Document
7"
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AIAA 78/-587R
Improving Large-Scale Testing Capabilityby
Modifying the 40-by 80-ft Wind Tunnel
K. W. Mort, P. T. Soderman, W. T. Eckert
Reprinted from Volume 16, Number 8, August 1979, Page 571.
Journal ol Aircraft This paper is declared a work of the U.S. Government and therefore is in the public domain.
AMERICAN INSTITUTE OF AERONAUTICS AND ASTRONAUTICS • 1290 AVENUE OF THE AMERICAS • NEW YORK, NEW YORK, N.Y. 10019 VOL. 16, NO. 8, AUGUST 1979 J. AIRCRAFT 571 ARTICLE NO. 77-587R
208179
NASA/TM, - 7q'-"
Improving Large-Scale Testing Capability by Modifying
the 40- by 80-ft Wind Tunnel
S_
Kenneth W. Mort," Paul T. Soderman,t" and William T. Eckertl" NASA Ames Research Center, Moffett Field, Calif.
lnteragency studies conducted during the last several years have indicated the need to improve full-scale testing capabilities. The studies showed thai the most _ffective trade between lest capability and facility cosl was provided by repowering the existing Ames Research Center 40- by 80-ft Wind Timnel to increase the maximum speed from about 100 m/s (200 knots) to about 150 m/s (300 knots) and by adding a new 24- by 37-m (80- by 120-ft) lest section powered for about a 50-m/s t100-knot) maximum speed. This paper reviews the design of the facility, a few of its capabilities, and some of its unique features.
Introduction flight-test hardware. Experimental investigations in the 40- by 80-ft wind tunnel have prevented catastrophic in-flight HE experience and insight gained by the successful failures of a number of advanced V/STOL research aircraft operation of the 40- by 80-ft wind tunnel over the past by exposing, during the wind-tunnel test, unanticipated three decades have shown both the value of the existing deficiencies in critical components (rotor, propulsion, or facility and the requirement for even greater test capabilities.
control systems). As a result, substantial savings have accrued Current and planned development of V/STOL aircraft and when tests in the 40- by 80-ft wind tunnel exposed deficiencies advanced helicopters have increased the need for a facility that were either fundamental to the concept or would have with a larger test section and higher test speeds. The planned required a prohibitive increase in the program cost to correct, modifications to the 40- by 80-ft wind tunnel at Ames The overall savings accrued have more than offset the con- Research Center will consist of increasing the power of the struction and operating costs of this facility for the last 35 drive motors and adding a new test section.
years, in addition, the research contributions this facility has This paper discusses the background of and justification made to such advanced aircraft concepts as tilt rotor and lift for the project, reviews the improved test capabilities for fan VTOL aircraft, advanced helicopters,-lifting-body model size and test speed, discusses the options that have been aerodynamics, etc. (although difficult to quantify in dollars) considered as possible enhancements to the aerodynamic test are unquestionably of far greater value than the appreciable capabilities, explains the improvement in the acoustic test money saved by ground testing rather than flight testing the capabilities, outlines some of the unique features of the design aircraft.
of the facility, and reviews the current status of the project.
In 1967, the Aeronautics and Astronautics Coordinating Background Delails Board (AACB) initiated studies of ground-based facilities in the U.S. The AACB's objective was to define those areas The value of full-scale subsonic testing has been con- where the limitations of existing facilities could be expected to vincingly demonstrated by the NASA 40- by 80-ft wind tunnel limit the performance or acceptability of future aircraft. The over the past 35 years. However, because of the increase in first step in this study was to review the degree to which size of aircraft and the technological changes in aircraft, existing facilities had met the test requirements of past air- facility capabilities for performing this type of testing must be craft development programs. Table I presents such a review improved (see, for example, Refs. 1-6). To meet this need, and shows how the current speed and size limits of the 40- by NASA is planning two major modifications to the 40- by 80-ft 80-ft wind tunnel have influenced various aircraft test wind tunnel. Figure 1 shows the modifications planned. The programs. As may be seen from the table, the size limitation first modification consists of repowering the drive, which will has primarily constrained the fixed-wing aircraft in- increase the maximum speed in the 12- by 24-m (40- by 80-ft) vestigations, while the speed limitation has primarily con- test section from about 100 m/s (200 knots) to about 150 m/s strained the rotary-wing investigations. The test constraint (300 knots). The second modification is the addition of a new due to size is caused by two basic factors: first, the long-term rectangular test section that will be about 24-m (80-ft) high by increase in aircraft size and, second, the continuing need to about 37-m (120-ft) wide. The maximum speed in the new test develop more lift from a wing of a given size. This latter section will be about 50 m/s (100 knots).
Past experience with the existing 40- by 80-ft wind tunnel has demonstrated the value of adequate ground testing of Table I Past programs affected by 40- b_, 80-ft wind funnel limits advanced aircraft (or critical components of these aircraft) before making the large financial investment required for Limit Program Size Speed Presented as Paper 77-587 at the AIAA/NASA Ames V/STOL C-8 Augmentor Wing X Conference, Moffett Field, Calif., June 6-8, 1977; submitted June 16, F-14 X 1977; revision received Jan. 15, 1979. This paper is a work of the U.S. F-15 X Government and therefore is in the public domain. Reprints of this F-Ill X article may be ordered from AIAA Special Publications, 1290 Avenue X Electra (prop. whirl) X of the Americas, New York, N.Y. 10019. Order by Article No. at top US/FRG fighter X of page. Member price $2.00 each, nomnember, $3.00 each. X AH-56A comp. helicopter X Remittance must accompany order.
ABC rotor X X Index categories: Testing, Flight and Ground; Subsonic Flow; X Bell H .P. helicopter Acrodynamics. X XH-51A comp. helicopter "Research Scientist. Member A IAA.
XC. 1421iIt wing X "i'Rcsearch Scientist, U.S. Army Aviation R&D Comrnand, X-22 tilt duct X Research and Technology 1 aboratoric,,, Aeromechanics I.ahoralory.
572 MORT, SODERMAN, ANDECKERT J.AIRCRAFT
factor requires a larger testsection (relative to thewing) to
as a function of test-section speed. As shown, the greatest alleviate wind-tunnel wall-constraint effects onthecomplex ratio or potential capability would result from using a self-
flowfields thataretypical of high-lift conditions. Thetest
correcting test section. A vented test section (passive venting)
constraint dueto speed is primarily caused by theneed to
with a residual error of 0.5 deg in angle of attack results in improve the capability and efficiency ofadvanced rotorcraft.
approximately the same test capability as solid walls with an error in angle of attack of about 2 deg. Wall effects of this magnitude are correctable to within that tolerance, thus, little Test Capability Improvement is gained by venting the test section. Despite the potential Aerodynamic Testing superiority of self-correcting test sections and vented test The increased capability in model size and speed which will sections, it was decided not to implement either scheme.
result from the planned modification is shown in Fig. 2.
Self-correcting test sections have not been developed for Allowable model size is shown as a function of test-section multienergy or powered-lift models. Experimental work to speed. The size boundary (developed using Ref. 7) is shown as date has been on two-dimensional models (see, for example, a band because the allowable size varies considerably, Refs. 9-12). There have been limited three-dimensional depending on the details of the model, the purpose of the theoretical studies on very simple model configurations (see, tests, etc. At low speeds, where wake constraint effects caused for example, Refs. I1 and 13). The judgment has been made, by the wind-tunnel walls reduce the allowable model size, the after considering these studies, that implementation of self- addition of the new test section will cause a substantial in- correcting techniques may ultimately be practical at small crease in the size capability. This increase in size capability scale for three-dimensional powered-lift testing; however, will substantially improve the capability for testing V/STOL implementation at full scale would require very complex and and powered-lift aircraft, as well as rotorcraft, at low speeds.
expensive systems and does not appear practical. In addition, The increase in speed planned for the existing test section (also flow-measuring instrumentation and computer capability shown in Fig. 2) will allow more representative studies of would be required for full-scale implementation. The variety high-performance rotorcraft than presently possible.
of model configurations, range of model sizes, and model It is appropriate at this point to review the aircraft (listed on support requirements required for full-scale testing would Table I) for which the size and speed capabilities of the also substantially increase the complexity.
existing 40- by 80-ft wind tunnel are inadequate, and to assess Test sections with passive venting for powered-lift testing the adequacy of the modified 40- by 80-ft wind tunnel to meet were studied in Refs. 14-21. Large plenum chambers appear the requirements of these aircraft. The modified facility necessary (although Ref. 19 describes promising results with would eliminate all of the wind-tunnel speed and size only small floor and ceiling plenums). Even though venting limitations for testing these aircraft with the exception of the attenuates wall effects, there are residual errors. To date, limitations associated with the Electra propeller whirl techniques for correcting these errors have not been nearly as problem, which required a combination of test section speed accurate or reliable as the application of corrections for solid- and size that is not provided by the modified 40- by 80-ft wind wall effects. As Fig. 3 shows, the solid wall with a 2-deg error tunnel. From this assessment, as well as consideration of long- (or correction) in angle of attack is essentially equivalent to term trends in aviation and the associated technical problems, vented walls with a 0.5-deg error in angle of attack, but the 2- it is believed that a good cost/benefit judgment has been deg error is correctable, while the 0.5-deg error is not. In achieved with the proposed modifications to the 40- by 80-ft addition, the venting requirements appear to depend on the wind tunnel.
model geometry and, for a given model, can depend on model attitude and lift. Such highly variable venting requirements Aerodynamic-Testing Enhancement Considered demand a complex, automated venting system. An additional Studies were made to examine the possibility of enhancing drawback, although less serious, is that venting increases the the test capability of the new test section. The goal was to wind-tunnel drive power required; it is estimated that venting eliminate or reduce the aerodynamic measurement errors and would increase the power for the new test section by about 7 uncertainties caused by the test section walls so that larger to 8°7"0. Because of these considerations, it was decided not to models could be tested. "Self-correcting" arrangements, such implement vented walls.
as those of Refs. 8-13, and vented-wall configurations, such It is likely that, over the next few years, wall correction as those of Refs. 14-21, were considered. Figure 3 illustrates techniques will be improved so that larger models can be some results for a uniformly loaded wing with a ratio of wing- tested. The potential gain in size (using the techniques in Ref.
loading to aspect-ratio of 958 N/m -_ (20 Ib/ft 2). The ratio of 7) is illustrated in Fig. 4. As can be seen, if large corrections wingspan to test section width for a given correction is shown can be made reliably, the testing capability compares 12 x 24-m _:_.:._ 5;!_':_':: ,_._:;_:::":_!i IMPROVED FACILITY 120 - (40 x 80-ft) TEST SECTION 100 m/s_150 m/$ 3O (200 knots-300 knots) Z _ ...... _:_'_:_:':_'_'?i_: < 80 - 25 24 x 37-m ft m 20 (80 x 120-ft) re TEST SECTION U 15 E 50 m/= < 40 - (100 knots) _i.!:! !i : PRESENT 1 [i 27 MW_100 MW ' 40x 80-ftl WIND TUNNELI i I_ (36,000 hp-_135,000 hp} 0 - 0 50 100 150 m/s I I I I 0 100 200 300 knots TEST SPEED Fig. I 40- h) 80-ft '_,ind tunnel modifications.
Fig. 2 Improvement in full-scale test capabilil).
AUGUST 1979 LARGE-SCALE WIND TUNNEL TESTING CAPABILITY 573 -- ANGLE-OF- QUARTER 1.0 SELF-CORRECTING ATTACK CHORD FLAP WITH BOUNDARY- "1" _ ERROR, I LAYER CONTROL
\
0.1% TO 1% OF
,.0
WING CHORD BLOWING JET_ _ BOUNOARV-LAYER RAKE I,-, , I I ._ APPROX.
0 25 50 2 CHORDS m/s Fig. 5 Ground-effect testing using floor blowing.
I I I 0 50 100 knots TEST SPEED Fig. 3 Potential enhancemenl of test capabilily at a wing loading/aspect ralio of 958 N/m z (20 Ib/ft z ).
I.-- Z ttl I" _ CORRECTIONS u_6 1"01 SELF-CORRECTING t,_ tK.
_ MODERATE w a. Z o o,> 00.5 cJ 4 k- zu t_ SMALL ..J
aa, 6;Jb2
Ip- _\\\\\\"_. ACCEPTABLE : 0 25 50 rn/s 0 0.4 0.8 1.2 1.6 2.0 2.4 l I I TRAILING-EDGE HEIGHT/WING CHORD 0 50 100 Fig. 6 Test ranges for fixed ground, biotin g jet, and mo_ing bell.
knots TEST SPEED Fig. 4 Potential of impro_,ed .all ¢orreclions at a wing Ioading/aspecl ratio of 958 N/m 2 (20 Ib/ft 2 ).
for powered-lift models and is substantially simpler to im- plement than a moving belt system. In view of this, studies on the blowing jet are continuing to be sponsored with the ob- favorably with that for self-correcting concepts in the 30 to 50 jective of eventual implementation in the new test section.
m/s test-speed range, which is the range of interest for powered-lift or STOL configurations. (Test speeds below 30 Acouslic Testing m/s will probably require, for practicality and economy, In addition to the improvement in aerodynamic test approaches to error reduction other than analytically capability to be achieved by the modification, the acoustic test determined wall corrections, self-correcting test sections, or capability will also be substantially improved. Over the past 5 standard venting techniques.) The techniques for determining to 10 years, aeroacoustic research has become increasingly large corrections remain to be developed. However, it is important, and many facilities have been modified or con- expected that this will be done in the next few years. In ad- structed for this work (see Ref. 5). in the existing 40- by 80-ft dition, it appears likely that, if the instrumentation and wind tunnel many experimental investigations of aeroacoustic computer techniques being developed for self-correcting test phenomena have been performed (see, for example, Ref. 27).
sections were employed for normal test sections, large The capability for performing acoustic studies in the modified corrections could be determined accurately by proper 40- by 80-ft wind tunnel will be improved by reducing the test measurements. This means of determining wall corrections section background noise as well as by increasing the test coltld be implemented in the new test section whenever the speed and test section size.
technology is developed.
The new drive system will be much quieter than the present drive system, even though the power will be increased to Steady-State Ground-Effect Testing nearly four times that of the present system. In Fig. 7 the new For a number of years, NASA Ames has been sponsoring drive background noise is compared with the present level. As studies on the use of a blowing jet to energize the boundary shown, there will be a substantial reduction in background layer on the floor of the test section for performing ground- noise. The drive noise will be lowered by reducing the fan effect testing (Refs. 22-26). Figure 5 illustrates the concept inflow distortion and by employing a fan with a much lower schematically. The jet is approximately two wing chords tip speed. The fan tip speed will be 115 m/s (337 ft/s) com- forward of the wing. The condition of the boundary layer is pared with 185 m/s (607 ft/s) for the present fan. The fan continuously monitored under the wing at the quarter-chord inflow will be improved by changing the fan arrangement as so that the jet may be properly set and adjusted as required. shown in Fig. 8. The fan will be located upstream of the The capability for performing ground-effect testing is support struts instead of downstream. The new drive requires illustrated in Fig. 6, which is reproduced from Ref. 25. Lift a set of stators which will be located at least two rotor-blade coefficient is shown as a function of the ratio of flap trailing- chords downstream of the rotor to minimize interference and, edge height to wing chord. The suggested test ranges for fixed hence, noise.
ground, blowing jet, and moving belt are shown. As can be The drive system was developed with the aid of studies seen from the figure, the blowing jet is substantially better performed on 1.8-m- (6-ft) diam fan models. The model noise than the fixed ground board, but not as good as the moving was scaled and increased to represent the noise of the six full- belt. However, the capability of the blowing jet is acceptable scale drive fans and then compared with the noise of the
574 MORT, SODERMAN, AND ECKERT J. AIRCRAFT
160 -- PRESENT_ o 150-- _PRESENT DRIVE FANS
0"- .-J ANS
- 1 / v/ PROPOSED LOW-SPEED _140-- >
130 ..... REPOWERED FANS
n,- w • "_, "'"",\'J_c HIGH SPEED _O120 D.
- --"_"_ LOW SPEED Q I I I _ 110 50 100 150 IMPROVED LOW SPEED m/s I I I I I I I I I I I 11111 lg 63 250 1000 4000 100 200 300 31.5 125 500 2000 knots ONE-THIRDOCTAVEBANDCENTERFREQUENCIES. Hz TEST SPEED Fig. 9 Drive sound power level at 100 m/s in 12- by 24-m (40- by 80- Reduction in test-section background noise.
Fig. 7 ft) test section.
---- PRESENT FAN LOUVER: PROPOSED FAN PROPOSED PRESENT STAIOR PROPOSED LOUVERS_.,J [,-_--_11_ _ I[ 1 //, _ % i •.. .... J/ .... "'_ (40) / I QUIET ACOUSTIC I _,_"/ I DRIVE TREATMENT/
\'%, _" I FLOW / : /-
( I DIVERTERS _'_/ (143) DIMENSIONS IN m fit) MOTOR SUPPORT STRUTS Fig. 8 Change in drive s)slem geometry.
O FLIW STRAIGHTENERS ACOUSTIC TREATMENT present drive fans. The results are shown in Fig. 9. Sound Fig. l0 Special features.
power level is shown in one-third octave bands for a test- section speed of 100 m/s (200 knots). The lower three curves were scaled from the model studies. Three fans were tested: a high-speed fan (tip speed 191 m/s, 627 ft/s) and two low- of the generally light wind conditions at the site. Critical low- speed fans (tip speed 115 m/s, 377 ft/s). The two low-speed speed testing will be scheduled for low-wind conditions which normally exist about 8 h each day. Many experimental in- fans were quieter than the high-speed fan, and the improved low-speed fan was quietest. The improvements were made by vestigations have been conducted to define the inlet treatment improving the rotor and stator airfoil shapes and by im- required to achieve satisfactory flow characteristics in the test proving the quality of the fan inflow. section under all wind conditions (see, for example, Ref. 28).
The model development process was important to the The present design does not incorporate all of the technology optimization of the drive system. The resulting drive system developed during these experiments, because, from a will be significantly quieter than the present drive system. cost/benefit viewpoint, to specify that the 24- by 37-m (80- by Since the existing facility is used extensively for acoustic 120-ft) test section must have excellent flow quality at low studies, the reduction in background noise will be a very speeds under all conditions of external winds did not appear important improvement. to be justified. If future use of the facility shows that test scheduling is being seriously curtailed by external wind conditions, the technology for additional inlet treatment is Special Features available.
Several special features of the modified facility are shown The flow straighteners will be lined with acoustic treatment i_n Fig. 10. To control the airflow in either the closed return or to lower the wind-tunnel noise in the surrounding community the nonreturn circuit, there are various louvers and flow (see Ref. 29). The exhaust acoustic treatment will be in both deflectors. The louvers at the locations indicated on the figure the nonreturn and closed-return circuits (as shown in Fig. 10), are relatively simple, two-position devices that either open or and will reduce fan and model noise propagation out of the close the flow passage. However, the two sets of flow wind tunnel (in the nonreturn mode) and fan noise diverters shown are more complicated. The set at the in- propagation into the present test section (in the closed-return tersection of the two circuits is the most complex and has been mode).
under extensive study. As shown, the forward portion of the device is in one of two positions, depending on the circuit Status being used. Not only is it important to minimize the power Various studies have been used to develop the design of the loss due to these deflectors, but it is important to minimize facility. Model tests were used to develop the flow diverter their wake so that its effect on fan noise is negligible. The set of diverters near the exit is less complex and consists essen- and louver systems, inlet and exit systems, and the drive tially of straight-sided extensions of the present turning vanes system. The diverters and louvers were studied at about 1/10 supported by a pivot system. scale, the inlet and exit systems at 1/50 scale, and the drive As shown in Fig. 10, flow straighteners at the inlet reduce system at about I/7 scale. Additional studies are being the effects of crosswinds on the flow quality in the test sec- performed on test-section systems, on acoustics, and on the tion. This minimal treatment appears to be acceptable because use of a blowing jet on the floor of the test section for ground- AUGUST 1979 LARGE-SCALE WIND TUNNEL TESTING CAPABILITY 575 effect testing. Tests for optimization of the acoustic treatment t l Bernstein, S. and Joppa, R.G., "Reduction of Wind Tunnel Wall Interference by Controlled Wall Flow," NASA CR-2654, March are underway, as well as a study of inflow turbulence effects 1976.
on fan noise. Final design on the modification began in 1974 t2Berstein S. and Joppa, R.G., "Development of Minimum and should be completed by mid-1979. Construction should Correction Wind Tunnels," Journal of Aircraft, Vol. 12, April 1976, be completed by mid- 1981.
pp. 243-247.
13Erickson, J.C., Jr., "Application of the Adaptive-Wall Concept Concluding Remarks to Three-Dimensional Low-Speed Wind Tunnels," NASA CR- A review of aircraft development trends, as well as ex- 137,917, Sept. 1976.
laCull, M., "Progress Report on Experimental Program to In- tensive operational experience with the existing 40- by 80*ft vestigate Various Test Section Configurations in Conjunction with wind tunnel, indicate the need for a major increase in the Development of Boeing V/STOL Wind Tunnel," The Boeing speed and size capability of this facility. The planned Company, Vertol Division, Morton, Pa, D8-0696, June, 1967.
modifications will increase the maximum speed of the 12- by tSTemplin, R.J., "Recent Trends in Low-Speed Wind-Tunnel 24-m (40- by 80-ft) test section from 100 to 150 m/s (200 to Design and Techniques," Proceedings of the International Congress 300 knots), and will provide a second test section 24 by 37 m on Subsonic Aeronautics, Annals New York Aeadenty of Sciences, (80 by 120 ft) with 50 m/s (100 knots) speed capability for Vol. 154, Art. 2, Nov. 22, 1968, pp. 1055-1073.
low-speed, powered-lift, and rotorcraft testing. Design studies ]6Binion, T.W., Jr., "An Investigation of Several Slotted Wind have shown that in spite of the large increase in wind-tunnel Tunnel Wall Configurations with a High Disc Loading V/STOL Model," Arnold Engineering Development Center, Tenn., AEDC- drive power, the use of a low tip speed fan will result in a drive TR-71-77 (AD 723294), May 1971.
that is quieter than the existing drive. These modifications will IVGrunwald, K.J., "Experimental Investigation of the Use of enhance the capability of the modified facility to perform Slotted Test-Section Walls to Reduce Wall Interference for High-Lift- propulsion and airframe noise research, as well as Model Testing," NASA TN D-6292, 1971.
aerodynamic research.
18Parker, A.G., "Use of Slotted Walls to Reduce Wind-Tunnel The 40- by 80-ft wind tunnel has proved to be a valuable Boundary Corrections in Subsonic Flow," AIAA Journal, Vol. 12, tool in aerospace research. The planned modifications will Dec. 1974, pp. 1771-1772.
expand and enhance its already unique capabilities and utility.
19Binion, T.W., Jr., "An Experimental Study of Several Wind The improvements will be made by additions to an existing Tunnel Wall Configurations Using Two V/STOL Model Con- facility to minimize costs. This method of expansion of test figurations," Arnold Engineering Development Center, Tenn., capability should prove to be a useful and viable approach for AEDC-TR-75-36, July 1975.
other, smaller facilities as well.
2°Mann, M.J., "Low-Speed Upwash Interference on a Transport Model in a Rectangular Slotted-Wall Wind Tunnel," NASA TM X- References 3218, 1975.
I Kelly, M.W. and Hickey, D.H., "Requirements and Design 21Hansford, R.E., "The Removal of Wind Tunnel Panels to Considerations for a New Full-Scale Subsonic Wind Tunnel," NA TO Prevenl Flow Breakdown at Low Speeds," AGARD Conference Seminar on General Problenzs Relating to Aerodynanlic Testing Proceedings 174 on Wind-Tunnel Design and Testing Techniques, Facilities, Institute Franco Allemand de Recberches de Saint-Louis, Paper 41, AGARD-CP-174, Oct. 6-8, 1975 May 4-7, 1971.
22Hackett, J.E. and Praytor, E.B., "Ground Effect for V/STOL -'Kelly, M.W., Mort, K.W., and Hickey, D.H., "Full-Scale Aircraft Configurations and Its Simulation in the Wind Tunnel, Part Subsonic Wind Tunnel Requirements and Design Studies," NASA I--lnlroduction and Theoretical Studies," NASA CR 114,495, No','.
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3Kelly, M.W., McKinney, M.O., and Luidens, R.W., "The -'3 Hackett, J.E., Boles, R.A., and Praytor, E.B., "Ground Effect Requirements for a New Full-Scale Subsonic Wind Tunnel," NASA for V/STOL Aircraft Configurations and Its Simulation in the Wind TM X-62,106, Feb. 1972; AIAA Paper, San Diego, Calif., Jan. 17-19, Tunnel, Part lI--Experimcntal Studies," NASA CR 114,496, Nov.
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4 Kelly, M.W., "Meeting the Challenge of Advanced Helicopters," "_Hackett, J.E., Praytor, E.B., and Caldwell, E.O., "Ground Vert_flite Magazine, Vol. 19, March/April 1973, pp. 4-8.
Effect for V/STOL Aircraft Configurations and Its Simulation in the sWilliams, J., Proceedings of AHS Panel Discussion on the Wind Tunnel, Part Ill--Application to the NASA Ames 40- by 80-ft Requirements for a New Large Subsonic Wind Tunnel for Research Wind Tunnel," NASA CR 114,497, Nov. 1972.
and Det,elopment Te_ting on V/STOL Aircraft, Oct. 1972; RAE Tech Memo Aero 1472, Jan. 1973.
25Hackett, J.E. and Boles, R.A., "High Lift Testing in Closed 6Mort, K.W., Kelly, M.W., and Hickey, D.H., "The Rationale Wind Tunnels," JournalofAircraft, Vol. 13, Aug. 1976, pp. 597-604.
and Design Features for the 40- by 80-/80- by 120-Foot Wind Tun- '6Hackett, J.E., Boles, R.A. and Lilley, D.E., "Ground nel," AGARD Conference Proceedings 174 on Windtunnel Design Simulation and Tunnel Blockage for a Jet-Flapped, Basic STOL and Testing Techniques, Paper 9, AGA R D-C P-174, Oct. 6-8, 1975.
Model Tested to Very High Lift Coefficients," NASA CR-137,857, 7Heyson, H.H., "Rapid Estimation of Wind-Tunnel Corrections March 1976.
_,_ith Application to Wind-Tunnel and Model Design," NASA TN D- -'7Hickey, D.H. and Kelly, M.W., "Aspects of Large-Scale, 6416, 1971.
Subsonic, Wind-Tunnel Design for Propulsion Noise Research," 8Sears, W.R., "Self-Correcting Wind Tunnels," Aeronautical presented as Paper 73-1279 at the AIAA/SAE Propulsion Con- Journal, Vol. 78, Feb./March 1974, pp. 80-89.
ference, Las Vegas, Nev., Nov. 6-7, 1973.
9Vidal, R.J., Erickson, J.C., Jr., and Catlin, P.A., "Experiments -_SMort, K.W., Eckert, W.T., and Kelly, M.W., "The Steady-State with a Self-Correcting Wind Tunnel," AGARD Conference Flow Quality of an Open Return Wind Tunnel Model," Canadian Proceedings 174 on Windtunnel Design and Testing Techniques, Aeronautics and Space Journal, Vol 18, Nov. 1972, pp. 285-289.
Paper II, AGARD-CP-174, Oct. 6-8, 1975.
l°Goodyer, M.J., "A Low Speed Self Streamlining Wind Tunnel," 29Scharton, T.D., Sawley, R.J., and Wilby, E.G., "An Acoustic A GARD Conference Proceedings 174 on Windtunnel Design and Study for the Modified 40× 80 Foot Wind Tunnel," Bolt, Beranek, Testing Techniques, Paper 13, AGARD-CP-174, Oct. 6-8, 1975. and Newman, Inc., NASA Contract NAS 2-8330, Feb. 10, 1975.