Document
N_ Tedmk_l Memorandum 86920
AIAA-85-0090
The NASA Altitude Wind Tunnel:
Its Role in Advanced Icing
Research and Development
N85-15758
(M&S &-T M- 86 92 0) _fi5 _ASA *LIZTUDE W_ND
TUNM_.L _AWT): I_S BC[E I_ ACVANCED ZCING
_ESE&RC_ &_D DEVELG/_tENY (NAS*) 21 p
Onclas
HC &O2/MF A01 CSCL lqB
13457
G3/09
Bernard J. Blaha and Robert J. Shaw
Lewis Re_,arck Ce_er
Clevelo.d, Ohio
Prepared for the
Twenty-third Aerospace Sciences Meeting
sponsored by the American Institute of Aeronautics and Astronautics
Rmo, Nevada, January 14-17, 1985
|
...... "" .... IV" ..... _ / ,4 L THE NASA ALTITUDE WIND TUNNEL: ITS ROLE IN ADVANCED ICING RESEARCH AND DEVELOPMENT Bernard J. Blaha and Robert J. Shaw National Aeronautics and Space Administration Lewis Research Center Cleveland, Ohio 44135 Abstract technological advances have provided the impetus to seek all-weather capability for all aircraft Currently experimental aircraft icing research classes including some which had no previous is severely hampered by limitations of ground icing requirement. This in turn has resulted in a simulation facilities. Existing icing facilities revised interest in aircraft icing research and do not have the size, speed, altitude, and icing also a renewed interest in icing ground test facil- environment simulation capabilities to allow accu- ities. As will be discussed in this paper the rate studies to be made of icing problems occurring future of icing research is severely limited by the for high speed fixed wing aircraft and rotorcraft.
lack of appropriate ground test facilities. In Use of the currently dormant NASA Lewis Altitude response to this identified need the NASA, Lewis Wind Tunnel (AWT), as a proposed high speed pro- Research Center has proposed the modification of pulsion and adverse weather facility, would allow the now dormant Altitude Wind Tunnel (AWT) into a many such problems to be studied. The character- new, larger, higher performance, wind tunnel test istics of the AWT related to adverse weather simu- facility appropriate for future propulsion system lation and in particular to icing simulation are integration and severe weather R&D, including icing discussed, and potential icing research programs and heavy rain. Since its initial proposal the AWT using the AWT are also included.
has been studied by the DOD, FAA, Industry, and AGARD. As a result of significant outside endorse- Introduction ment including the aforementioned groups, as well as the Aeronautics and Space Engineering Board The National Aeronautics and Space (ASEB) and the Aeronautics and Astro,lautics Coor- Administration, through the work conducted by the din3ting Board (AACB), NA3A is proceeding in its Lewis Research Center in Cleveland, Ohio, is com- efforts to obtain Formal Project Approval, which mitted to the continued advance_nent of technology will then allow start of construction.
for new and improved aircraft propulsion systems.
Several recently completed studies_, _ on a The characteristics of the proposed AWT for national assessment of the future outlook for aero- simulating adverse weather conditions, particularly nautics concluded that major advances in aircraft icing, are discussed in this paper. Potential capabilities are yet to be made. Major conclusions icing research programs using the AWT are also of these studies indicate that to achieve these included.
advances aggressive p_ograms must be pursued in propulsion and in the integration of the propulsion Icing Requirements system with the total aircraft. They also con- cluded that it will be essential that these As stated previously, new technologies are advanced systems have expanded capabilities for evolving which are leading to new and different adverse weather operation, primarily icing and aircraft systems. These new technologies are hav- heavy rain.
ing a profound impact on the future of Icing R&D.
Examples of some of these new developments are To date the country has had excellent test shown in Fig. I which include advanced engines and facilities, however new technologies are evolving engine cycles which will have less bleed tolerance which are ]eading to new and different propulsion to provide the hot gas now used for icing and pro- systems for new and different aircraft with pulsion system ice protection. Also listed here expanded capabilities like severe weather opera- are advanced materials (e.g., composites) which may tion. These new technologies, in turn, will impact the use of future hot air and electrothermal require new and different test facilities to accom- ice protection systems; advanced airfoils which may plish them. Examples of these new technologies are be either more or less sensitive to ice accretion; advanced engine and convertible engine cycles, high highly integrated vehicle configurations, which speed propeller concepts, materials (e.g., compos- also now need protection; advanced avionics and ites), controls, and advanced avionics. These new controls; and new mission requirements. As indi- technologies, in turn, are leading to future air- cated in Fig. 2 all the current Icing Research and craft systems which are much more complex in desiqn Development (R&D) applications will have to be and mission, and wi]l require a higher degree of revisited and new technologies developed for pro- system inteqration compared to present systems.
tecting items such as new propulsion system compo- The technica] challenges associated with the nents including inlets, nacelle structures, achievement of these new systems are substantial carburetors, and new and more exotic (higher bypass and will require a combination of continued ratio) fans and new high speed propellers. New ice advances in both computational capabilities and the protection concepts will also be required for fixed availability of test facilities with unique char- wing and now rotary wing aircraft which will also acteristics which currently are either severely incorporate advanced airfoils, high llft devices, ]imited or do not exist at a]I.
and new fuselage structures.
The need for increased severe weather operation These recent and continuing advances have pro- of new and different aircraft systems is an exa_le vided the impetus to seek all-weather capability of where the above is particularly true. Recent for, as indicated in Fig. 3, a11 aircraft classes including some which had no previous requirement. icing analysis methods; and new and/or improved Today due to technological advances in avionics icin 9 test facilities.
and flight controls, nearly all helicopters and general aviation aircraft can and are equipped to Early in the new icing effort, NASA and others operate under Instrument Flight Rules (IFR) condi- surveyed the,_clng test facilities throughout the free world°'_u capable of meeting the modern tions. Yet only a few military helicopters have icing clearances, and no U.S. civil helicopter has requirements. These investigations revealed a FAA certification for flight into forecasted ictng serious lack of ground icing research facilities conditions. Many of today's general aviation air- for testing in particular helicopters, missiles, craft are certified for operation in an icing general aviation, and higher performance aircraft.
environment, but they rely on ice protection tech- These results were further corroborated in the nology that is over 20 years old. The relatively other studies already mentioned, which helped formulate the NASA program. _'" small payload fraction and low power margins of these smaller aircraft mean that their ice protec- tion systems must be ltght in weight and low in Icing Facility Requirements power consumption. Similarly, because of high fuel costs, today's large commercial transports need As shown in Fig. 5, the phenomena of icing is complex and requires that many variables be simul- lighter and more efficient ice protection systems.
Tomorrow's transport aircraft will need alterna- taneously controlled for correct simulation. Ice tives to the current hot air ice protection systems can form in two ways, either the droplets freeze because bleed air will be scarce on the more effi- immediately upon impact, forming rime ice, or cient higher-bypass-ratio engines or high speed shortly after impact forming glaze ice. Each case results in a different ice formation with different turboprop engines. Ice protection will also be imperative for the advanced highly survivable mil- growth patterns and each case can have serious and itary aircraft and cruise missiles whici_ will fly deleterious effects on the aerodynamic performance long distances at low altitudes where icing is of of the surface upon which it forms. AS shown, the major concern. Thus, all of these types of air- growth of the ice is a function of many variables, craft now share common icing requirements: highly including velocity, temperature, pressure, liquid effective, lightweight, low-power consuming deice water content, droplet size, airfoil size and and anti-icing systems, and detailed knowledge of shape, and also vehicle angle-of-attack. There- the aerodynamic penalties due to ice on aircraft fore, proper icing testing requires not only con- surfaces. trol of more test variables, but also more exacting control than is required for aerodynamic testing.
To fully exploit the potential of the future Likewise, icing phenomena cannot be simply scaled as is the case with aerodynamic testing. As shown energy efficient, high performance, survivable, and operationally flexible aircraft requires that they in Fig. 6, ice accretion varies greatly with the be capable of successfully operating in all weather size of the object. Attempts to properly scale ice accretion have been made over the years, but were conditions. Continuing pressures to expand flight never verified as being successful. In these envelopes, geographic routes and flight frequency indicates that the effects of weather phenomena attempts it was found that adjustments must also such as icing, heavy rain, and snow must continue be _de in velocity, liquid water content, time of to be explored and this can only be done under accretion, and droplet size to begin to achieve any realistic environmental conditions. Successfully degree uf success. As a result, little work has achieving all-weather operational capability in been done with subscale models and icing has been future highly survivable military aircraft and in considered a "full-scale" R&D area. Ice scaling current and future rotorcraft requires a more research is continuing, but no validation of any thorough understanding of the potential adverse scaling law has yet been achieved.
effects of weather on these concepts. For example, the understanding and development of icing tech- As a result, icing R&D has been conducted at nology requires evaluation under realistic atmos- full scale in either flight or in the few ground pheric conditions. Today these conditions can only test facilities capable of testing full scale hard- be found via costly, time consuming and risky ware. The largest current icing facility in the flight testing. U.S. is the 6 by g Ft Icing Research Tunnel (IRT) at the NASA Lewis Research Center. A typical test In response to the nationally recognized need conducted in this facility, shown in Fig. 7, for new icing and severe weather R&D, NASA has includes full scale equipment which generally results in severe blockage effects. This tunnel organized a new aircraft icing research program at Lewis to help solve the icing problem for modern is heavily used, but has been identified by the aircraft. As schematically indicated in Fig. 4 Icing community as being extremely limited, because this program is broad based, and covers both basic of these blockage effects and also limitations in speed and altitude capabilities. The test capa- research and engineering applications. Specific elements of the program are summarized in Table I bilities of the IRT are shown in Fig. 8 and com- and an expanded description of the overall program pared to the flight conditions described in the is included in Ref. 3. This program is well coor- Federal Aviation Regulations (FAR) Part 25 require- dinated among the various NASA Centers, the FAA, ment for certification of flight into forecast DOD, universities, industry, and some foreign gov- icing conditions. Although temperature simulation of the IRT is good (-20 to +32 °F), the IRT does ernme,ts. In fact. the program was formulated with the help of all these groups thrgugh a series of not simulate altitude pressure and the top speed workshops and special studies, a'/ Inputs from is only Mach 0.4 when the tunnel is empty. Large these studies identified four main needs for future blockage models limit the top speed to something less, Consequently, the IRT does not adequately icing R&D: new and more efficient ice protection concepts; improved icing instrumentation; advanced simulate flight, in particular for the current and perature variations from 60 °F down to less than future higher performance aircraft configurations. -40 °F. Standard wind tunnel components of the facility are shown in Fig. 13. The original AWT had a 20 ft diameter circular solid wall test sec- Consequently, as summarized in Fig. 9, no suitable ground facility exists for proper simula- tion with a top speed near Mach 0.6. This will be tion of flight altitude conditions, at correct replaced with a new slotted octagonal configuration flight speeds, with large or full scale equipment.
more appropriate for high subsonic speeds. Wall As previously identified, this lack of appropriate boundary layer air will be bled through the slots ground test facilitie', has, and will continue to into a plenum surrounding the test section and then severely limit or impede the technology development re-injected downstream of the test section, to necessary for future higher performance aircraft. minimize model blockage effects. At the downstream Without the availability of appropriate test facil- end of the test section variable finger flaps will ities, flight and test development times will have also be used. Two 30 000 hp electric motors will to increase, particularly for the new and more com- be provided to power a two-stage variable speed plex flight systems. What is needed, therefore, is fan with variable inlet and interstage guide vanes.
a new test facility with the simultaneous capabil- The number of guide vanes (28 per stage) was ities as summarized in Fig. lO. These requirements selected to minimize the interactions between the vanes and rotor blades and thus minimize the fan include the concurrent pressure and temperature simulation of altitude, the ability to test large noise. A heat exchanger will provide cooling to or full scale hardware, and with speeds to near remove the heat added to the tunnel air by the Mach 1.0. To properly simulate flight test condi- drive fan and test models, and to simulate the tions over an extended length of test area the desired altitude static temperatures. Flow condi- facility should also be a wind tunnel. Sizeable tioners will consist of honeycomb sections and removable screens. The screens would be removed capacity flow services are required for propulsion system operation and/or simulation with large during adverse weather testing. When in place, the engine simulators. This requires large capacity flow conditioners will provide good quality (low turbulence) flow to the test section.
air services for engine exhaust flow scavenging and make up air supplies to replace that removed by the exhaust system. Lastly, this facility must have Special features which make the AWT a unique the capability of water spray Systems appropriate test facility are shown in isometric view in for icing and heavy rain simulation. Fig. 14. To prevent contamination by exhaust h- products, an engine exhaust scoop will be located In response to the identified need NASA Lewis downstream of the test section. This scoop will I.: has proposed the modification of an existing dor- provide both altitude exhaust and engine exhaust mant wind tunnel facility that will simultaneously flow scavenging via the laboratory Central Exhaust satisfy all the test requirements needed to suc- System. This scoop will be variable in the pitch cessfully evaluate and resolve the technological plane to allow for engine angle-of-attack variation W_ challenges of future high potential aeronautical and will have variable and interchangeable tip vehicles.
geometries. It will also be cooled to allow test- ing even with afterburning engines. Make-up air will be injected just downstream of the fan. A Proposed Altitude Wind Tunnel (AWT 1 Plenum Evacuation System (PES) will be used to Fhe AWT, seen in Fig. l], was built in 1944 and pump the test section plenum to allow for testing served NACA as a low speed (M_ 0.6) propulsion with model blockages near 12 percent at high sub- wind tunnel until 1958. In 1958 NACA became NASA sonic speeds. The pumped plenum air will also be and the technology focus shifted from aeronautics re-injected into the tunnel just aft of the fan.
to space. Because the AWT had low pressure (high Heated turning vanes will be used in the two cor- altitude) and low temperature capability and also ners upstream of the fan to prevent and remove ice because the aircraft research of the time was in buildups during adverse weather testing. The heat the high speed flight regime, it was decided to exchanger will be con- netted to a 21 000 ton convert this wind tunnel into a much needed vacuum capacity, Freon-22, two-phase refrigeration system.
test facility. In ]960 the AWT then became the This system will permit operation at static temper- Space Power Facility and served NASA in this capac- atures that are encountered at altitudes varying ity until 1970. The tunnel was converted into two from sea level to in excess of 55 000 ft. Lastly, large vacuum tanks and was used for many space a removable water spray bar system will be inserted into the tunnel upstream of the bellmouth for applications, including extensive Launch Vehicle testing. Since 1970 the facility has been unused adverse weather testing (i.e., icing, heavy rain, and would thereby provide a cost effective founda- ard possibly snow). This system will have the tion upon which to build a new, needed wind tunnel capability to introduce water in various droplet sizes and water content levels. Details of how the facility.
spray bar system will be inserted and removed from A schematic of the proposed new AWT is shown the tunnel are also shown in the figure. The spray in Fig. 12 with its projected test capabilities.
system will thereby be interchangeable with the tunnel flow conditioners so that flow quality may The tunnel will have a nominal 20 ft, octagonal, slotted high subsonic speed test section. Mach be preserved for aerodynamic and propulsion tests.
numbers ranging from near 0 to more than 0.9 will As seen in Fig. 14, this large tunnel will be be achievable with large blockage models (lO to 12 extremely flexible and provide some unique func- tions and capabilities which are not found in any percent) including complete operating propulsion systems. The Lewis central altitude exhaust system other existing wind tunnel.
will provide pressure altitude varlation from sea level to greater than 55 000 ft. The tunnel Details of the high speed test sac&ion are refrigeration system will allow tunnel total tam- elaborated in Fig. 15. The octagonal cross section geometry was selected because it provides a conven- Further exa_les of typical test programs, ient configuration for incorporating the necessary which demonstrate the diverse flexibillty of the special features. This configuration allows for AWT, are shown in Figs. 20 to 24. These test pro- the convenient use of bleed slots arld reentry flaps grams could include general aviation, rotorcraft, which are necessary for testing of large blockage highly survivable military, and advanced missile models at high subsonic speeds. Furthermore, the systems. The tunnel would also be appropriate for flat walls allow the extensive use of high quality advanced CTOL propulsion system testing. In each optical windows necessary for icing testing and for case, multiple tests could be performed such that the advanced laser data systems. This design also either performance, acoustics, or severe weather allows for the easy replacenw_nt of the tunnel side- data could be obtained. This feature could reduce walls with acoustically treated panels more appro- the amount of testing in any particular program.
priate for acoustic testing of advanced propulsion AS shown in Fig. 24, a unique type of testing could systems. The test section will include both floor be performed in the AWT. Because the AWT would and side wall model mounting. The floor mount will have cold walls, IR signatures of hot exhaust could include a multlco_)onent force balance for system be n_asured which could not be obtained in an performance testing. Models will be installed and unrefrigerated wind tunnel. An expanded discussion removed from the test section through a bottom of the capabilities of AWT for propulsion system entry which will lower the tunnel floor to the shop and system integration testing is included in level below on screwjacks. The two lower tunnel Ref. 11.
quarter panels will rotate out of the way so that full span models can be easily installed. These The proposed AWT Project schedule is _hown in features will provide for rapid model access Fig. 25. A Preliminary Engineering Study (PER) was thereby increasing the productivity of the tunnel.
completed in December, 1984. As was mentioned earlier, NASA is seeking approval for a construc- The nominal pressure altitude/Mach number tion (CoF) start. If approval is soon obtained, capability of the tunnel is shown in Fig. 16.
final design will start in FY 1985, construction Because many of the potential programs to be run will begin in FY 1986, and the facility will be in the AWT may be sensitive and/or proprietary, ready for checkout and calibration in FY 1990. At special features are being incorporated into the present, to support the final design and to insure tunnel design. These features, as shown in the success of the tunnel in meeting its test Fig. 17, include separate model buildup and check- requirements an extensive modeling program has been out areas, restricted access to the flow circuit, initiated at Lewis. A detailed description of this secure data systems and control rooms, and special program is included in Ref. 12. In this program, office locations.
every tunnel component will be tested at l/IO scale, first alone and then integrated into a com- As was _w_ntloned previously, one of the plete full circuit loop. Additional model testing requirements for a new propulsion wind tunnel is is being conducted at Lewis using the other wind special support services for propulsion testing tunnel facilities, particularly the IRT. As including high capacity air systems. A listing of described in Ref. 13 extensive use is also being a series of these support services already existing made of advanced analytical techniques and analy- and operating at Lewis is shown in Fig. 18. These sis. Analytical modeling of the tunnel controls is system capabilities are unique in NASA and include also underway. These efforts will include the compressed air supplies capable of continuous development and use of a real time system simulator operation at high pressures and weight flows, and whereby the dynamics of the tunnel controls and provide a capital net worth exceeding $200 million.
systems can be studied before the tunnel is first started. As indicated in Fig. 25 these modeling The AWT will, therefore, provide the U.S.
efforts are projected to be very extensive during Aeronautics industry with a needed, truly unique, final design, but will drop off as construction and diverse wind tunnel for future propulsion sys- continues. Eventually, a complete scale model of tem integration and icing R&D. The unique capabil- the AWT will come to be which can either be used as ities _f the AWT to perform icing research are a pilot tunnel for AWT or as a separate research summari,ed in Fig. 19. The AWT Mach number and facility.
altitude capabilities are shown compared to the FAA icing limit and also to the IRT. The AWT will pro- Concludinq Remarks vide the necessary altitude simulation and speeds to cover the complete range of interest and will It is projected that many advances can and will have a test section area approximately six times be made in future aeronautical systems. However, larger than the IRT. Lastly, because of its these advances will only come from a more thorough special capabilities, complete propulsion systems understanding and resolution of integrated systems may be tested over the con_lete required range of technology and solutions to adverse weather related test conditions. As shown in the figure, in con- problems. This will require new unique analytical trast to the tsolated nacelle tested in the IRT, and experimental capabilities appropriate to suc- the AWT wtll be able to test complete full scale cessfully evaluate these systems integration prob- advanced high speed propeller propulsion systems.
lems under realistic operating conditions. Not all Performance testing would be done at the same time of the required facilities necessary to achieve as icing testing. In summary, this figure Shows these goals exist today.
that the AWT wlll be a significant improvement over the capabilities o _ the IRT which is the largest As indicated in Fig. 26 the proposed modifi- and most heavily used icing facility in the U.S. cation to the now dormant Lewis Altitude Wind Plans also include capabilities to perform heavy Tunnel provides a cost effective means of achieving rain testing, and if the technology is ever devel- a critically needed unique capability that does not oped for snow slmulations, the AWT will also be currently exist within the U.S. This factllty has appropriate.
been substantially endorsed as _ettng the needs of 6. Breeze, R. K.; and Clark, G. M.: Light Tran- the icing research community. The extensive modeling effort that is supporting the facility sport and General Aviation Aircraft Icing Research Requirements. (NA-81-110, Rockwell design is the most complete and comprehensive that NASA has ever undertaken. These efforts will International Corp.; NASA Contract NAS3-22186.)
NASA CR-1652go, Ig81, thereby insure the success of this program and will play a major role in maintaining U.S. superiority 7g in Aeronautics well into the 21st Century. Koegeboehn, L. P.: Commercial Aviation Icing Research Requirements. NASA CR-165336, 1981.
References 8.
01sen, William: Survey of Aircraft Icing Sim- ulation Test Facilities in North America. NASA I.
Executive Office of the President: Report of the Aeronautical Policy Review Committee. TM-81707, 1981.
Office of Science and Technology Policy (OSTP).
9.
November 9, 1983. Taylor, F. R.; and Adams, R. J.: National Icing Facilities Requirements Investigation.
2. FAA-CT-81-35, 1981.
Aeronautics and Space Engineering Board (ASEB), Commission on Engineering and Technical Sys- I0.
tems, National Research Council (NRC): Aero- Rotorcraft Icing - Status and Prospects.
nautics Technology Possibilities for Year 2000: AGARD AR-166, 1981.
Report of a Workshop, 1984.
II.
Chamberlin, R.; and Miller, B. A.: Altitude 3. Wind Tunnel (AWT) - A Unique Facility for Pro- Reinmann, J. J.; Shaw, R. J.; and 01sen, W. A., pulsion System and Adverse Weather Testing, Jr.: NASA Lewis Research Center's Program on Icing Research. NASA TM-83031. AIAA Paper AIAA Paper 85-0314, 1985.
83-0204, 1983.
12. Abbott, J. M.; Diedrich, J. H.; Groeneweg, J.
4. Aircraft Icing. FAA-RD-78-109, NASA CP 2086, F.; Povinelli, L. A.; Reid, L.; Reinmann, J.
1978, J,; and Szuch, J. R.: Analytical and Physical Modeling Program for the NASA Lewis Research 5. Center's Altitude Wind Tunnel (AWT). AIAA Peters,n, A. A.; Dadone, L.; and Bevan, A.: Rotorcraft Aviation Icing Research Review and Paper 85-0379, 1985.
Recommendations. (D210-11662-I, Boeing Vertol Co.; NASA Contract NAS3-22384.) NASA 13. Towne, C. E.; Povinelli, L. A.; Kunik, W. G.; CR-165344, 1981. Muramoto, K. K.; Hughes, C. E.; and Levy, R.: Analytical Modeling of Circuit Aerodynamics in the new NASA Lewis Altitude Wind Tunnel. AIAA Paper 85-0380, 1985.
• 4 _ll_r _ J _4 TABLE I. - ELEMENTS OF NASA'S AIRCRAFI ICING PROGRAM Ice protection systems Pneumatic deicers for helicopters Electrothermal deicers Glycol fluid systems Electromagnetic impulse deicers Icephobics Icing instrumentation Cloud instrument evaluation Ice detectors Experimental methods Icing research tunnel Airfoil performance in icing Testing with artificial ice F, elicopter test rigs Icing scaling laws Analytical methods Computer codes for • Water droplet trajectories • Ice accretion modeling • Aero performance penalties • Transient deicer analysis Flight research Validation data for icing simulation facilities Instrument evaluation Icing cloud data Meteorology Helicopter performance
NEW DEVELOPMENTS IMPACTING FUTURE
ICING R&D
• ZERO BLEED ENGINES
• COMPOSITE MATERIALS
• ADVANCED AIRFOILS
• HIGHLY INTEGRATED VEHICLE CONFIGURATIONS
• ADVANCED AVIONICS/ELECTRONICS
• NEW MISSION REQUIREMENTS
FIGURE 1
ICING R&D APPLICATIONS
• PROPULSION SYSTEMS
- INLETS
- NACELLE STRUCTURES
- CARBURETORS
-FAN AND PROPELLERS
• FIXED AND ROTARY WING AIRCRAFT
- ADVANCED AIRFOILS
-HIGH LIFT DEVICES
-FUSELAGE
FIGURE 2 N FIGURE 3 FIGURE 4 FIGURE 5
mR_'l" or u oa, W m a _
©
FIGURE 6
ALL WEATHER APPLICATIONS
ICINGAND HEAVY RAIN ICING TESTOFA FULLSCALE GENERAL AVIATION INLET/NACELLE IN THELEWISICING RESEARCH _ pRT) C-M-1345 Figure 7.
EXTENT OF ICING R&D TEST REQUIREMENTS
FIGURE 8
ICINGR&D TEST REOUIREMENTS SUMMARY
• FLIGHT ALTITUDE SIMULATION
• CORRECTFLIGHT SPEEDS _"
• FULL/LARGE SCALE_
J SUITABLE FACILITY
NON'EXISTENT
FIGURE 9
NEW TEST FACILITY REQUIREMENTS
• CONCURRENT PRESSUREAND TEMPERATURE
SIMULATION OF ALTITUDE
• LARGE SCALE TEST ARTICLES
• FULL SUBSONICSPEEDRANGE
• WIND TUNNEL CONFIGURATION-
AERODYNAM ICS/ACOUSTICS
• PROPULSION SYSTEM OPERATION/SIMULATION
• ICING,HEAVY RAIN CAPABILITY
I NO EXISTING NOR PLANNEDFACILITY MEETS THESE NEEDS I
I I FIGURE 10 "" "='
ALTITUDE WIND TUNNEL
C-84-1412
Figure Ii.
PROPOSED ALTITUDE WIND TUNNEL
FIGURE 12
OF pOC_R _,-."--''
D OF P_,.-,-, Q, ..... :,
STANDARD WIND TUNNEL COMPONENTS
TWO 30K HP ORfYE MOTORS 20 ft diam FLOW TEST SECTION ..... CONDITIONERS D FIGURE 13 E mr F --...
D
ALTITUDEWIND TUNNELSPECIALFEATURES
_._ _._ _/STEAM HEATEO TURNN6 VANES _e"_ . ,, _ ENGINE EXHAUST MAKE-UP AIR _ " t::--.. _ _ ,"" _. _" '_-_, EXCHANGER
f/ "++++ ...+,,,_" +#,'.' _ f/I+l+l
+..j . , +*, _ ,
/ ..... "W?--++5_,> - i_
SYSTEM / . " _Pr _,',_ . -_ HEMOVEAeLE FLOW FIGURE 14 ....
TEST SECTIONFEATURES
• OCTAGONALCROSSSECTION
• SLOTTEDWALL WITH PLENUM
• FLAT WINDOWS
• FORCEBALANCE
• REMOVABLE ACOUSTICALWALLS
• BOTTOM MODEL ENTRY
• BARD MODEL ACCESS
'T
FIGURE 15 F- ze,,- FIGURE 16
FEATURES FORSENSITIVEPROGRAMS
• ABILITY TO ROUTINELY HANDLESENSITIVE MODELS
-THREE SEPARATEMODELBUILDUP AND CHECKOUT ROOMS
-MODEL STORAGE PROVISIONS
• RESTRICTED ACCESSTO FLOW CIRCUIT
• DATA SYSTEM AND DATA SECURITY
-STAND ALONEDATA ACOUISITION/ANALYSIS/DISPLAY CAPABILITY
.SHIELDED
•RESTRICTED PHYSICALACCESS
-SECUREDOFF LINE DATA REDUCTION AND STORAGE CAPABILITY
• OFFICE SUPPORT
FIGURE 1 7
ALTITUDEWIND TUNNEL SPECIAL SUPPORTSERVICES
*CENTRAL COf]TINUOUS COMPRESSED AIR SUPPLIES
- 450 Ib/sec AT 45 psi
- 400 Ib/sec AT 150 psi
- 76 Ib/sec AT 450 psi
oCENTRAL EXHAUST SYSTEM
-480 Ib/sec AT SEA LEVEL PRESSURE AND 700 F
FIGURE 18 FIGURE 19
GENERALAVIATION VEHICLES
/--\
\__/
• PERFORMANCE
• ICING
• COOLING
• AEROELASTICS
• ACOUSTICS
FIGURE 20
ADVANCEDROTORCRAFT PROPULSION
IN AWT
/
• HIGH SPEED PERFORMANCE
"ACOUSTICS
• ICING
• CONTROLS
FIGURE 21
STOVL AIRCRAFT WITH FIO0
ENGINE IN AWT
I I
/
t I
I _,\1
I I
\w/
| I
• PERFORMANCE
• ICING
FIGURE 22
®
HIGH BYPASS RATIO TURBOFAN
PROPULSIONSYSTEMS
I
• PERFORMANCE
• ACOUSTICS
• AEROELASTICS
• ICING
FIGURE 23
SPECIAL PURPOSE VEHICLES
_l _ l' /--\
• SIGNATURES
• ICING
• PERFORMANCE
FIGURE 24
OF POOR QJ,_,L_ ,,'_"
FIGURE 25
ALTITUDE WIND TUNNEL
PROVIDES NEW NATIONAL AERONAUTICAL PROPULSION AND ICING R&D CAPABILITY C-84-I347 Figure Z6, I