NACA Conference on Aircraft Loads 1957

8/8/2019 NACA Conference on Aircraft Loads 1957

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LIST OF CONFEREES

The following were registered a t t he NA Conference on AircraftLoads, St ruc tur es, and Fl ut te r, Langley Aeronautical Laboratory, LangleyFie ld , Va . , March 5 , 6, and 7, 1957:

ABBOT", I r a H.ABRAHAM,LewisADAMS, Fred T.ADAMS, Laurence J.ALBERI, AmericoALEXANDER,Marvin M .AZLEN, RobertANBRO, Paul A.

. ANDERSON, Enoch A.ANDEXSON, Franklin C.

ANDERSON, John W.ANDERSON, Melvin S.ANTILL, Richard B.ASHLEY, Holt

NACA HeadquartersSubcommittee on Aircraft StructuresNaval A i r Development StationGlenn L. Martin CompanyRepublic Aviation CorporationConvairC i v i l Aeronautics AdministrationGrumman Aircraft Engineering Corp.Boeing Airplane CompanyBell Aircraft Corporat ion

ConvairNACA - Langley LaboratoryRaytheon Manufacturing CompanySubcorrnnittee on Vibration and Flutter

BADGER, David M. Northrop AircraftBAILEX, Robert A. Lockheed Aircraft CorporationBAILEY, Freder ick J . , Jr. NACA - Langley LaboratoryEAIRD, Eugene F.BAKER, Wilfred E.BANAS, ConradBANNER, Richard D.BARCUS, Ronald

BARD, Donald 0.EARKER, Ly n n M .BARTER, John W.BATTERSON, Sidney A.MUM, C . PhilemonBECKER, John V.BEEBE, JohnBEELER, De ElroyBEEXOVITS, Avraha,mBERTRAM, Mitchel H.BEZEATCHEM(0, John W .BIRNBAUM, SidneyBITZ, Eugene A.BLAND, W i l l i a m M . , Jr.BOBBITT, Percy J .BOCK, Charles D.BOGEMA, Bernard L.

Subcommittee on Vibration and FlutterB a l l i s t i c Research LaboratoryUnited Aircraft CorporationNACA High-speed Flig3t StationBendix Aviation Research Labs.

Chrys e r CorporationSandia CorporationDouglas Aircraft CompanyNACA - Langley LaboratoryBureau of AeronauticsNACA - Langley LaboratorySubcommittee on Low-Speed AerodynamicsNACA High-speed Flight StationNACA - Langley LaboratoryNACA - Langley LaboratoryGoodyear A i r c r a f t CorporationG l e n n L. Martin CmpanyNaval Ordnance LaboratoryNACA - Langley LaboratoryNACA - Langley LaboratoryArma Div. , American Bosch Arma Corp.Glenn L. Martin Company

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BOND, Aleck C .BOONE, Paul W .

BOSWINKLE, Robert W . , J r ,BOUTON, InnesBOWMAN, John C .BRASSAW, Lloyd L. , J r .BREMER, Paul H.BRESSETTE, Walter E .BREUHAUSER, W . 0 .BREWER, Jack D.BROOKS, Thurman P.BROOKS, W i l l i a mA., J r .BROWN, Clinton E .BROWN, Harvey H.BUCKLEY, Edmond C.

CAHEN, George L.CARLSON, Harry W .CARLSON, Wendell C .CASS, Lorne E .CATHAWAY, Russ G.CAYWOOD, W i l l i a m C .CENTER, Kenneth W.CHILTON, James P .CLARK, L t . John B.CLARKE, Martyn V.CLEVENSON, Sherman A.COL;EsIAN, Thomas L.

COOPER, Ralph D.

COWGILL, Lee C .COX, Joe F.CRABILL, Norman L.CRAWFORD, Davis H.CREEL, RalphCROWLEY, John M .CROWLIE, John W .CULBERTSON, P h i l i p E.CUNNINGHAM, Herbert J .CYPHERS, Richard J . , J r .CZARNECKI, K. R .

COONEY, Thomas v.

DAUM, Fred L.

DAVIS, Don D., J r .DAWSON John R .DEARING, DavidDEEP, Raymond A.

NACA - Langley LaboratorySubcommittee on Aircraft Structural

NACA - Langley LaboratoryNorthrop AircraftRyan Aeronautical CompanyBel l Aircraf t Corpora t ionLockheed Aircraft CorporationNACA - Langley LaboratoryCornell Aeronautical LaboratoryNACA HeadquartersMcDonnell Aircraft CorporationNACA - Langley LaboratoryNACA - Langley LaboratoryNACA HeadquartersNACA - Langley Laboratory

Materials

G l e n n L. Martin CompanyNACA - Langley LaboratoryRaytheon Manufacturing CompanyLockheed Ai rc ra ft CorporationLockheed A ir cr af t CorporationApplied Physics LaboratoryGrumman Aircraft Engineering Corp.Douglas Aircraft CompanyA i r Research an d Development CommandC i v i l Aeronautics BoardNACA - Langley LaboratoryNACA - Langley LaboratoryNACA High-speed Flight Stat ionDavid Taylor Model Basin

Lockheed Aircraft CorporationCessna Aircraft CompanyNACA - Langley LaboratoryNACA - Langley LaboratoryBureau of AeronauticsOffice of Naval ResearchNACA HeadquartersConvairNACA - Langley LaboratoryGrumman A i r c r a f t E ngin eerin g Corp.NACA - Langley LaboratorySubcommittee on Vibration and Flutter

NACA - Langley LaboratoryNACA - Langley LaboratoryBeech Aircraft CorporationRedstone Arsenal

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DeHART, Robert C .DeLANCEY, Lawrence M.DEPTULA, Capt. Alfred R.

DEUTSCKMAN, Jerome N .D I E M , Capt. Walter S. USNDOETSCH, E. K.DOLEN, W i l l i a mK .DONELY, P h i l i pDOTY, Ralph JohnDOUGHEBTY, James E . , J r .DOW, NorrisDRALFY, Eugene C.DRYDEN, Dr. Hugh L.DUBERG, John E .DUKES, W ilfred H.DUNCAN, L t . C m d r. B. B.DUNN, Maurice B.

DUSSAULT, JohnDYE, L t . F. E ., J r .

EXREMY C l i f f o r d 0 .ENTZ, P h i l l i p H.EPSTEIN, Alber tERICKSON, A l b e r t L.ERTHAL, John F.EVANS, A lber t J .EVEIETT, M a j . P h i l l i p E .

FALABELLA, Gaetano, J r .FANTI, Roy

FAVOR, Ronald J .FEDZIUK, Henry A.FELLER, W i l l i a m V.FERNANDEZ, JoseFETTIS, Henry E.FINK, MartinFISCKLZB, Jerome E .FITZGERAIJ), F redF LY W, R i c h a r d W.FORMHALS, Edwin J.FOSTE8, Charles R .FRALICH, Robert W .FRANCEX, Capt. R. G.

FRICK, Charles W .FROST, Richard C.FURLONG, G. Chester

( R e t . )

Armed Forces Sp ec ia l Weapons Pr oj ec tNaval Ordnance T e s t Sta t i onOff ice Deputy Ch. of S ta f f,

Bel l A i rc ra f t Corpora tionCommittee on AerodynamicsWright Air Development CenterSandia CorporationNACA - Langley LaboratoryBoeing Airplane CompanySubcommittee on Aircraft StructuresGeneral Electric CompanyNACA - Langley LaboratoryNACA HeadquartersAeronutronic SystemsBel l Ai rcraf t Corpora t ionBureau of AeronauticsBoeing Airplane Company

Cessna Aircraft CompanyBureau of Aeronautics

Development, USAF

ConvairBoeing Airplane CompanySubconunittee on Aircraft LoadsNACA - Ames LaboratoryNaval A i r Material CenterNACA HeadquartersAF Development F ie ld Repr esenta tive

Air Force Cambridge Research CenterUnited Aircraft Corporat ion

Batte l le Memorial Inst i tuteNACA - Langley LaboratoryNACA - Langley LaboratoryConvairW r i g h t A i r Development CenterUnited Aircraft CorporationDouglas Aircraft CompanyRadioplane CompanySperry-Utah Engineering Lab.Bureau of AeronauticsSubcommittee on Aircraft LoadsNACA - Langley LaboratoryA i r Force Special Weapons Center

ConvairConvairArnold Engineering and Development

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GARCIA, Manuel A.GARRICK, I. E.GARVIN, John B.GATES, Ordway B. , J r .G A W , W i l l i a mGENIESSE, Major Eugene W ., J r .GEUDTNER, Walter J . , J r .GILLIG, Franklin J .GILRUTH, Robert R .GIONFRIDDO, Maurice P.GLASSMEYER, C l i f f o r d E . , J r .GLOVER, Louis S.GOLDIN, RobertGOMZA, AlexanderGORANSON, R . FabianGRANT, Frederick C.GREEB, Major Edwin H.GRIFFIN, Edward J .

GRIFFITH, George E .GRIFFITH, Max 0.GRINSTED, Frank

HALL, Bertrand M.HALL, John B., J r .HALSEY, Robert M.HAMILTON, W i l l i a m T.

HAMMERBERG, Fr i t ch io fHAMMIL, John P.HARDRATH, Herbert F.HARRINGTON, Joseph H.

HARRIS, Thomas A.HARTLEX, Richard M .HAVILAND, John K.HAYDEN, Harold J .HEDGEPETH, John M.HEIMERL, George J .HEITHECKER, Heinrich A.HELDENFELS , Richard R .HELLEBRAND; E m i l A. H.HENDERSON, Arthur, J r .HENDERSON, CampbellHEPPER, Richard H.HESS, Robert W .HOFF, D r . Nicholas J .HOFFMAN, Niles R.HOGE, H. J .HOLTZ, John R .

Naval A i r Miss i le T e s t CenterNACA - Langley LaboratoryNACA - Langley LaboratoryG l e n n L. Martin CompanyRadioplane CompanySubcommittee on High-speed AerodynamicsConvairCornell Aeronautical LaboratoryNACA - Langley LaboratoryAir Force Cambridge Research CenterBat te l le Memoria l Ins t i tu teApplied Physics LaboratorySubcoinnittee on Aircraft LoadsGrumman Aircraft Engineering CorporationNACA XeadquartersNACA - Langley LaboratoryArnold Engineering Development CenterBureau of Aeronautics

NACA - Langley LaboratoryBureau of AeronauticsB ri t i sh Ministry of Supply

Douglas A i r c ra f t CompanyNACA - Langley LaboratoryB e l l Air cra f t Corpora t ionSubcommittee on Aerodynamic S t a b i l i t y

Civil Aeronautics AdministrationDouglas A i r c ra f t CompanyNACA - Langley LaboratorySubcommittee on Aircraft Loads

NACA - Langley LaboratoryDavid Taylor Model BasinChance Vought AircraftBoeing Airplane CompanyNACA - Langley LaboratoryNACA - Langley LaboratoryHolloman A i r Development CenterNACA - Langley LaboratorySubcommittee on Aircraft StructuresNACA - Langley LaboratoryNorth American AviationMcDonnell Aircraft CorporationNACA - Langley LaboratorySubcommittee on Aircraft StructuresWright A i r Development CenterSubcommittee on Aircraft LoadsDavid Taylor Model Basin

and Control

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HOOD, Manley J .HOOVER, Isaac H.HUBBARD, Harvey H.HUNT, Melvin W .HUNTLEY, H. W. , J r .HUSIC, W i l l i a m J .HUSTON, Wilber B.

ISWLSE3?, L t . Col. Orson A.

JACKSON, David W .JENKINS, David R .JOHNSON, Aldie E., Jr .JOHNSON, H. ClayJOHNSON, Harry W.JOHNSON, J . AldridgeJOHNSON, J . B.

JOHNSTON, W i l l i a mM.JORDAN, Gareth H.JORDAN, Peter F.

KAECHELE, Lloyd E .KAMEACK, Elmer L .KAPLAN, AbnerE;AsTEN, Herman G.KEARNS, John P.KEENER, E a r l R .KEIRSEX, Robert L.KEITH, Bobbie C.KERR, T . H.KINNAMAN, Edward B.KINTAS, JohnKLEXSS, h d r . N. J .KORDES, Eldon E .KOTANCHIK, Joseph N.KRAMER, Edward H.KRICKENBERGER,

K R O U , Wilhelmina D.KRUSZEWSKI, Edwin T.KUGEL, H. E .KUHN, PaulKULLAS, Alber t J .KUNZE, Frank C.

L t . Cmdr. C u t e r F., J r .

LaFRANCE, Jeremie U., J r .L A I D L AW, . W i l l i a mR.

NACA - Ames LaboratoryCi vi l Aeronautics AdministrationNACA - Langley LaboratoryNorth American AviationSubconrmittee on Aircraft LoadsCi vi l Aeronautics AdministrationNACA - Langley Laboratoryb e d Forces Sp ec ia l Weapons Pr oj ec t

Wright A i r Development CenterBate l le Memoria l Ins t i tu teAvco Manufacturing CorporationSubcommittee on Vibration and FlutterRaao-Wooldridge CorporationLockheed Ai rc ra ft CorporationSubcommittee on Aircraft Structural

Lockheed A ir cr af t CorporationNACA High-speed Flight StationGlenn L. Martin Company

Mater ia ls

Rand CorporationG l e n n L. Martin CompanyRamo-Wooldridge CorporationRam0-Wooldridge CorporationApplied Physics LaboratoryNACA High-speed Flight Stat ionDouglas A i r c r a f t CompanyWright A i r Development CenterBr i t i sh Minis t ry of Supply

Boeing Airplane CompanyBeech Aircraft CorporationNaval A i r Material CenterNACA - Langley LaboratoryNACA - Langley LaboratoryFa i r ch i ld A i r c ra f t D ivi sionArmed Forces S pec ia l Weapons Pr o j ec t

National Bureau of StandardsNACA - Langley LaboratoryWright Air Development CenterNACA - Langley LaboratoryGlenn L. Martin CompanyChrysler Corporation

Glenn L. Martin CompanySubcommittee on Vibration and Flutter

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LAMPROS, Alexander F.LANDEBS, C . B.LANDES, Paul E .LANDON, John M .LAUTEN, W i l l i a mT., J r .

LAUVER, Dean C .LECAT, RobertLEISS, AbrahamLEONARD, Robert W .LEVY, SamuelLI , Ta Chung-HengLIEBOWITZ, HaroldLOCKE, Frederick W . S. , J r .LOFTIN, Laurence K., J r .L O W , Ted L .LoPRESTI, AntonioLU, HoshenLUCAS, John W .

MAGRATH, H. A.MALVESTUTO, Frank S.MARCOPULOS, GeorgeMARTIN, James E .MARTIN, John W ., J r .MARTZ, Ronald B .MATHAUSER, Eldon E .MAY, Ralph W ., J r .MAYE3, Frank S ., J r .MAZELSKY, BernardMcCASKILL, A l l a n M.McCULLAND, Clyde M.McFARLAND, Donald R .McGINNESS, Capt. W . T.McGOWAN W. A.McHlJGH, James G.

McLEMORE, H. ClydeMcWITHEX, Robert R.MELLQUIST, V ictor G.

MEREDITH, L t . Oliver D.MERZ, Ernes t J .ME=, StephenMICHEL, DouglasMI=, Chester W .MILLER, L t . C m d r. Jack N.MILWITZKY, BenjaminMIROWITZ, L . I.MISKAM, Frederick C.

Naval Air Miss i l e T e s t CenterLockheed A ir cr af t CorporationMcDonnell Aircraft CorporationChrysler CorporationNACA - Langley LaboratoryBureau of AeronauticsFairchild Guided Missile DivisionNACA - Langley LaboratoryNACA - Langley LaboratoryGeneral Electric CompanyConvairOffice of Naval ResearchBureau of AeronaQticsNACA - Langley LaboratoryBoeing A i r p l a n e CompanySubcommittee on Aircraft StructuresRepublic Aviation CorporationJ e t Propulsion Laboratory, C . I . T.

Wright A i r Development CenterNACA High-speed Flight Stat ionNorth American AviationChance Vought AircraftGlenn L. Martin CompanyWright A i r Development CenterNACA - Langley LaboratoryNACA HeadquartersDouglas A i r c r a f t CompanyLockheed Ai rc ra ft CorporationDouglas Aircraft CompanyDouglas Aircraft CompanyNACA - Langley LaboratorySubcommittee on Aircraft StructuresBureau of AeronauticsSubcommittee on Aerodynamic Stability

NACA - Langley LaboratoryNACA - Langley LaboratoryAircraf t Indus t r ies Associa t ion of

Air Research and Development CommandGlenn L. Martin CompanyConvairS u b c m i t t e e o n Vib ra ti on and F l u t t e rMcDonnell Aircraft CorporationBureau of AeronauticsNACA - Langley LaboratorySubcommittee on Vibration and F l u t t e rDouglas Aircraft Company

and Control

America

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MITCHAM, Grady L.MOL?3L;LA, Roslo J.MOORE, Lt. Col. John T., Jr.M O M , Herbert D.

MORDFIN, LeonardMORGAN, Homer G.MUGLER, John P., Jr.MULHERN, John J.MULTHOPP, HansMUNDO, Charles J.MUNTER, Paul L.MURRAY, James J.

MYERS, Boyd C., I1MYKYTOW, Walter J.

NAUGHTON, S. K.NEELY, Robert E.NELSON, Lloyd W.NEWBERGER, Eli S.NEWBY, Clinton T.NEWELL, Joseph S.NOREM, Allan G.NORTON, David A.

O'BRIEN, R. J.OFFTERMATT, Lt. C m d r. Wilbur F.OGNESS, Arthur M.O ' M A L m , James A., Jr.OSBORN, Earl P.

PAIMER, Carl B.PAYSON, Peter

PEARSON, Albin 0.PELOUBET, Raymond P.PERSH, JeromePETERS, Roger W .PETERSON, James P.PFAFT, George C., J r .PHILLIPS, Franklyn W.PHILLIPS, William H.PEBCE, Harold B

PIERPONT, William G.PINES, SamuelPOLHAMUS, Edward C.

U. S. Naval Ordnance Experimental UnitNaval Air Material CenterAir UniversityAircraft Industries Association of

National Bureau of Standards

NACA - Langley LaboratoryNACA - Langley LaboratoryNaval Air Development StationGlenn L. Martin CompanyArma Div., American Bosch h a orp.Republic Aviation CorporationOffice of Ordnance Research,

NACA HeadquartersSubcommittee on Vibration and Flutter

America

u. s. Army

Wright Air Development CenterFairchild Aircraft DivisionLockheed Aircraft CorporationCivil Aeronautics AdministrationSubcommittee on Aircraft LoadsNorth American AviationAerophysics Development CorporationBoeing Airplane Company

Wright Air Development CenterBureau of AeronauticsNorthrop AircraftSubcommittee on Low-Speed AerodynamicsGrumman Aircraft Engineering Corp.

NACA HeadquartersSubcommittee on Aircraft Structural

NACA - Langley LaboratoryConvairNaval Ordnance LaboratoryNACA - Langley LaboratoryNACA - Langley LaboratoryGlenn L. Martin CompanyNACA HeadquartersNACA - Langley LaboratoryNACA - Langley LaboratoryBeech Aircraft CorporationRepublic Aviation CorporationNACA - Langley Laboratory

Materials


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PRESS, HarryPRIDE, Richard A.

RACISZ, S tan ley F.RAPHAEL, ColemanRARING, Richard H.RAY, George D.RAYBURN, LeoREEVES, Walter R .REGIER, Arthur A.REID, Dr. Henry J . E .REL;LER, John O . , J r .RHODE, Richard V.RICHARDSON, W i l l i a mB.ROBERTS , W i l l i a m M.ROBERTSON, Donald K.ROC&, Jean A.

RODDEN, W i l l i a m P.ROLFE, R i a l E. , Jr.ROSCHE, Melvin G.ROSECRANS, Richard J .ROSEN, B . WalterROSENBAUM, RobertROTH, A. L.ROTIWYER, E a r lROZELLE, Harold L.RUMSEY, Charles B.RUNSTAD, Harold J .RUNYAN, Harry L., J r .

SALZBERG, Leo F.SAUNDERS, John J .SCRNEIDER, W i l l i a m C .SCHNITT, ArthurSCHREIBER, RalphSCHUMACHER, E rnes t A .SCHUMACHER, John G.SCROOC, D. J .SECKIER, SamSEIDE, PaulSENTKECR, Lawrence C.SHARPE, Lawrence W.SHATUNOFF, S tanley

SHAW s. L.SHORTAL, Joseph A.SHOWERS, Nathan

NACA - Langley LaboratoryNACA - Langley LaboratoryB e l l Air cra f t Corpora tionRepublic Aviation CorporationNACA HeadquartersCommittee on Aircraft ConstructionRaytheon Manufacturing CompanyConvairSubcommittee on Vibration and FlutterNACA - Langley LaboratoryNACA - Ames LaPoYatoryNACA HeadquartersNaval Ordnance LaboratoryArnold Engineering Development CenterGlenn L. Martin CompanyAF Development Field Office, NACA

Langley Aeronautical Laboratory

North American AviationMcDonnell Aircraft CorporationNACA HeadquartersNACA - Langley LaboratoryAvco M anufacturin g Corp orationSubcommittee on Vibration and FlutterDouglas Aircraft CompanyGoodyear Aircraft CorporationRyan Aeronautical CompanyNACA - Langley LaboratoryBoeing Airplane CompanyNACA - Langley Laboratory

Wright A i r Development CenterGlenn L. Martin CompanyBureau of AeronauticsBel l Aircraf t Corpora t ionRadioplane CompanyGlenn L. Martin CompanyConvairRam0-Wooldridge CorporationBureau of AeronauticsRamo-Wooldridge CorporationCivil Aeronautics AdministrationA i r Force Cambridge Research CenterArma Di vis ion , American Bosch Arma

Douglas Aircraft CompanyNACA - Langley LaboratoryWright A i r Development Center

Corporation


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SHUFFI;EBARGm, C. C .SILSBY, Noman S.SKIMMIN, Kenneth D.SMITH, Randall C .SMITH, Turner L.SOCKS,, G l e n n A .

SOULE, Hart ley A.SPAHR, J . RichardSPAUUIING, E . H.SP-, Robert F.STACK, JohnSTAHL, Harold F.STAUFFER, Warren A.STEIGER, Arthur R .STEIN, Bland A.STEVENS, C l i f t on D.STEVENS, John E .STEVENSON, C . H.STONE, Ralph W . , Jr.STONEY, W i l l i a m E . , J r .STOWELL, Elbridge Z.STRASS, H . K u r tSTRINGHAM, R. H.S Rodney D.SUTLIFF, JohnD.SUTTON, Fred B.SWANSON, L t . Col. Arthur R.

SWANSON, Theodore B.SWIHART, John M.

TARGOFF, Walter'IIARNOWER, GeraldTETENS, Robert C .THEODORSEN, D r . TheodoreTHIBODAUX, Joseph G., J r .THOMPSON, Floyd L.THOMPSON, Robert F.THOMSON, Robert G.THORITE, Charles J .THORSON, Kenneth R .TOPP, LeRoy J .'IIRACY, Alfred C .TROHA, Charles C .

TUOVILA, W . J.UJIHARA, Ben H.UNDERWOOD, W i l l i a m J .

NACA - Langley LaboratoryNACA - Langley LaboratoryWright Air Development CenterFa i r ch i ld A i rc ra f t D ivi s ionBall is t ic Research LaboratoryChrysler Corporation

NACA - Langley LaboratoryNACA - Ames LaboratorySubcommittee on Aircraft StructuresBureau of AeronauticsNACA - Langley LaboratoryChance Vought AircraftLockheed Ai rc ra ft CorporationA i r Force Special Weapons CenterNACA - Langley LaboratoryGoodyear Ai rc ra ft CorporationSubcommittee on Aircraft StructuresCommittee on Aircraft StructuresNACA - Langley LaboratoryNACA - Langley LaboratoryNACA - Langley LaboratoryNACA - Langley LaboratoryDouglas Aircraft CompanyConvairRyan Aeronautical CompanyNACA - h e s LaboratoryOffice of the Assis tant Secre tary

Aero-General CorporationNACA - Langley Laboratory

of Defense

Cornell Aeronautical LaboratoryChance Vought AircraftTemco Aircraft CorporationRepublic Aviation CorporationNACA - Langley LaboratoryNACA - Langley LaboratoryNACA - Langley LaboratoryNACA - Langley LaboratoryU. S. Naval Ordnance T e s t Sta t ionBoeing Airplane CompanyBoeing Airplane CompanyDouglas Aircraft CompanyBureau of Aeronautics

NACA - Langley LaboratoryNorth American AviationNACA Liaison Officer,

Wright-Patterson AFBx v i i


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UPDEGRAFF, Richard G.UTTER, Emmet

VANDREY, Dr. FriedrichVETERITO, A. MichaelVOLLMECKE, Albert A.VOLLO, Samuel D.VOSTEEN, Louis F.

WA L L S , James H.WATKINS, Charles E.WATSON, Robert E.WEBB, Everette L.WEIDMAN, Deene J.WEINBERGER, Robert A.WEISMAN, YaleWEST, F. E., Jr.WESTRUP, Robert W.WESTVIG, RogerWHALFY, Richard E.WHELDON, Wilbert GWHITE, Richard P., JrWILLIAMS, Walter C.WILLIAMSON, Leonard H.WILSON, Herbert A., Jr.WILSON, James N.WOLCOTT, Verne V.WOOLSTON, Donald J.

YATES, William B.YORK, Joseph E.YOUNG, George E.YOUNG, Willis H.

ZENDER, George W.ZISFEIN, Melvin B.ZLOTNICK, Martin

NACA - Langley LaboratoryBeech Aircraft Corporation

Glenn L. Martin CompanyFairchild Guided Missile DivisionCommittee on Aircraft Construction

Bell Aircraft CorporationNACA - Langley LaboratoryBureau of AeronauticsNACA - Langley LaboratoryBoeing Airplane CompanyBoeing Airplane CompanyNACA - Langley LaboratoryBureau of AeronauticsSubcommittee on Aircraft LoadsNACA - Langley LaboratoryNorth American AviationNorth American AviationNACA - Langley LaboratoryNorthrop AircraftCornell Aeronautical LaboratoryNACA High-speed Flight StationCivil Aeronautics AdministrationNACA - Langley LaboratoryJet Propulsion Laboratory, C.I.T.ConvairNACA - Langley LaboratoryGlenn L. Martin CompanyLockheed Aircraft CorporationNorthrop AircraftBureau of Aeronautics

NACA - Langley LaboratoryBell Afrcraft CorporationAvco Manufacturing Corporation

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AIRCRAFTLOADS


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U

. w"#-Nm1-75363 f 'FACTORSAFFECTING LOADS AT HYPERSONIC

By Ar thur Henderson, Jr., and Mitch el H. Bertram


SUMMARY

This paper gives a brief summary of current loads information a thypersonic speeds. Sev era l methods which th e design er can employ i nestimating th e loads on various a ir c r a f t components ar e discussed. Thepaper deals with the characterist ics of both slender and blunt configura-t ions and touches upon the effects of boundary-layer and aerodynamicin ter ference .

INTRODUCTION

The calculat ion of loads a t hypersonic speeds requ ires th e use oftechniques with which many desig ners ar e no t very f am ili ar . The methodsbased on l inear o r second-order theory, which were widely used a t super-sonic speeds, are inadequate for s lender configurat ions a t hypersonicspeeds and, of course, a re completely inap plica ble t o configurationswith blunt noses o r leading edges.

I n t h i s p ap er it i s shown t h a t certain simplifying features whichallow good design approximations of lo ad s t o be made wit h a minimum of

e f f o r t e x i s t a t hypersonic speeds. I n ad dit ion , some of t h e unsolvedproblems a sso cia ted with hypersonic phenomena m e poin ted ou t.

SYMBOLS

a

A

C

I

Cn

speed of sound

constant

lo ca l chord leng th

mean, aerodynamic cho rd

sec tio n normal-force coeff icien t


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2

CP

d

h

K

2

M

S

t

v

X

U

7

6

A

E

0

h

a

7

cp

loca l su r f ace p re s su re coe ff i c i en t

maximum body diameter

height of wedge

dhypersonic si m il ar i t y parameter, M -1

length of nose o r wedge

Mach number

pressure

radius

Reynolds number

a rc l eng th

thickness

v e l o c i t y

distance from nose o r 1ea.ding edge i n body-axis system

angle of at tack

r a t i o of s p e c i f i c h e a ts

f l ap de f l ec t ion ang le

incremental value

distance between adjacent streamlines

cone shock angle

sweepback angle

cone semiapex angle

time

meridian angle 19a


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3

Subscripts :

0 2 free-stream condit ions

B(W)

MAX maximum

S shoulder

t based on thickness

d based on diameter

body i n presence of wing

DISCUSSION

There a re sev era l methods which the designer can employ i n a rr iv in ga t an estimate of th e loads on th e various a ir c r a f t components. Beforedi sc us si ng them, however, it i s i n s t ru c t ive t o consider, qua l i t a t i ve ly,how hypersonic phenomena differ from supersonic.

Although hypersonic flow introduces many problems which were notencountered a t supersonic speeds, it also in t roduces cer ta in s impl i fy ingfea tur es; and aerodynamicists have not been long i n taking advantage ofthem. For example, one source of simplification a t hypersonic speeds i st h e f a c t t h a t , i n the exact shock equations, the Mach number term i susually squared and often appears i n th e denominator. Thus, as the Machnumber increases, these t e r m s become insignif icant; thus relat ivelysimple expressions oft en yie ld accurate approximations for ce rta in f lowproper t ies a t hypersonic speeds.

Slender Configurations

Characterist ics of hypersonic f low. - Some simplifying features ofhyper sonic f l ow a re i l l u s t r a t ed i n f i gu re s 1 and 2. One of the character-i s t i c s of hypersonic f low i s i t s tendency toward two-dimen sionality wheni n contac t wi th s lender bodies or surfaces. (See f i g . 1.) The upperhalf of f igure 1 depicts a sharp-leading-edge sweptback w i n g i n a lowand i n a high Mach number flow f i e l d . There ar e two stream lines th e samedis tance E apar t . A s shown by the dashed l in es , th e f ie ld s of influencefrom each disturbance point along the leading edge spread across the w i n g

i n supersonic f low, whereas they ar e confined t o a r el at iv el y narrowregion i n hypersonic flow.p a i r w i l l s t r i ke t he l ead ing edge l a t e r t han the l e f t one, t he t ime la g

I n addit ion, th e r i gh t s t reaml ine of each

being AT = E tan *. Obviously, as th e Mach number in cr ea se s, t h e time$0

M a


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4

lag decreases; thus a t high Mach numbers the right streamline strikest h e leading edge a t almost the same time as t h e l e f t one.as f a r a s t h e f l u i d i t s e l f i s concerned, it f e e l s as though it i s p r a c t i -cally two dimensional.

Consequently,

The bottom hal f of f ig ur e 1 depicts the fundamental b a s i s of t h e

generalized shock-expansion method as applied t o s lend er three-dimensionalbodies. A s Eggers and Savin ( r e f . 1) have shown, so long as the d iver-gence of streamlines along the body i s negl igibl e, th e f low on the bodysurface and the associa ted f low f ie ld w i l l be essential ly two-dimensionali n natu re; conseque ntly, two-dimenqional shock-expansion the ory can beused t o analyze the f low about slender bodies of revolution.

Hy-personic similarity l a w. - The designer has another powerful to o la t h i s d i sposa l i n th e form of the hypersonic s im i la r i t y law (see , fo rexample, refspoin ts on s imi la r ly shaped bodies a re id en t ic a l i f , for the two bodies,t h e product of free -stre am Mach number and thi ckn ess r a t i o i s a constant .

2 t o 4), which s ta tes tha t the pressures a t corresponding

The physical concept behind the hypersonic similari ty l a w i s i l l u s -t r a t e d q u a l i t at i v e l y i n f i gu re 2. Two marbles a r e shown, each ro l l i ngtoward i t s own wedge. The upper marble w i l l r i s e a he igh t h i n t helength 2 1 wi th the ve loc i ty VI, while the lower marble w i l l r i s e t hesame he ig ht h i n the longer length 22 = A 2 1 but with the higher veloc-i t y V2 = AV1. The r a t i o of leng ths and ve lo ci t ie s i s such th a t bo thmarbles r i s e the same height h i n th e same leng th of time; th at i s ,they both experience the same change of ve lo ci ty and, consequently, eachmarble w i l l impart the same amount of momentum t o i t s part icular wedge.If the marbles are thought of as a i r molecules and the wedges as cor-responding slopes on two s imi lar bodies, a direct analogy with the hyper-son ic s imi l a r i t y l a w i s immediately apparent.

The approximate region i n which th e hypersonic s im il a ri ty l a w i sapplicable has been determined by Lees (ref. 5 ) t o be about as shown i nf i gu re 3 for cones. This region i s determined by the condit ion that thecone shock angle O s i s l e s s tha n 24'. Thus, t h e maximum cone anglefor good corre la t ion a t hypersonic speeds w i l l be about 20.revolution such as og ives a r e e s sen t i a l l y con ica l a t the nose and decreasei n s lopes the rea f te r. Therefore, i f the nose of any pointed body i sabout 20 o r l e s s , it should corre la te w e l l w i t h t h i s l a w.t h i s means the f ineness r a t i o should be about 3 o r more.

Bodies of

For ogives,

Figure 4 presen t s t he p re s su re- r a t i o d i s t r i bu t ion on ogives.

so l id l in es a re the ch ara c te r i s t ic so lu t ions of Rossow ( r e f . 4), each ofwhich i s for a t least two different combinations of rJr, and Z/dwithin the range shown a t th e lower ri gh t. Although M, = 1 2 w a s t h e

The


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5

highest value of used in th e calculat ions, it should be pointed outt h a t t h i s v a l u e i s not meant t o be taken as an upper l i m i t .m e the tangent-cone approximations of Pro bstei n and Bray (r e f . 6). ForK 2 1, they appl ied the tangent-cone approximation t o Lees' resul t whichi s fo r the case when the shock l i e s fa i r l y close t o the body; and f o r

K < 1, t h a t i s , when the shock i s well removed from the surface of thesle nd er bodies, the tangent-cone approximation i s app li ed t o Kd,rmsh'sr e su l t i n l i n em ize d supe r son ic flow.

Also shown

Van Dyke has pointed out i n his work on the hypersonic small-disturbance theory ( r e f s . 7 and 8) t h a t t he r ange of app l i cab i l i t y o fthe hypersonic s imi lar i ty l a w can be extended t o th e t ran son ic range byrepla cing th e Mach number term with the Prandtl-G lauert sj in il m it y fa c-

t o r fK. he degree t o which t h i s co rre lat ion i s successful i si l l u s t r a t e d i n f ig u re 5 for cones with semiapex angles of 5O, loo, l5O,and 20'. I n t h i s f igure $/ tan2 u i s p lo t t ed aga ins t \ I F 1 a n u

f o r a Mach number ran ge from 1.15 t o hypersonic speeds.ended when sonic velocity appems on the cone surfaces.i s s e e n t o b e ex c e ll e n t.

Each curve i sThe correlat ion

The correlat ion for b l u f f cones as suggested by Newtonian theoryi s presented in f i gu re 6, where Cp s in2 0 i s p lo t t ed aga ins t a. Fort h e ranges of Mach number and u shown, a good approximation t o th e

Ipressure on the surface of a bluff cone i s cp -2.2.

sin2 0

Shock-expansion the ory .- The use of two-dimensional shock-expansiontheory t o pre dic t the pressures on slender bodies of revoluti on a t zeroangle of att ack a t hypersonic speeds i s w el l known. Eggers and h i sa s s o c i a t e s ( r e f s . 1 and 9 ) have shown tha t , p rovided condi t ions a t thenose are known from either conical theory or experiment, the gener alize dshock-expansion method can be used f o r slender bodies of revolution a tangle of atta ck.

Figure 7 shows a comparison of th e shock-expansion theory d t h ex peri-ment f o r an og iva l 6ose a t an angle of at tack of 150 and a free-streamMach number of 5.05. The symbols show the experimental pressure coef-f ic ients a long the top , s i d e , and bottom meridians. The th eo re ti ca l pre-di ct io ns begin with th e assumption of c onic al flow a t t h e nose.curves use the theoretical cone approximation of Savin ( ref . 10) as t h es t m t in g point f or th e shock-expansion calcula t ions , and th e dashed curveuses e xperimenta lly determined conditions on th e nose cone as the s ta rt in gpoin t fo r the ca lcula t ions . Obvious ly, a re l i ab le th eo re t ic a l method i spreferable f o r design work. It i s seen th a t on t h e bottom meridian, whichwould be of most interest for loads considerat ions, the shock-expansion

The solid


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6

calcu lat ion s agree with experiment for both the t he or et ic al ly and th eexperimental ly determined start ing .condit ions; that i s , S a v i n ' s t h e o r e t i c a lcone results combined with shock-expansion theory w i l l give good desig:iestimates of maximum loads on sharp noses a t angle of attack.

Effect of Blunting

I n cases where high heat - tran sfer ra te s a re expected, t he use ofblunt leading edges and noses i s dic ta ted .re su l t s fo r both two-dimensional c i rcu lar cy l inders ( r e f L . 11 and 1 2 )and hemispherical-nose bodies of rev olu tio n (r e f. 1 3 ) .z i r cu l a r cy l inde r s ( f i g . 8) would be appli cabl e both t o bodies of revolu -t l o n a t high angles of at ta ck and t o the leading edges of bl unt sweptbackwings.It shows the manner i n which the press ure m t i o va rie s w ith meridian angle,and it i s good f o r a wide range of sweepback an gl es .Mach number inc re as es , t h e band of exp erim enta lly determin ed pr es su re

ratios converges toward the theoretical curve of Goodwin (ref. 12) shownby th e dashed li ne . Penland ( r e f . 11) has shown t h a t pmax can be de te r-mined on yawed circular cylinders for sweepback angles from 0 t o about75 a t M, = 6.9 by using the normal component of &. Thus, the abso-lu te pressure d i s t r i bu t io n on the windward s ide of yawed circulas cylinderscan be obtained.

Figures 8 and 9 present

The r e su l t s fo r

Figure 8 i s e s s e n t i a l l y a double corre la t ion of pressure ra t ios .

Also note tha t , as t h e

The results for hemispherical noses in f igu re 9 show excellent agree-In t h i s f i gu re t he p re ssu re-coeff ic i ent r a t i oent with Newtonian theory.

i s p lo t t ed aga ins ts een, t he p re s su re -coe ff i c ien t r a t i o i s independen t of Mach number. Thecurve of C P , w aga ins t & i n the upper r i gh t of th e f ig ure shows

t h a t , f o r values of & grea ter than about 4, %,w i s e s s e n t i a l l ya constant on the order of 1.8. Thus, with the aid of Newtonian theory,Cp can be close ly estimated, and f o r values of M, above about 4, h eCp di st ri bu ti on w i l l be e s sen t i a l l y i nva r i an t w i th &.

s / r, which i s t he a r c ang le i n r ad i ans . As can be

A s w a s mentioned prev ious ly, many of t h e exa ct flow parameters canbe c los ely approximated with simple expressions i n th e hypersonic-flowregime. For example, for y = 1.4 and Mm >> 1, it can be shown t h a tthe r a t i o of f ree -s tr eam s t a t i c p re ssu re t o s t agna t ion pre s su re on ablunt-nose body i s approximately 0.777/&'. ( T h i s r a t i o i s determined

.L

\


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7

i n the appendix.)incurred by using the approximation i s shown by the sketch.t h e e r r o r i s only 4 percent and itdecreases r ap id ly t he rea f t e r w i thincreasing I%,. It can also beshown that for I%, >> 1, t h e r a t i oof l o c a l pre ss ur e t o maximum pres-sure i s approximately equal t o t her a t i o of l o ca l p re ssu re coe ff i c i en tt o maximum press ure co ef fi ci en t.Consequently, th e r a t i o of th e lo ca labsolu te pressure t o th e f ree-st ream 0 2 4 6 8 1 0

The percentage error in t he t rue va lue of p , / ~A t M, = 3,

% ERROR IN

~~

p aPMAX

s t a t i c p re ss ur e i s given by M a

- = - Mco2 cp ; h a t i s , a t hyper-P, 0.777 Cp,&msonic speeds, th e ab solute pressure a t any poin t on a blunt nose i sd ir e c t ly prop ort ion al t o th e square of th e Mach number.

fo r any given a l t i t ud e , t he abso lu t e p re ssu re d i s t r i b u t i on on a hemi-

I n p a r t i cu l a r ,m

p McQzsphe r ica l nose i s given by - - os2(s / r ) f o r 0 6 s / r 6 1 . 3 radians .pw 0.777

I

The fac t tha t the exper imenta l presswes devia te f rom the theore t ica lpressures beyond about 1 . 3 radians i s due t o a combination of entropy,v o r ti c i t y , and boundary-layer ef fe ct s, which, of course, Newtonian the orydoes not include. For th e Mach numbers considered herein, th e e f f e c t s ar eneg l ig ib l e as f a r as loads m e concerned.however, th es e e f f e c t s become inc rea sin gly important.

As t h e Mach number i s increased,

Figure 10 shows how, as a r e s u l t of en tropy gain , th e sur fac e pre s-

sure a t th e shou lder va ri e s with Mach number. The model i n f ig u re 10 i sa two-dimensional f l a t sl ab with a sonic-wedge le adin g edge.sur es were ca lcu lat ed by simple inv is ci d shock-expansion theory. I t canbe seen th a t , as M, increases , th e shoulder pressure increa ses t o verylaxge valu es. The pre ssu re on th e shoulder of blunt-nose bodies and blun t-lead ing-edge wings would fol lo w t h e same t r e n d wit h Mach number.

The pres-

Figure 11 presents theoret ical ly and experimental ly determined pres-sure distr ibutions on a blunt- leading-edge f l a t p l a t e f o r a free-streamMach number of about 7. The theo re t ica l ly determined pressure d is t r i -bu tio ns were approximated by assuming sonic-wedge leadin g-edge co nd it io ns .Also i n d ic a te d i n t h e f i g u r e i s the va lue of th e pressure r a t i o f o r noentropy gain and zero vort ici ty.pressure exis t s on the f l a t plate was perfect ly acceptable a t lower super-sonic speeds. A t hypersonic speeds, however, t h e las ge entropy gainthrough the normal shock and the large entropy and vort ici ty gradients

The assumption that free-stream stat ic


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induced in th e f low f i e l d by the h igh shock curvature res ul t i n the typeof pressu re di st ri bu ti on shown by th e s o li d curve.culated f o r t h e i nv is ci d flow a t a Mach number of about 7 f o r the sonic-wedge leading-edge configuration shown a t t he upper l e f t .curve i s the exper imenta l pressure d is t r ibut ion for the b lunt - leading-edge p l a t e shown on th e r ig h t a t one value of Reynolds number.ference between these two curves i s due t o the presence of th e boundaryl aye r, I f t h e Reynolds number had been lower, th e boundary la y er wouldhave been thicker and the separation of these two curves would have beengrea ter ; th e converse being tr u e i f th e Reynolds number had been higher.It should a l so be poin ted out tha t , as the Mach number increases, no tonly does the leve l of p/pw a t th e shoulder increase but a l so the ra te ofdecr ease with distan ce becomes le ss , s o tha t the ent ropy and vor t ic i tye f f e c t s a r e spre ad over a gr ea te r dis tan ce a t hi gher Mach numbers.

T h i s curve was cal-

The dashed

The dif-

Effec t of Boundary-Layer Separation

When re a l f l u id ef fe ct s , including boundary layer s, ar e broughtin to th e pictu re, th e consequences of boundary-layer sepa ratio n must a l s obe considered.important, although it can sometimes be neglected.

A t hypersonic speeds boundary-layer separation i s of t en

Figures 1 2 and 1 3 i l l u s t r a t e examples of boundary-layer separat ionwhich must be c ons ide red and boundary-layer se pa ra tio n which may be,neg -lec ted . Both the body wi th conica l f la re (" f la r ed sk i r t " ) shown i n f ig-u re 12 and the body with flapped wing shown in f i gu re 1 3 were tes te d a tM, = 7. (See refs . 1 4 and 15, respect ive ly. )se pa ra ti on poin t moves rearward along the f lared-skir t body with increasingReynolds number i s i nd i ca t ed by the so l id l i ne in f i gu re 12.press ure-c oeffi cient di st r ib ut io ns f or two extreme posi t ions are shown

above with corresponding symbols.design for separated or unseparated f lows i s obvious. For unseparatedf low the s k i r t p re s sure i s about what would be expected in the absenceof visco s i ty, whi le the laminar separa t ion region ess ent ia l l y pro tec tst h e s k i r t from direc t contac t wi th the f ree s t ream.

The manner i n which t h e

The body-

The importance of knowing whether t o

On the other hand, a larg e port ion of the upper surface of the w i n gwi th t r a i l ing -edge f l a p ( f i g . 13) i s i n a separated-flow region and therei s es se nt ia l l y no effec t on the upper sur face pressure coeff ic ie nt . F ig-ure 13 shows the f lap deflected 16O; however, the same effects would betr u e with a negative f l ap deflect ion. The loads on th e upper surfacesof wings a t angle of at tac k i n hypersonic f low are es se nt ia l ly negl ig ib lewhether se para t ion e xi s t s or not; the difference between free-stream pres-sure and vacuum i s s o small i n comparison with th e pres sures on th e lowersurface th a t , for a l l pr ac t i ca l purposes, the upper sur face can be neg-l ec t ed in l oads ca l cu l a t i ons .

25


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The separated flow on the lower surface i s confined t o a r e l a t i v e l ys m a l l region.sur fac e loa ds would be af fe ct ed more than shown i n fi g u re 13.tude of th e loa ds induced would al s o depend on t h e co nd itio n of t h e bound-a r y l a y er, t h a t i s , whether it i s l aminar or turbulent .

I f t h i s sepa rat io n poin t were t o move forward the lowerThe magni-

There i s as yet not enough knowledge about separation a t hypersonicspeeds t o be ab le t o p red ict when o r where separation w i l l o c c w f o re i ther laminar o r turbulent f low.

Aerodynamic Interference

Another f i e l d which i s r e l a t i v e l y unexp lo red a t hypersonic speedsi s th at of aerodynamic int erfere nce and th e ro le th a t in t e r f e r ence p l aysin altering the expected loads on any component.

One phase of t h e in te rf er en ce problem w a s invest igated by building

& = 6.85Some preliminary results are

a scale model of a configuration which had previously been tested a t& = 3.36.i n th e Langley 11-inch hypersonic tunn el.presented herein.

(See refs . 16 and 17.) This model w a s t e s t e d a t

Figures 1 4 and 15 present t he spanload d i s t r ib ut io ns on the wingalone and on the w i n g i n the presence of th e body a t an angle of a t t a c kof 15' f o r I& = 3.36 and & = 6.85, respect ive ly. The ov er a l l t rend sof t h e r e s u l t s a t & = 3.36 a re abou t what would be exp ect ed . Ther e s u l t s a t M, = 6.85 show the la rge loca l ized effec t which the th ickboundary la y er pla ys i n in te rf er en ce between adj ace nt components. Theindic ated po si t i on of t he boundary lay er w a s taken from sch l ie ren p ic tur esa t a = 0 on t h e sharp-nos e body. The th ic kn es s and condit ion of th e

boundary layer a t t h e wing-body jun ctu re a t a = l 5 O i s not known. Notea l s o th e e ff e c t of nose shape on t h e loadings. The blu nt nose decreasedthe wing loadings.Mach number, it i s t o be expec ted tha t t he g rea t e r l o s se s i ncu r r ed by adetached shock a t higher Mach numbers w i l l more se r ious ly aff ec t th eloadings not only on the wing but also on a l l components within the regionof inf luence of the h ighly ro ta t io na l pa r t of t h e flo w f i e l d a s s o ci a te dw i t h b l u n t noses.

Although the decrement was not appreciable a t t h i s

The shock-expansion theory predicts the loading on the wing alonea t & = 6.85 f a i r l y wel l. The . resu l t s of the & = 3.36 t e s t s a renot a f a i r t e s t of t h e adequacy of shock-expansion t heor y sinc e a ta = 15 O t h e leading-edge shock i s detached a t & = 3.36.

Figure 16 prese nts th e i nterf erenc e loading on th e body due t o thepresence of t h e wing a t an angle of a tt ac k of l 5 O f o r I& = 3.36 on the

I 245


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1c.

sharp-nose body and Mm = 6.85Th e or i ent a t io n of th e wing and body wi th respect t o the load- dis t r i but io ncurve i s as indicated.

on the sharp- and blunt-nose bodies.

Mach number ap pa re nt ly does no t p la y an i mp or ta nt r o l e i n i n t e r -f e r e n c e e f f e c t s i n t h i s Mach number range, as evidenced by the fac t tha t

t h e genera l t rends of t h e interference loading curves on the sharp-nosebody a t . Mm = 3.36 and = 6.85 do not d i f f e r widely. The e f f e c t ofnose shape on body interference loadings i s evidenced by th e re la t i vedisplacement of t h e curves with the square and diamond symbols, and, asalready mentioned, th e sig nif ic ance of t h i s type of in t er fere nce w i l lprobably increase with increasing a. Also of i n t e r e s t i s t h e f a c tt h a t t h e maximum in ter ference loading f o r each of t he t h r ee curves w a s50 t o 60 per cen t of the corresponding body-alone loa ding .

CONCLUDING RFMARKS

This paper has summarized briefly current loads information a thypersonic speeds. Se ve ra l methods which th e de sign er can employ i nest imating t h e loads on various aircraft components have been discussed.The paper has cons idered the charac ter i s t ics of b o t h slender and bluntconfigurat ions and the effects of boundmy-layer Ad aerodynamic inter-fe re nc e. Many problems s t i l l confront t h e designer - t h e e f f e c t o n t a i lloads of t h e wing f l o w f i e l d and i t s associated high-energy wake and thee f f e c t of th e body f low fi e l d and i t s h igh ly ro t a t i ona l f l o w f o r b lun t - nose bod ies . In addi t ion , th e eff ec t on loads a t hypersonic speeds ofthe i ne r t deg rees of freedom of the components of the a i r (molecularv ib ra t ion , d i s soc i a t i on , and ion i za t ion ) i s e s s e n t i a l l y unknom.


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PSPENDIX

PRESSURERELATION APPROXIMATIONS

The ratio of free-stream static pressure to stagnation pressure fora blunt-nose body can be calculated approximately from the following exactrelation (see ref. 14, eq. (100))

1

+ 1, equation (1) can be written:ince - - 17 - 1 7 - 1

As becomes large, equation-(2) is closely approximated by

For 7 = 1.4, equation (3) becomes

28


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In addition, since

and

cp = -(x - I)402 pm

the ratio of local surface pressure coefficient to maximum surfacepressure coefficient becomes

which, with the aid of equation (4) , yields for large Ea,

,


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REFERENCES

1. Eggers, A. J . , J r. , and Savin, Raymond C.: A Unified Two-DimensionalApproach t o th e Ca lcu lat ion of Three-Dimensional Hypersonic Flows,With Application t o Bodies of Revolu tion. NACA Rep. 1249, 1955.

(Supersedes NACA TN 2811.)

2. Tsien, Hsue-Shen: S im il ar it y Laws of Hy-personic Flaws. Jour. Math.and Phys., vo l. XXV, no. 3, Oct. 1946, pp. 247-251.

3. Ehret, Dorris M ., Rossow, Vernon J . , and Stevens, Victor I . : AnAnalysis of the Applicabil i ty of th e Hy-personic S im il ar it y Law t othe Study of Flow About Bodies of RevolutiGn a t Zero Angle ofAttack. NACA TN 2250, 1950.

4. ROSSOW, Vernon J . : App licab ility of th e Hypersonic Si mi la ri ty Rulet o Pressure Dist r ibut ions Which Include the E ffe cts of Rotat ionf o r Bodies of Revo lution a t Zero Angle of Attack. NACA TN 2399,1951

5. Lees, Le st er : Note on th e Hypersonic S im il ar it y Law f o r an UnyawedCone. Jo ur . Aero. Sci . (Rea ders ' Forum), vo l. 18, no. 10, Oct. 1951,pp. 700-702.

6. Pr ob st ein, Ronald F., and Bray, Kenneth N. C. : Hypersonic Similarityand th e Tangent-Cone Approximation f o r Unyawed B o d i e s of Revolution.Jour. Aero. Sci. (Readers' Forum), vol. 22, no. 1, Jan. 1955,pp. 66-68.

7. Van Dyke, Milton D. : Applications of Hy-personic Small-DisturbanceTheory. Jou r. Aero. Sci ., v ol. 21, no. 3, Mar. 199, p. 179-186.

8. Van Dyke, Milton D. : A Study'of Hypersonic Small-Disturbance Tbeory.NACA Rep. 1194, 1954. (Supersedes NACA TN 3173.)

9. Eggers, A. J . , J r. , Savin, Raymond C., and Syvertson, Clarence A.:The Generalized Shock-Expansion Method and I t s Applicat ion t o BodiesTraveling a t High Supersonic Air Speeds. Jou r. Aero. Sci ., vol . 22,no. 4, Apr. 1955, pp. 231-238, 248.

10. Savin, Raymond C . : Application of the Generalized Shock-ExpansionMethod t o I ncl ine d Bodies of Revolution Travelin g a t High SupersonicAirspeeds. NACA TN 3349, 1955.

ll. Penland, J i m A .: Aerodynamic Characteri8tics of a Circular CylinderNACA TN 3861,t Mach Number 6.86 and Angles of Attack.6p t o 90'.

7.-

1957. (Supersedes NACA RM L54Al4.) t..> 30

-A.,*&*


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14 .f

12. Goodwin, Glen, Craeger, Marcus O., and Winkler, Ernest L.: Investi-gation of Local Heat-Transfer and Pressure Drag Characteristics ofa Yawed Circular Cylinder at Supersonic Speeds. NACA RM A55H31,1956.

13. Crawford, Davis H., and McCauley, William D.: Investigation of theLaminar Aerodynamic Heat-Transfer Characteristics of a Hemisphere-Cylinder in the Langley 11-Inch Hypersonic Tunnel at a Mach Numberof 6.8. NACA 'I!N 3706, 1956.

14. Becker, John V., and Korycinski, Peter F.: Heat !Transfer andPressure Distribution at a Mach Number of 6.8 on Bodies WithConical Flares and Extensive Flow Separation. NACA RM ~56~22,1956.

15. Ridyard, Herbert W., and Fetterman, David E., Jr.: AerodynamicCharacteristics of a 6-Percent-Thick Symmetrical Circular-ArcAirfoil Having a 30-Percent-Chord !Trailing-Edge Flap at a MachNumber of 6.9. NACA RM L56B24, 1956.

16. Hill, William A., and Kaattari, George E.: Force and Pressure- (iDistribution Investigation to High Angles of Attack on All-MovablePiangular and Rectangular Wings in Combination With a Body atSupersonic Speeds. NACA RM ~ 5 6 ~ 1 2 , 956.

17. Kaattari, G. E.: Pressure Distributions on Piangulaz and RectangularWings to High Angles of Attack - Mach Numbers 2.46 and 3.36.NACA RM A54JI2, 1955.

18. Ames Research Staff: Equations, Tables, and Charts for CompressibleFlaw. NACA Rep. 1135, 1953. (Supersedes NACA TN 1428.)

.


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TENDENCY TOWARD TWO-DIMENSIONALITY OF FLOWAT HYPERSONIC SPEEDS

Figure 1

PHYSICAL CONCEPT FOR HYPERSONIC SIMILARITYLAW

- = V , ~ C C M ? = Kh d2 12

Figure 2

3%


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RANGE OF

24

20

16

CONESEMIAPEXI2ANGLE, 0

a

4

0

APPLICABILITYOF HYPERSONIC SIMILARITYLAW (K-M,~) FOR CONES

a o oM= Q)

HYPERSONIC SIMILARITY LAWAPPLICABLE IN THIS REGIONFOR UNYAWED CONES

2 4 6 8 O M 1 4 1 6 18

APPLICABILITYOF HYPERSONIC SIMILARITY LAWTO SLENDER BODIES

7 Q I = O O

6

5

METHOD OF CHARACTERISTICS

TANGENT-CONE(LEES APPROXIMATION)

TANGENT- GONE(KARMAN APPROXIMATION)

P 4 --pa 3

-

2

I3SMaS12

I3SaS12-

0.2 .4 -6 .8 IOV I

Figure 4


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CORRELATION OF THE,ORETICAL PRESSURE COEFFICIENTSON SLENDER UNYAWED CONES

COMBINED SUPERSONIC-HYPERSONIC SIMILARITY RULE

CPTANu 115 Mq) 10

Figure 5

CORRELATION OFON

THEORETICAL PRESSURE COEFFICIENTSBLUFF UNYAWED CONES

0 IO 20 30 40 50 600-

Figure 61 34


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PREDICTIONOF PRESSURE DlSTRl6UTlON ON OGIVE ATANGLE OF ATTACK

GENERALIZED SHOCK-EXPANSION METHOD; M a5.05

----- EXPERIMENTAL CONECONDITIONS

0 20 40 60 80 100% NOSE LENGTH

Figure 7

INVARIANCEOF CYLINDER PRESSURE RATIOWITHMACH NUMBER AND SWEEP ANGLE

M a 6.9M a 3.9

0 IO 20 30 40 50 60 70 80 90+.DEGFigure 835c .


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INVARIANCE OF PRESSURE-COEFFICIENT RATIO WITHMACH NUMBER FOR HEMISPHERICAL NOSES

GPp, MAX

1.0

.8

.6 0 4 8 1 2

.4 EXPERIMENT

.2

0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8s/ r

Figure 9

EFFECT OF ENTROPY ON SHOULDER PRESSURE OFSONIC- WEDGE FLAT PLATE

24 r20 t16 -

PSp a- 2-

8 -

4 -

I I I I I I I

0 2 4 6 8 10 12 14 16 18M a

Figure 10 36


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20

.2

EFFECT OF BOUNDARY LAYER AND ENTROPY ONFLAT-PLATE PRESSURES

M a W 7 ; y = 1.4L H X

T - C I #

SONIC- WEDGELEADING EDGE

-

0 IO 20 30x/ t

Figure 11

EFFECT OF SEPARATION ON PRESSURE DISTRIBUTIONON BODY OF REVOLUTION

M a " 7 ; Q = 0.8r

Gp .4i DSEPARATION

R d .4

I I I0 2 4 6 8 1 0 1 2

----

Figure 1237'


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EFFECT OF SEPARATION ON PRESSURESONTWO-DIMENSIONAL WtNG FLAP

a=16O; 8=Is0; M a m 7 ; R=1.65XIO6

_/.VACUUM0

UPPER SURFACE

EXPANSION

.2

.4

.6

0 .2 A .6 .e 1.0x / c. e l ' 1 I ' I ' 1 I I

Figure 13

INTERFERENCEON WINGLOAD DISTRIBUTIONAT a=15"M a = 3.36

2.0 kc SHOCK-EXPANSION THEORY

I.5

1.0

.5

0 25 50 75 100% SEMISPAN

Figure 1 4


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INTERFERENCE ON WING LOAD DlSTRlWTlON ATQ = E OM a 6.8 5

I 5 EDGE OF

NOSE

I -'q SHOCK- EXPANSlONI I0 25 50 75 100

O/oSEMISPAN

INTERF EREN CE ON BODY LOAD DISTRIBUTION QU ETO WINGQ =IS0

. I 2

-08

Acn B (w) -04

0

-.04

THEORY

I \ ' i . 9

5 6 7 8 9 10 II 12 13x/d

Figure 16

i d . 39

d


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SPAN LOADINGS DUE TO WING TWIST AT

TRANSONIC AND SUPERSONIC SPEEDS

By Frederick C. Grant and John P. Mugler, Jr.


SuMMP,ziY

Two similar tapered sweptback plan forms with L e same two spanw-seva r i a t i ons of t w i s t have been t e s te d i n t he Mach nmiber range from 0.8t o 2.0. The t e s t re s u l ts showed, i n general, ra th er good agreement withtheo re t i ca l p r ed i c t i ons of th e incremental span loadings due t o t w i s tf o r zero angle of at ta ck . The measured incre men tal span loa din gs duet o t w i s t general ly diminished w i t h increasing angle of at tack throughth e Mach number range .progressively vanished f r o m the t i p inboard wi th increas ing angle ofa t t ack .the re was no diff erenc e i n the span loadings of t h e f l a t and twistedwings.of the incremental loading due t o t w i s t w as s t a r t i n g a t the h ighes t anglesof a t t ack (nea r 20').

A t a Mach n W e r of ,O.g, the incremental loadings

For the h ighes t angles of a t tack (about 20') a t Mach nmiber 0.9,

A t the higher supersonic speeds, a similar v an is hi ng a t t h e t i p s

For angles of attack lower than about 20' a t supe rson ic speeds, noimportant change in t he shape of th e incremental loadings occurred,al though the s t r eng th of the loading diminished with increasing angleof a t t ack .

INTRODUCTION

The t h i n wings of modern high-speed a irp la ne s deform app rec iab lyThe changes i n a i r loading due t o these deformations have

A n aerodynamically important formi n f l i g h t .not been exten s ive ly inves t iga ted .of deformation i s t w i s t , or change in angle of a t t a c k a t a given span-wise s t a t io n on a wing. A s pa r t o f a research program.on the loads duet o wing t w i s t , t w o simple spanwise t w i s t dis t r ibut ions have been tes teda t th e Langley Aero naut ical Iab ora tor y i n th e Mach nmiber range from 0.8t o 2.0. For a complete airplane with stores and nacelles act ing on thewing, the t w i s t di st ri bu ti on along th e span may be r at he r complicated.It i s hoped t h a t the loadings due t o simple t w i s t d i s t r i b u t i o n s w i l l ,by superposit ion, give th e *loadings due t o complica ted d is t r ib ut i ons .

$3


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2

A

b

C

cAV

cn

M

X

Y

a

A

aa

e 0e

_ _* l ee .

5 -- * e *- - -

e a re * ,

SYMBOLS

a sp e ct r a t i o

span

chord

average chord

sec t io n normal- force coe ff ic ien t

Mach nuuiber

dynamic pressure

th ickness

chordwise distance

spanwise distance

angle o f a t t a c k

incremental normal-force c oe ffi ci en t

i ricremental l i f t i n g pressure

sweepback a t quarter chord

t a p e r r a t i o

MODELS

The wings t e s t e d and t h e t w i s t var ia t ions which were bui l t i n a r eshown i n f ig u re 1. The wings tested a t t ransonic speeds had an aspectr a t i o of 4, 45' of sweepback a t th e qua rte r chord, and a t a p e r r a t i oof 0.17. The semispan wing tested a t supersonic speeds had an aspectr a t i o of 3.5, 50 of sweepback a t th e quar ter? chord, and a t a pe r r a t i o

of 0.20.t h e body c e n t e r l i n e t o 3 percent a t and beyond halfway t o t he t i p .The thickness of the supersonic y&ngs w a s a constant 5 percent .caaiber was b u i l t i n t o t h e t r a ns o n ic w ings.

The thick ness of th e trans oni c wings var ied from 6 percent a tA small

All th e wings t es te d had the


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a , * C -b*.a . .,- a. -a

3

0

same 65A-series thickness distribution and the same spanwise variationso f b u i l t - i n t w i s t .which i s a t t a in ed by a l in ea r and quadra t ic var ia t io n wi th spanwisepos i t ion .washed out, f o r the p os i t iv e d i rec t ion assumed in t h i s paper.were te st ed i n each speed range t o provide a reference t o which th etwisted wings might be compared.

The t w i s t angle a t t he t i p w a s 6 O , i n every case,

The t ips a re a t a lower angle 02 a t t ack than the roo t , orFla t wings

INCRFSIENTAL LOADING

Figure 2 shows the span loadings on the f l a t wing and the l inear lytwis ted wing a t M = 1 .6 and a t a = 12'. The di ff er en ce i n th es e spanloadings, o r incremental span loading, i s a l s o shown. Incremen tal spanload ings formed i n t he same manner w i l l be th e ba s is of comparison betweenl i n e a r t h e o r y a n d t h e t e s t r e s u l t s a t the other Mach numbers and angles ofa t t ack .

The incremental loading shown i n f igur e 2 ' i s th e is ol at ed e ff ec tof spanwise wing t w i s t with, of course, the nonlinear influence of angleof at tack and thickness neglected. I f real wings behave as do the wingsof l i ne ar theory, the incremental loading f o r a given spanwise t w i s t d i s -t r i b u t i o n w i l l not change with angle of attack. For t h i s c a s e t h e i n c r e -ment i n normal fo rc e produced by 6 O of t w i s t i s 13 percent of the f l a t -p l a t e n orm al-fo rc e c o e ff i c i e n t. T hi s i l l u s t r a t e s t h e f a c t t h a t , f o r agiven o ver a l l accuracy i n predic t ing the loading on a tw iste d wing, t heaccuracy of prediction of the incremental loading can diminish as t h eangle of at tack increases . -

PREDICTIONS AT ZEBO AJYGLF: OF ATTACK

I n o r de r t o e l i n i n a t e , a s f a r as poss ib le , th e inf luence of angleof at tac k, th e root angle of at ta ck may be se t t o zero. The pred ictedand measured incremen tal span load ings due t o t w i s t with the root angleof a t ta ck se t t o zero ar e shown i n f igure s 3 and 4.

Wings With Linear Twist

Transonic speeds.- Figure 3 shows the comparative theoretical andexperimental increme ntal span load ings for t he t r anson ic l i nea r ly tw i s t edwing. The se ct io n loa din g paramete r Acn C/CAV i s p lo t t ed aga ins t.the spanwise posikionspanwise po si ti on of t h e wing-body jun cture .

2y /b , and the ve r t i ca l da shed l i ne i nd i ca t e s t he

I

' t . .: 1.

. +., c

42


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4

A t M = 0.90, the agreement between th e data and theory i s f a i r l y

(See ref . 1.) The predictiongood.approximating th e presence of th e body.i s be t t e r outboard than it i s nearer th e body.

The theory shown i s a l i f t i ng - su r f ace t heo ry w i t h a provis ion for

A t M = 1.20, t h e r e i s close agreement between the data and theory,even though the va l id i ty of l inear theory i s becoming questionable ast h e Mach nuniber approaches one. The th eo ry used a t supersonic speedsf o r subsonic leading edges i s tha t g iven i n r e f e rences 2 an d 3.addit ion, t h e boundary condit ions w e r e o n l y approxima te ly s a t i s f i e d i nt h e t he o re t i ca l compu ta tions f o r t he t r anson ic w i n g s a t N oattempt was made t o account f o r th e presence of th e body.t h e r e s u l t s a t M = 1.20of t h e body on th e incremental span loadings.

In

M = 1.20.A feature of

i s the apparent absence of any marked influence

Supersonic speeds.- Figure 4 hows the incremental span loadingswfth zero roo t angle of a t ta ck fo r the supersonic l in ea r l y twis ted wingsa t M = 1 .6 and 2.0.

Figure 4 shows t h a t t h e d a t a are about 20 percent lower than pre-d i c t ed va lues . A s predic ted , th e loading i s s l ig ht ly weaker a t t he h ighe rMach number. The shock waves caused by t h e th ic kn ess seem t o have nomore effect on t h e span loadings a tthe leading edge i s supersonic a t M = 2.0 and shock waves due t o th ic k-ness must c e r t a i n l y b e more severe. The theo ry used a t M = 2.0 i sg iven i n r e f e rence 4.

M = 2.0 than a t M = 1.6, although

Wings With Quadratic Twist

Transonic speeds.- Figure 5 shows the incremental span loadingson th e wings with quad rat ic twist as measured and predicted a t t ransonicspeeds.

The agreement with theory i s again rather good a t M = 0.90. Thei s about the same as i t w a s i n the case of thegreement a t M = 1.20

wings with l inear t w i s t .a t M = 1.20.

Again there i s no apparent body

Supersonic speeds.- Figure 6 shows the predicted andmenta l loadings f o r th e wing wi th quadra t ic t w i s t a t M =f o r M = 2.0 are no t ye t ava i l ab l e .

The agreement i s b e t t e r i n t h i s c ase th an it w a s f o r

e f f e c t

measured incre-1.6. Data

t h e l i n e a r l ytwi ste d wing a t t h i s Mach number.as co qa re d wi th abou t 20 pe rcen t fo r t hew ing wi th l i ne a r t w i s t . T h i smust be pa r t l y due t o t h e f ac t t h a t t he ave rage ang le of t w i s t over th eplan form i s lower than it was i n t h e c a se o f t h e l i n e a r l y tw i s te d wings.

The values are o n l y 7 percent l o w e r

I' .. 43


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e $0

5

LIFTING PRESSURE DISTRIBUTIONS

Figure 7 shows the chordwise lifting pressure distribution corre-sponding to two of the incremental span loadings previously shown.lifting pressure coefficient Ap/q is plotted against the chordwisepositionattack, and the spanwise station at which the data were taken is 0.7of the semispan. Distributions for both linear and,quadratic twist areshown. These distributions are typical of other spanwise stations atthis Mach number.shown for both twist variations.

The

x/c; distributions are for Mach number 1.6, zero root angle of

Linear-theory predictions of the lifting pressure are

For the wing with linear twist, the agreement with theory is good.

(See ref. 3 . ) SinceThe level of agreement is comparable %o that indicated by recent pressuremeasurements made on a zero-thickness delta wing.a zero-thickness delta wing exactly satisfies the boundary conditionsof linear theory, the agreement with theory obtained on such a wingtypifies the best that can be expected. To have s imi l a r agreement ona wing with 5-percent thickness is surprising. The agreement for thewing with quadratic twist is even better than that for the wing withlinear twist.agreement observed in the integrated loadings for the wing with quadratictwist.

The fine agreement shown here was reflected in the good

R U D I C T I O N S AT A N G U OF ATTACK

All the incremental loadings that have been shown thus far werefor zero root angle of attack.

incremental loadings will not change with angle of attack, or, in otherwords, the twist will produce the same change in loading whether or notthe wing is at an angle of attack. Of course, this simple predictionis not borne out by the data.

According to the linear theory, the

Transonic Speeds

Figure 8 shows the effect of angle of attack on the span loadings

Dataat M = 0.90. In this figure, instead of incremental loadings, the totalspan loadings are shown for the flat and linearly twisted wings.for angles of attack of bo , 8 O , and 12' are shown. For the transonicwings at angle of attack, incremental aeroelastic twists occurred which

amounted to about 10 percent of the 6 of built-in twist at 12' angleof attack. Figure 8 shows that the shape of the incremental loadings(the vertical difference between curves) changes markedly with angle ofattack while the strength of the incremental loading greatly diminishes.

4%


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6

A tl i n e a r t w i s t carrying th e le ss er load. Between a = bo and a = 8 O ,t he f l ow sepa ra t e s a t t he t i p of t h e f l a t wing, and a ti s sep ara ted outboard of about 60 percent of the semispan.wing a t a = 8 O , however, has much th e same type of span loading as a ta = 4 and the flow appears unseparated. A t a = 12' both wings aresep ara ted outboard of about 40-percent semispan.a l ready small a t a = 12O, e ff ec t ive ly vani shes a t the h igher angles ofa t t a c k . A t th e h igher angles, then, the re i s no di ff er en ce between th ef l a t and twisted wings. Simi lar resul ts have been obtained on the wingwi th quadra t ic t w i s t . A t Mach number 1 .2 t h e r e s u l t s a r e c o ns is te n t w i t htho se t o be shown f o r th e superson ic wings, bu t values w i l l not bepresented.

a = bo, b ct h wings show t h e same s o r t of span load ing, th e wing with

a = 8 the flowThe twisted

The incremental loading,

Supersonic Speeds

I n f i g u r e 9 , th e percent of t he t he or e t ica l loading which must beused t o ob ta in a good f a i r i n g through the data i n th e outboard regions(beyond semispan), where most of the incremental l i f t i s located, i sp lo t t ed aga ins t t h e roo t ang le of a t t ack .t h i s p l o t i s t h e r a p i d d ec re as e of t h e e f f e c t i v e l i n e a r t w i s t with angleo f a t t a c k .t h e M = 2.0 M = 1.6d a t a a r e f o r a subsonic leading edge. A s was mentioned previously, aless accura te predic t ion of the incrementa l loading i s acceptable a t t h ehigher angles of at ta ck . Even i f 100 pe rcent o f t he t heo re t i ca l load ingf o r t h e l i n e a r t w i s t were used t o pred ic t t he l oad ing a t 12' angle ofa t ta ck , th e &>-percent d i fference indica ted by f ig ure 9 would come t oan er r or of about 12 pe rcen t i n p red i c ti ng th e t o t a l l oading . A b e t t e res t imate o f the incremental loading, such as t h e f r a c t i o n s of t he t heo-

r e t i c a l loading indica ted by the curves , could lead t o a n e g l i g ib l e e r r o ri n t h e t o t a l lo ad in g.

The most striking feature of

There i s no marked effect of t h e Mach nurriber, al thou ghdata a r e f o r a supersonic leading edge and the

For the wing with quadrat ic t w i s t , o n l y t h e M = 1 .6 data, orsubsonic-leading-edge data, a re av ai la bl e. However, t h e re i s no reasont o e xp ec t t h a t t h e Mach number e f f e c t s w i l l be any stronger than theywere for t he l i nea r ly tw i s t ed wing . F o r th e wing w i t h quadrat ic t w i s t ,f i g u r e 9 shows t h a t the good predict ion of the incremental loading a tzero angle of a t t a c k i s coupled with a s low drop in e ffec t ive t w i s t ast he ang le of a t t a c k i n c re a se s .p red i c t i on a t zero angle o f attack and more rapid drop with angle ofa t tack observed on the l inear ly t w i s t e d wing.

This cont ras t w i t h t h e r e l a t i v e l y po or er

There i s l i t t l e change in th e shape of th e incremental loadingsf rom 12' a n gl e of a t t a c k t o a w u t 20'.incremental load ings van ish on the outboard region s of t h e wing i n amanner s i m i l a r t o t ha t observed a t M = 0.90.

I n t h e neighborhood of 20

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7

CONCLUDING RESIARKS

At higher subsonic speeds the theoretical predictions at zero angleof attack of incremental span loads due to twist were fairly good.Because of separation effects, these predictions failed as the angle ofattack increased. At the highest angles, there was no difference in theloadings of the flat and twisted wings. At low supersonic speeds, thepredictions at zero angle of attack were better although the validityof the linear theory is becoming questionable.speeds, the predictions at zero angle of attack were generally largerthan the actual loadings. The prediction was better for the wings withlower average twist. At angles of attack up to 12O, factors were appliedto the theoretical incremental loading which give good agreement withthe data.loading steadily diminished with angle of attack.

At the higher supersonic

Through the Mach number range of 0.9 to 2.0 the incremental

REFERENCES

1. Crigler, John L.: Comparison of Calculated and Experimental LoadDistributions on Thin Wings at High Subsonic and Sonic Speeds.NACA TN 3941, 1957.

2. Heaslet, Max. A., and Lomax, Harvard: Supersonic and Transonic SmallPerturbation Theory.Vo l . VI of High Speed Aerodynamics and Jet Propulsion, sec. D,ch. 3 , W. R. Sears, ed., Princeton Univ. Press, 1954, pp. 186-206.

General Theory of High Speed Aerodynamics.

3. Lomax, Harvard, Heaslet, M a x . A., and Fuller, Franklyn B.: Integralsand Integral Equations in Linearized Wing Theory. NAC!. - - p . 1054,1951. (Supersedes NACA TN 2252.)

4. Kainer, Julian H.: Equations. for the Loading on Triangular WingsHaving Supersonic Leading and Trailing Edges Due to Various BasicTwist Distributions. Jour. Aero. Sci., vol. 20, no. 7, July 1953,pp. 469-476-

3 . Boatright, William B.: Experimental Study and Analysis of Loadingand Pressure Distributions on Delta Wings Due to Thickness and toAngle of Attack at Supersonic Speeds. NACA RM ~56114, 956.

I 46


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8

MODEL CONFIGURATIONSTRANSONIC SUPERSONIC

A.4; Ac/4= 45O; X=0.15 A.3.5; AC,=5O0 j X.0.20N AC A 6 5 A 2 0 6 , ROO T N A C A 6 5 8 0 0 5NACA 65A2 03 ,0 .5 b /2 TO TIP

%%T,DEG -6

?$?tT,DEG -6

-I 0 I 0 I2y/ b 2y/b

Figure 1

ILLUSTRATIVE EXAMPLEOF AN INCREMENTAL SPAN LOADINGCY=12" M=1.6

.*r/--

.4

.2

1

INCREMENTAL'I J

0.5 I20

47 Figure 2


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INCREMENTAL SPAN LOADINGS ONWING WITH LINEAR TWISTQ = O O

M=0.90 M11.20

0 .5 I o2y/b

Figure 3

INCREMENN SPANLOADINGSON WINGWITH UNEAR TWISTCY=OO

M.1.6

AM=2.0

Figure 4

48


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INCREMENTAL SPAN LOADINGS ON W ING WITH QUADRATIC TWISTa soo

M.0.90 M.1.20

AC" C

'AVI I I

0 .5 I.o

INCREMENTALSPAN LOADINGS ON WING WITH QUADRATICMllST

azo0; Mz1.6

ACnC

'AV -.I THEORY

-,2' I I0 0.5 I

Zy/b

Figure 6


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INCREMENTAL LIFTING PRESSURES ON TWISTED WINGSM . 1 . 6 ; a z o o ; 2y /ba0 .7

0

- P -.2-.4

LINEAR TWIS T QUADRATIC TWIST

*Ll-&+---

- P THEORY -pb --e HEORY09'- 1

III

JI

I I I I

Figure 7

EFFECT OF ANGLE OF ATTACK ON SPAN LOADINGM=0.90

O -- FLAT WINGD--- TWISTED WING

a = 4 O Q 580 a 1120I1.2*IL,\.4 I

hFa.

0 .5 I

~~ I

I

0 .5 I

1:\0 .5 I

2y/b 2y/b 2y/b

Figure 8s i p


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-0

O/o OFTHEORETICAL50 -

LOADING ----

I I I

-------

I I I


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* .m e

FLIGHT MEASURENENTS AND CALCULATIONS OF WING LOADSAND

LOADDISTRIBUTIONS AT SUBSONIC, TRANSONIC, AND

SUPERSONIC SPEEDS

By Frank S. Malvestuto, Thomas V. Cooney, and E a r l R. Keener

NACA High-speed Flight Station

P re se nt ed i n t h i s r e p o r t i s a summaxy of l o c a l and n e t angl e-of -at ta c k wing-panel loads measured in f l i g h t on s i x a i rp l a ne s .a comparison of the se lo ads measured in f l i g h t with ca lc ul at io ns based onsimple theory i s presented.

I n addit ion,

INTRODUCTION

A t the High-speed Flight Station of the National Advisory Committeefor Aeronautics, ful l-scale research i n t h e f i e l d s of s t a b i l i t y , p erform-ance, and lo ad s i s conducted with a v a r i e t y of completely instrumentedresearch and mi l i ta ry- type a i rp lanes .

I n th e pres ent paper, th e aerodynamic loads a s pe c t of t h i s f l i g h tThe presentation w i l l involve a summary of l oca lesearch i s considered.

and n e t ang le-of -atta ck wing-panel load s measured in f l i g h t on a v a r i e t yof ai rpl ane s flown during th e past 5 o r 6 years. I n addit ion, a prelim-inary comparison of these loads measured i n f l i g h t and the correspondingloads cal cul ate d by simple theory i s presented. The ob jec t of t h i s corn-par ison i s t o a s s es s t h e a b i l i t y of sim ple t h e o r e t i c a l t e ch ni qu es t o p re-d i c t th e f light -measured loads for a var ie t y of conf igura tions .curs ory comparison of th e f l i g h t measurements wit h comparable wind-tunnelresults has been made. I n a gene ra l s ense , t h e f l i g h t resu l t s v e r i f y t h etunnel f indings .been added.

Only a

For the convenience of the reader, a bibliography has

Figure 1 depi c ts wi th plan-view out l in es th e a i rp lanes t o be d is -cussed i n t h i s repo rt . The wing panels are darkened t o emphasize th e

fac t tha t only the wing loadsw i l l be

considered.A n

inspect ion ofthe individual sketches and geometric data shows that there i s a goodcoverage of Wing sweep, plan form, aspect ra t io , and th ickness . I naddi t ion , the X - l E wing has 2 pos it i ve incidence and t h e D-558-11 winghas 3 of posi t ive incidence.f l i g h t s of the se a irp lan es varied,from,*; x 106 t o 6 x 106 per foot .a l t i tu de var ied from 25,000 fe e t t o 65,000 feet.

The fr ee- str eam Reynolds number f o r th eThe


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0 .*..

. a .e. . .* . . e 1 - - - *. - -2

SYMBOLS

A

b

b

C

Cav

Cf

Cn

CN

cNa

CP

K P

H

M

%o

t

X

Y

a

se

*LE

aspect rat o

wing-panel spanflap span

chord

average chord

flap chord

section normal-force coefficient

net normal-force coefficient

slope of wing-panel normal-force coefficient

pressure coefficient

pressure coefficient differential between upper and lower surfaces

altitude

Mach number

free-stream Reynolds number

thickness

distance along x-axis

distance along y-axis

angle of attack

elevon deflection

leading-edge sweep

53


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3

THEORIES CONSIDERED

A few prel iminary remarks regarding the theories used for the wing-panel load ca l cu l a t i ons w i l l be made. The w i n g s ar e assumed t o be r i g idf l a t p l a t e s and of ne gl ig ib le th ickness . In add it i on, t he e ff ec t of t hefuselage interference on the wing loads was approximated by assuming thef u se l ag e t o a c t as a pe r fe c t r e f l ec t ion p lane loca t ed a t the wing-fuselage juncture.load on one panel of a symmetrical wing with i t s root chord coincidentwith th e wing-fuselage juncture. It i s r ea l i z ed t ha t t h i s app roxima tiont o t h e f u se l ag e i n t e r f e r e n c e i s su bj ec t t o improvement; however, it i sf e l t t o be s u f f i c i e n t f o r t h e p r es e nt s t ud y. With th es e assumptions i nmind, t h e wing theo rie s used fo r load predic t ions are g iv en i n t h e f o l -lowing table:

On t h i s ba s is , t h e w i n g load i s pred ic t ed a s t he

Theor ies used f or ca lcu la t io n of w i n g loads -Subsonic

(0 .5 C M < 0.85)~

A l l wings: l inearl i f t i n g s u r f a c e( r e f s . 1 o 4)

Transonic( M = 1.0)

Swept wing: l i n e a rl i f t i n g s ur fa ce( r e f s . 5 and 6)

Unswept wing: two-dimensional f l a tplate; two-dimensional doublewedge (refs. 7

and 8 )

Supersonic(M 2 1 .2)

A l l w i n g s : l inea rl i f t i n g s ur fa ce( r e f s . 9 t o 16)

I n th e subsonic range, fo r a l l wings, l inea r theory w a s applied.:See refs . 1 o 4.) These subs onic ca lc ul at io ns were made up t o a Machnumber of 0.85, alth ough i n th e neighborhood of t h i s Mach number, tran-sonic mixed-flow conditions no doubt exist.cu la ti on s were made only f o r a free-stream Mach number of 1.0. I n t h i srange, f o r the swept wings, th e lin ea r theory presented by Mangler (r e f . 5 )which i s i n essence Jones ' s lender-wing theo ry ( re f . 17) modified forl i ne ar iz ed sonic-f low condi%ions w a s applied. For th e unswept w i n g , a t aMach number of 1.0, use w a s made of t h e r e s u l t s of Guderley and Yoshihaza( r e f . 8 ) f o r a double-wedge se ct io n and th e r e s u l t s of Guderley ( r e f . 7)f o r a f l a t p l a t e of neg lig ib le thick ness . For th e supersonic Mach numberrange, th e well-known l if t i ng -s ur fa ce the ori es were applied.

In t he t r anson ic range, ca l -

i 54'


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LOADINGDISTRIBUTION

I n the discu ssion of f l i g h t res ul ts , the chordwise and spanwiseloadings for the unswept-wing X-1E airplane, the swept-wing D-358-11 a i r -

plane,and the delta-wing JF-10% airplane are considered and then a

fo rc e summary f o r all s ix a i rp l anes i s given.

Some id ea of t he f l i g h t Reynolds number, al ti tu d e, and angle-of-a t tack excurs ions for these a i rp lanes can be determined from figure 2.Reynolds number i s given on a per-foot basis and for free-stream con-d i t i o n s .ber obtained.approximately 1 x 10 6 t o 4 106.approximately 25,000 t o 65,000 f e e t . On the r ight-hand s ide of f igure 2the hatched boundary i s indica t ive of the maximum angle-of-attack excur-sions obtained i n f l ig ht . The discussion of the angle-of-attack wingloads w i l l be within the region shown by the dashed boundary.

The

The open circular symbol represents the maximum Reynolds num-It i s noted th at t h i s f l i g h t Reynolds number va ries from

The al t i tude covers a range from

In f igu re s 3 t o 6 are presented the chord loadings and span loadingsThe s o l id l i ne represents the theory; the openThe dashed line through the open circles repre-

fo r t he X - D w i n g panel.symbol, t h e f l i g h t d at a.s en t s " f a ir ed" f l i g h t dat a .u re 4 ndica te the panel normal-force coe ffic ien tat ta ck a t which the chord and span loadings ar e shown. Consider f i r s tthe chord loadings of f i gu re 3, t h a t i s , t h e v ar i at i on of L?Ep, thel i f t i n g p re s su r e, w it h x / c , the normalized distance from the leading

edge. These re su lt s ar e f o r a span s ta t ion - = 0.46. The symbol b'b ' / 2

denotes the external panel span. The chord loadings are shown f o r Mach

numbers of 0.8, 1.0, and 1.9. For each Mach number th e chord loadings a reshown for t w o angles of a t t ack , a low angle and a high angle.of the high angle of at tack i s l imi ted by the a va i la b i l i ty of the da ta.The angle of a t t ack i s always the angle of attack of the w i n g panel.M = 0.8, the calculated level and variat ion of the chord loading comparesfavorab ly wi th t he f l i g h t da ta .available finite-span unswept-wing theory.shown here i s th e fl at -p la te two-dimensional theor y of Guderley.t h e l e v e l of t h e l i f t i n g p r e ss u r e i s not predicted herein, t he va r i a t i oni s similar t o the f l ight-measured var ia t ion fo r both angles of attack.

The sketches on the left-hand side of fig-CN for the angles of

The magnitude

A t

For a Mach number of 1.0, t h e r e i s noThe th eo re t ic a l var ia t ion

Although

A t supersonic speed and low angle of atta ck, th e comparison of f l i g h tA t the higher angle of at tac k, the loading di s-nd theory i s acceptable.

t r i b u t i o n i s not predicted by theory al though the level of the l o c a l load-can be ca lcu lat ed.two addi t io nal spanwise stat io ns, 'o ne near the roo t and one ne w the t i p ,are shown in figures 5 and 6.

The midspan chord loadings and the chord loadings a t


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a .a .

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I f t he span - load d i s t r i bu t ions ( f i g . 4) are considered, it i s notedt h a t , a t M = 0.8 and M = 1.9, the calculated span loading comparesfavorably with th e flight-measured loading. For M = 1.0, the span loadingwas not calculated, s ince, as mentioned prev ious ly, t h e two-dimensionalr e s u l t s of Guderley were used; however, t h e f l i g h t data have been faired.The shapes of t he span-loading curves stro ngl y resemble each other f o r th e

t h re e Mach numbers shown.

For the swept-wing D-55SII airplane the chordwise and span-loaddi s t r i bu t io ns fo r t he wing panel a r e shown in f igures 7 t o 10.s o l i d l i n e r e p re s e nt s t h e c a l cu l a ti o n s and th e open cir cu la r symbol, th ef l i g h t measurements. The panel normal-force co ef fi ci en ts correspondingt o the angles of a t tac k cons idered are indica ted i n the ske tches on t h elef t -hand s id e of f ig ure 8.a r e f o r a spanwise s t a t io n c lose t o t h e midsemispan location.subsonic and supersonic speeds, the theory al lows the calculat ion of thele v el and va ri at io n of the chord loading except a t the high angle ofattack for the supersonic Mach number. A t M = 1.0, th e measured d i s t r i -bu tion of t he l i f t i n g p re s su re EP i s not ca lcula ted by the l inear

theo ry. Theory giv es a zero loading behind the l inear ized sonic shockt h a t s t a r t s from the leading edge of th e s treamwise t i p of t h e w i n g panel.It i s p o s s i b l e t o o b t ai n a nonzero loading by minor al te ra ti o ns of t hewing-tip geometry s o th a t , fo r the por t ion of t he wing behind the l i ne -ar ized shock, th e l oc a l span increases wi th increas ing longi tudin a l posi -ti on ; and hence l i f t i s produced. (See re f . 17.) A d i scuss ion of t h i sa r t i f i c e i s given in the report by Mangler ( r e f . 5) mentioned earlier.The midspan chord loadings and th e chord loadings near t he ro ot and t i pa r e shown i n f igures 9 and 10 .

Th e

The chord loadings presented i n f i g u r e 7For the

The span loading for the swept-wing D-558-11 i s presented in f i g -u re 8.compares favo rably with th e f l i g h t measurements. For M = 1.0, t h e c a l -cula ted loading , espec ia l ly a t th e h igh angle of a t ta ck ( l l o ) , does notrepresent t he experiment because of th e i na bi l i ty of t he theory t o pre-d i c t t h e l e v e l of t h e loads i n t h e v i c i n i t y of t h e r o o t and t i p r eg io ns .A t an angle of at tack of lIo, h e CN of the panel i s approximately 0.8.It i s p o s s i b l e t h a t s e p ar a ti o n e f f e c t s a t t he roo t and t i p a r e impor tantf o r t h i s configurat ion. I n addit ion, the simple end-plate correct ionused herei n f o r fusela ge i nt er fe re nc es may be approximate.r ega rd the app l i ca t ion o f an analysis such as tha t repor ted by Cr ig ler( r e f . 6) f o r wing-body i nt er fe re nc e a t sonic speeds would improve thepred ic t ion of th e loading i n t h e v i c i n i t y of t h e r o o t.

A t subsonic and supersonic speeds th e ca lcu la ted d i s t r ib ut io n

I n t h i s

The flight-measured loads for the wing panel of the 60 delta-wingJT-102A ai rp la ne a r e considered ne xt.view of the w i n g .

I n f i g ur e 11 i s shown an explodedNote the two fences loca ted i n th e forward port ion of

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t h e w i n g and the elevon surface which i s opera t ive dur ing f l ight .wing has c on ic al camber and a ref lex ed t i p .wing-panel loads, the ef fe ct of the fences and th e ef fe ct s of th e con icalcamber and th e re flexe d t i p m e neglected; however, th e e ff ec t of t h eelevon has been considered.

ThisFor the ca lcula t ion of the

I n f i g u r e s 1 2 and 1 3 are shown the chord loading and the span loadingf o r t h i s a i r p l a n e .from 3 O t o 5O ) the calculations of the chord loadings compare favorablywit h th e measurements. Up-elevon de fl ec ti on i s negat ive . The fa c t th a tthe loading a t the leading edge i s not predicted i s

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NACA Conference on Aircraft Loads 1957