Lomax, Harvard; and Inouye, Mamoru: Numerical Analysis of Flow Properties About

Blunt Bodies Moving at Supersonic Speeds in an Equilibrium Gas. NASA TR R-204,

1964.

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