PDF/ocr_dataset_1046.jsonl

{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 20, "total_pages": 33, "image_filename": "19930085977_p20.jpg", "text": "```markdown\nM = .78\nCONFIDENTIAL\nM = .91\nNACA RM L9H22\n\nVertical distance\nabove bump, in.\n6\nM1\n.72\n.73\n4\n2\n0\n8\n10\n12\n14\n16\n18\nM1\n.74\n.75\n.76\n.77\n.78\n.79\n.80\n\n6\nM1\n.84\n.85\n4\n2\n0\n8\n10\n12\n14\n16\n18\nM1\n.86\n.87\n.88\n.89\n.90\n.91\n.92\n.93\n.92\n\nNominal boundary-layer thickness\n\nM = 1.00\nM = 1.17\n\nVertical distance\nabove bump, in.\n6\nM1\n.92\n.93\n4\n2\n0\n8\n10\n12\n14\n16\n18\nStation on bump, in.\nM1\n.94\n.95\n.96\n.97\n.98\n.99\n1.00\n1.01\n1.02\n1.03\n1.04\n\n6\nM1\n1.07\n1.08\n1.09\n1.10\n1.12\n4\n2\n0\n8\n10\n12\n14\n16\n18\nStation on bump, in.\nM1\n1.13\n1.14\n1.15\n1.16\n1.17\n1.18\n1.19\n1.20\n1.21\n\nCONFIDENTIAL\nNACA\n\nFigure 6.— Typical Mach number contours over transonic bump in region of model location.\n19\n```", "timestamp": "2026-07-22T05:45:14.439376+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 22, "total_pages": 42, "image_filename": "19930085938_p22.jpg", "text": "NACA RM No. L9B04\n21\n\n<!-- Image (92, 109, 922, 902) -->\n\nFigure 3.- General arrangement of model 237-7B. (All dimensions are in inches.)", "timestamp": "2026-07-22T05:45:14.735041+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 32, "total_pages": 51, "image_filename": "19930085962_p32.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 31\n\nLift coefficient, $C_L$\n\nElevator deflection, $\\delta_e$, deg\n\n(a) M, 0.20.\n\nFigure 12.—The variation of lift coefficient with elevator deflection for various angles of attack of the tail.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:45:15.546723+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 23, "total_pages": 46, "image_filename": "19930085972_p23.jpg", "text": "NACA RM L9B18\n21\n\nPitching-moment coefficient, $C_m$\n\n$i_t$ (deg)\nNo cutout\nFaired cutout\ntail off\ntail off\n\nLongitudinal-force coefficient, $C_x$\n\nAngle of attack, $\\alpha$, deg\n\nLift coefficient, $C_L$\n\nFigure 6.- Aerodynamic characteristics of a variable-sweep model with and without faired wing cutout. $\\Lambda = 0^\\circ$.", "timestamp": "2026-07-22T05:45:21.342888+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 4, "total_pages": 12, "image_filename": "19930091987_p4.jpg", "text": "# National Advisory Committee for Aeronautics\n\nHeadquarters, 1724 F Street NW., Washington 25, D. C.\n\nCreated by act of Congress approved March 3, 1915, for the supervision and direction of the scientific study of the problems of flight (U. S. Code, title 50, sec. 151). Its membership was increased from 12 to 15 by act approved March 2, 1929, and to 17 by act approved May 25, 1948. The members are appointed by the President, and serve as such without compensation.\n\nJEROME C. HUNSAKER, Sc. D., Cambridge, Mass., Chairman\n\nALEXANDER WETMORE, Sc. D., Secretary, Smithsonian Institute, Vice Chairman\n\nHON. JOHN R. ALISON, Assistant Secretary of Commerce.\nDETLEY W. BRONK, Ph. D., President, Johns Hopkins University.\nKARL T. COMPTON, Ph. D., Chairman, Research and Development Board, Department of Defense.\nEDWARD U. CONDON, Ph. D., Director, National Bureau of Standards.\nJAMES H. DOOLITTLE, Sc. D., Vice President, Shell Union Oil Corp.\nR. M. HAZEN, B. S., Director of Engineering, Allison Division, General Motors Corp.\nWILLIAM LITTLEWOOD, M. E., Vice President, Engineering, American Airlines, Inc.\nTHEODORE C. LONNQUIST, Rear Admiral, United States Navy, Deputy and Assistant Chief of the Bureau of Aeronautics.\n\nDONALD L. PUTT, Major General, United States Air Force, Director of Research and Development, Office of the Chief of Staff, Matériel.\nJOHN D. PRICE, Vice Admiral, United States Navy, Vice Chief of Naval Operations.\nARTHUR E. RAYMOND, Sc. D., Vice President, Engineering, Douglas Aircraft Co., Inc.\nFRANCIS W. REICHELDERFER, Sc. D., Chief, United States Weather Bureau.\nHON. DELOS W. RENTZEL, Administrator of Civil Aeronautics, Department of Commerce.\nHOYT S. VANDENBERG, General, Chief of Staff, United States Air Force.\nTHEODORE P. WRIGHT, Sc. D., Vice President for Research, Cornell University.\n\nHUGH L. DRYDEN, Ph. D., Director\nJOHN W. CROWLEY, Jr., B. S., Associate Director for Research\n\nJOHN F. VICTORY, LL.M., Executive Secretary\nE. H. CHAMBERLIN, Executive Officer\n\nHENRY J. E. REID, D. Eng., Director, Langley Aeronautical Laboratory, Langley Field, Va.\nSMITH J. DEFRANCE, B. S., Director, Ames Aeronautical Laboratory, Moffett Field, Calif.\nEDWARD R. SHARP, Sc. D., Director, Lewis Flight Propulsion Laboratory, Cleveland Airport, Cleveland, Ohio\n\n## TECHNICAL COMMITTEES\n\nAERODYNAMICS\nPOWER PLANTS FOR AIRCRAFT\nAIRCRAFT CONSTRUCTION\n\nOPERATING PROBLEMS\nINDUSTRY CONSULTING\n\nCoordination of Research Needs of Military and Civil Aviation\nPreparation of Research Programs\nAllocation of Problems\nPrevention of Duplication\nConsideration of Inventions\n\nLANGLEY AERONAUTICAL LABORATORY,\nLangley Field, Va.\n\nLEWIS FLIGHT PROPULSION LABORATORY,\nCleveland Airport, Cleveland, Ohio\n\nAMES AERONAUTICAL LABORATORY,\nMoffett Field, Calif.\n\nConduct, under unified control, for all agencies, of scientific research on the fundamental problems of flight\n\nOFFICE OF AERONAUTICAL INTELLIGENCE\nWashington, D. C.\n\nCollection, classification, compilation, and dissemination of scientific and technical information on aeronautics\n\nII", "timestamp": "2026-07-22T05:45:22.482964+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 58, "total_pages": 78, "image_filename": "19930082618_p58.jpg", "text": "```markdown\n.060\n.024\n.020\n.016\n.012\n.008\n.004\nSection drag coefficient, $c_d$\n-.2 -.1 0 .1 .2 .4 .8 1.2\nSection lift coefficient, $c_l$\nR\n$\\circ$ 0.7 $\\times$ 10$^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\nFlagged symbols denote\nstandard roughness\n\n.060\n.024\n.020\n.016\n.012\n.008\n.004\nSection drag coefficient, $c_d$\n+.2 +.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\nR\n$\\nabla$ 3.0 $\\times$ 10$^6$\n$\\square$ 6.0\n$\\diamond$ 9.0\nFlagged symbols denote\nstandard roughness\n\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0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"timestamp": "2026-07-22T05:45:23.162448+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 33, "total_pages": 42, "image_filename": "19930085914_p33.jpg", "text": "32\nNACA RM A9D25\n\nLift coefficient, $C_L$\nAngle of attack, $\\alpha$, deg\n\n$\\nabla$ $R=8 \\times 10^6$\n$\\Delta$ $R=20 \\times 10^6$\n$\\triangleleft$ $R=50 \\times 10^6$\n$\\triangledown$ $R=70 \\times 10^6$\n$\\nabla$ $R=90 \\times 10^6$\n\nfor $R=8 \\times 10^6$\n\n(b) $C_L$ vs $\\alpha$\n\nFigure 12.- Continued.", "timestamp": "2026-07-22T05:45:26.163195+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 10, "total_pages": 36, "image_filename": "19930086097_p10.jpg", "text": "8\nCONFIDENTIAL\nNACA RM A9H11\n\n$$c_d = \\frac{(t/c)^2}{\\sqrt{M_\\infty^2 - 1}} (2-\\eta)^2 - P_b \\frac{h}{c} + c_{df} \\quad (5)$$\n\nHere a minus sign is needed for the term involving $P_b$ since a negative value of $P_b$ corresponds to positive base drag. Equation (5) assumes that the flow does not separate at any point forward of the base. The subscripts 1 and 2 will be used to denote the double-wedge and the blunt-trailing-edge airfoil, respectively. It follows from equation (5) that\n\n$$c_{d_1} = 4 \\frac{(t_1/c)^2}{\\sqrt{M_\\infty^2 - 1}} + c_{df_1} \\quad (6)$$\n\nand\n\n$$c_{d_2} = (2-\\eta)^2 \\frac{(t_2/c)^2}{\\sqrt{M_\\infty^2 - 1}} + c_{df_2} - P_b \\frac{h}{c} \\quad (7)$$\n\nThe fractional difference in drag between the two airfoils is defined as\n\n$$\\frac{\\Delta c_d}{c_{d_1}} \\equiv \\frac{c_{d_2} - c_{d_1}}{c_{d_1}} \\quad (8)$$\n\nWith this definition, negative values of $\\Delta c_d$ correspond to a decrease in profile drag of the blunt-trailing-edge airfoil as compared to the double-wedge airfoil. If the two airfoils (fig. 1) are compared on the basis of equal thickness ratio $t_2=t_1=t$ and, if it is assumed that $c_{df_1} = c_{df_2}$, then substitution of equations (6) and (7) into (8) yields\n\n$$\\frac{(\\Delta c_d)_t}{c_{d_1}} = \\frac{-\\eta + \\frac{\\eta^2}{4} + \\frac{(-P_b \\sqrt{M_\\infty^2 - 1})\\eta}{4(t/c)}}{1 + \\frac{c_{df} \\sqrt{M_\\infty^2 - 1}}{4(t/c)^2}} \\quad (9)$$\n\nThe subscript t indicates that the thickness ratio is the same for both airfoils. If the airfoils are compared on the basis of equal section modulus $t_2^2 = t_1^2 \\frac{2-\\eta}{2-\\eta^3}$ and equation (8) becomes\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:45:33.879772+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 11, "total_pages": 28, "image_filename": "19930086092_p11.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 9\n\nunswept airfoil $\\left(\\frac{\\text{unswept rudder chord}}{\\text{unswept fin chord}} = 0.318\\right)$ was estimated to have a value of about 0.47. It can be shown that, in accordance with the concepts of simple sweep theory (reference 10), $a_{\\delta}$ should decrease in proportion to $\\cos \\Lambda_{c/4}$ if the airfoil is swept back and flap deflection is measured in a plane perpendicular to the hinge line. Applying this correction will give a value of $a_{\\delta}$ of 0.23 for the swept-back vertical tail.\n\nThe rudder effectiveness (equation (2)) computed by using the above values agreed with the experimental value:\n\nTheoretical $C_{n_{\\delta_r}} = -0.0011$\n\nExperimental $C_{n_{\\delta_r}} = -0.0011$\n\nRudder Hinge Moments\n\nAn estimation of the rudder-hinge-moment parameters can be obtained by applying the concepts of simple sweep theory to two-dimensional characteristics. Thus, when the correction for aspect ratio$^4$ is included, the equations for the hinge-moment parameters can be written as follows:\n\n$$C_{h_{\\alpha_t}} = c_{h_{\\alpha}} \\times \\cos \\Lambda_{c/4} \\times \\frac{2A_t}{2A_t+2} \\tag{3}$$\n\n$$C_{h_{\\delta_r}} = c_{h_{\\delta}} \\times \\cos^2 \\Lambda_{c/4} - c_{h_{\\alpha}} \\times a_{\\delta} \\times \\cos^2 \\Lambda_{c/4} \\times \\frac{2A_t}{2A_t+2} \\tag{4}$$\n\nThe two-dimensional values of $c_{h_{\\alpha}}$ and $c_{h_{\\delta}}$ have been estimated to be -0.0070 and -0.0125, respectively. By the use of these values in equations (3) and (4), $C_{h_{\\alpha_t}}$ and $C_{h_{\\delta_r}}$ were computed and can be compared to the experimental results:\n\n| | $C_{h_{\\alpha_t}}$ | $C_{h_{\\delta_r}}$ |\n| :--- | :--- | :--- |\n| Theoretical | -0.0022 | -0.0026 |\n| Experimental | -.0006 | -.0029 |\n\n$^4$Similar to the prediction of $\\Delta C_{n_{\\beta_t}}$, the aspect ratio of the tail, assuming a complete end-plate effect of the fuselage, will be used. This is twice the aspect ratio of the semispan tail.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:45:34.251759+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 4, "total_pages": 34, "image_filename": "19930082472_p4.jpg", "text": "```markdown\n2\nNACA TN No. 1797\n\ncharacteristics of swept wings in the moderate and high lift-\ncoefficient range. These poor characteristics are partially due to\nthe effects of sweep on the potential-flow field and partially due\nto the occurrence of separation over the wing. The potential-flow\neffects are sufficiently well understood and accurate methods of\nprediction are available (reference 1); consequently, they will not\nbe discussed further herein. The effects of flow separation, on the\nother hand, are only superficially known and the mechanism of flow\nseparation is quite obscure. Large increases of drag, sudden and\nlarge fore and aft shifts of the aerodynamic center, decreases in\nlift-curve slope, and eventually, of course, establishment of maximum\nlift coefficient have all been rather generally known to be effects of\nseparation. These effects, furthermore, are manifested at relatively\nlow lift coefficients (sometimes in the lower half of the lift-\ncoefficient range) and thus assume even greater importance than\ncorresponding effects on straight wings (wings with no sweep) where\nseparation is not experienced to any appreciable extent until\nmaximum lift is reached.\n\nTo properly approach the problems involved in alleviating the\neffects of separation, detailed information of the effects of sepa-\nration is necessary. To obtain this information, a large-scale 45°\nswept-forward wing was tested in the Ames 40- by 80-foot wind tunnel.\nThis report presents the results of force and pressure-distribution\nmeasurements, tuft studies, and boundary-layer measurements made to\nenable a detailed correlation between separation phenomena and the\nlongitudinal characteristics of the swept-forward wing.\n\nCOEFFICIENTS AND SYMBOLS\n\nThe data are presented in the form of standard NACA coefficients\nand symbols as defined in the following tabulation:\n\n| | |\n| :--- | :--- |\n| $C_L$ | lift coefficient $\\left( \\frac{\\text{lift}}{qS} \\right)$ |\n| $C_D$ | drag coefficient $\\left( \\frac{\\text{drag}}{qS} \\right)$ |\n| $C_m$ | pitching-moment coefficient computed about the quarter-chord point of the mean aerodynamic chord $\\left( \\frac{\\text{pitching moment}}{qS\\bar{c}} \\right)$ |\n```", "timestamp": "2026-07-22T05:45:35.776262+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 5, "total_pages": 21, "image_filename": "19930082474_p5.jpg", "text": "NACA TN No. 1799\n\nShortly after a student pilot first experiences forward flight, he is impressed with the necessity for having to control the helicopter constantly. The reasons for this situation are not immediately clear. It is a well-known fact that a flapping rotor tilts to the rear if speed is increased; thus the rotor tilt causes the machine to return to the original speed. This condition constitutes stability of the rotor with respect to speed. Wind-tunnel investigations of the subject fuselage (reference 4) have shown it to be unstable with respect to speed, but this instability is evidently outweighed by the rotor stability just discussed, inasmuch as measurements have shown that the stick-position gradient with respect to speed is stable. Furthermore, observation and measurements have indicated that the static stick-force gradient with respect to speed is small, but it has been either unstable, neutral, or stable, the gradient depending upon the pitching moments of the particular blades and upon the bungee configuration; however, the pilot's over-all impression that the helicopter is unstable is not greatly affected by these force gradients. The source of the difficulty, therefore, cannot be either stick-fixed or stick-free instability with speed.\n\nThe somewhat obvious conclusion is that the pilot's impressions are a result of the helicopter's instability with angle of attack. At least two logical sources exist for this instability with angle of attack:\n\n(1) When the helicopter rotor is subjected to an angle-of-attack change in forward flight, for constant rotational speed the advancing blades are subjected to a greater upward accelerating force than the retreating blades because the product of angle-of-attack change and velocity squared is greater on the advancing side than on the retreating side. The resulting flapping motion will then tilt the disk in the direction of the initial change and an unstable moment will result. This effect is a function of the tip-speed ratio and becomes more pronounced at higher speeds.\n\n(2) Wind-tunnel investigations of the fuselage of the subject helicopter have indicated that it is unstable with respect to angle of attack (reference 4).\n\nAlthough airplanes can and do exhibit instability with angle of attack at times, this condition is recognized as unsatisfactory and is generally prevented by keeping the center of gravity sufficiently far forward.\n\nThe effects of the instability with respect to angle of attack on the flight characteristics of the helicopter were subsequently investigated in more detail, first in the low-speed flight range and then at successively higher speeds. In maneuvers in which the stick was abruptly deflected from trim and held, the normal acceleration was found to build up at an increasing rate for a length of time detectable to the pilot. Furthermore, the acceleration and pitching velocity, at least for small stick deflections when the maneuver could be continued for a reasonable time, did not reach a maximum until 3 or 4 seconds had elapsed. The", "timestamp": "2026-07-22T05:45:36.354135+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 35, "total_pages": 36, "image_filename": "19930085912_p35.jpg", "text": "NACA RM No. E9C16\n33\n\n1105\n\n[Figure: Three graphs showing temperature profiles. The vertical axis is labeled \"Distance from duct wall, in.\" and ranges from 0 to 8. The horizontal axis is labeled \"Temperature rise, °F\" and ranges from 0 to 70. The graphs are labeled (a), (b), and (c). A NACA logo is present in the bottom right corner of graph (c).]\n\n(a) Temperature profile measured 13.2 inches downstream of orifice center line.\n\n(b) Temperature profile measured 23.1 inches downstream of orifice center line.\n\n(c) Temperature profile measured 33.1 inches downstream of orifice center line.\n\nFigure 14. - Typical temperature profiles measured 13.2, 23.1, and 33.1 inches downstream of 3/4-inch-diameter orifice. Tunnel velocity, 374 feet per second; plenum-chamber gas temperature, 537° F; plenum-chamber gas pressure, 2900 pounds per square foot absolute.", "timestamp": "2026-07-22T05:45:38.782282+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 33, "total_pages": 51, "image_filename": "19930085962_p33.jpg", "text": "32\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (89, 199, 851, 776) -->\n\nFigure 12. — Continued.\n(b) M, 0.50.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:45:43.243125+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 23, "total_pages": 42, "image_filename": "19930085938_p23.jpg", "text": "22\nNACA RM No. L9B04\n\n[Figure: Hull lines diagram with concentric circles labeled 11 1/2, 18, 25 1/2; side view with stations 7, 9, -2, 0, 1, 3, 5; bottom profile with stations FP, -2, 0, 1, 3, 5, 7, 9, 11 1/2, 18, 25 1/2; NACA logo]\n\nFigure 4.— Hull lines of model 237-7B.", "timestamp": "2026-07-22T05:45:47.643673+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 24, "total_pages": 46, "image_filename": "19930085972_p24.jpg", "text": "22\nNACA RM L9B18\n\nPitching-moment coefficient, $C_m$\n.2\n0\n-.2\n-.4\n-.6\n\n4\n(deg)\nNo cutout\n$\\Delta$ $O_3$\n$\\nabla$ $O_3$ tail off\n$\\diamond$ $O_3$\n$\\square$ $O_3$\n$\\triangle$ tail off\nFaired cutout\n\nLongitudinal-force coefficient, $C_X$\n.3\n.2\n.1\n0\n\nAngle of attack, $\\alpha$, deg\n16\n8\n0\n-8\n\nNACA\n\n-4 0 4 .8 12\nLift coefficient, $C_L$\n\n(a) Basic tail position.\n\nFigure 7.- Aerodynamic characteristics of a variable-sweep model with and without wing cutouts. $\\Lambda = 15^\\circ$.", "timestamp": "2026-07-22T05:45:52.021089+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 59, "total_pages": 78, "image_filename": "19930082618_p59.jpg", "text": "NACA TN 1945\n\nSection lift coefficient, $c_l$\n\nMoment coefficient, $c_{m_{c/4}}$\n\nSection angle of attack, $\\alpha_{sec}$, deg\n\nR\n○ 0.7 × 10⁶\n□ 1.0\n◇ 1.5\n△ 2.0\n▽ 3.0\n▷ 5.0\n◁ 9.0\n\nFlagged symbols denote standard roughness\n\nNACA\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\n\nFigure 13.— Aerodynamic characteristics of the NACA 4415 airfoil section, 24-inch chord.\n\n57", "timestamp": "2026-07-22T05:45:56.141886+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 34, "total_pages": 42, "image_filename": "19930085914_p34.jpg", "text": "NACA RM A9D25\n33\n\nLift coefficient, $C_L$\nPitching-moment coefficient, $C_m$\n\n| | |\n| :--- | :--- |\n| $\\nabla$ | $R = .8 \\times 10^6$ |\n| $\\triangleright$ | $R = 2.0 \\times 10^6$ |\n| $\\triangleleft$ | $R = 5.0 \\times 10^6$ |\n| $\\nabla$ | $R = 7.0 \\times 10^6$ |\n| $\\nabla$ | $R = 9.0 \\times 10^6$ |\n\n[Figure: Graph plotting Lift coefficient ($C_L$) against Pitching-moment coefficient ($C_m$) for various Reynolds numbers ($R$). The x-axis ranges from 0.4 to -0.8. The y-axis ranges from -0.4 to 1.0. There are five distinct curves corresponding to the legend values. A NACA logo is present in the bottom right corner of the plot area.]\n\nfor $R = .8 \\times 10^6$\n\n(c) $C_L$ vs $C_m$\n\nFigure 12.- Concluded.", "timestamp": "2026-07-22T05:46:04.654248+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 12, "total_pages": 28, "image_filename": "19930086092_p12.jpg", "text": "10 CONFIDENTIAL NACA RM A9F14\n\nCONCLUDING REMARKS\n\nAn investigation has been conducted to determine the effects of a vertical tail with the leading edge swept back $63^\\circ$ on the aerodynamic characteristics of a wing-fuselage combination employing a wing with $63^\\circ$ of sweep at the leading edge.\n\nAt angles of attack from $0^\\circ$ to $12^\\circ$, the effectiveness of the vertical tail was maintained to an angle of sideslip of $25^\\circ$ (the highest tested). At an angle of attack of $21^\\circ$, however, effectiveness was maintained only to an angle of sideslip of about $7^\\circ$; beyond $7^\\circ$ the directional stability was irregular.\n\nThe rudder was effective throughout the range of angles of attack and of angles of sideslip tested. At an angle of attack of $21^\\circ$, however, at angles of sideslip greater than about $9^\\circ$, the rudder effectiveness was considerably less than at the lower angles.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nREFERENCES\n\n1. Jones, Robert T.: Estimated Lift-Drag Ratios at Supersonic Speed. NACA TN 1350, 1947.\n\n2. McCormack, Gerald M., and Walling, Walter C.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.— Investigation of a Large-Scale Model at Low Speed. NACA RM A8D02, 1949.\n\n3. Reynolds, Robert M., and Smith, Donald W.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.— Subsonic Mach and Reynolds Number Effects on the Characteristics of the Wing and on the Effectiveness of an Elevon. NACA RM A8D20, 1948.\n\n4. Madden, Robert T.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.— Characteristics at a Mach Number of 1.53 Including Effect of Small Variations of Sweep. NACA RM A8J04, 1949.\n\n5. Madden, Robert T.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.— Investigation at a Mach Number of 1.53 to Determine the Effects of Cambering and Twisting the Wing for a Uniform Load at a Lift Coefficient of 0.25. NACA RM A9C07, 1949.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:46:12.125383+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 36, "total_pages": 36, "image_filename": "19930085912_p36.jpg", "text": "34\nNACA RM No. E9C16\n\n<!-- Image (136, 123, 880, 893) -->\n\nFigure 15. - Comparison of measured and calculated penetration coefficients as function of mixing-distance - diameter ratio.", "timestamp": "2026-07-22T05:46:13.392582+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 4, "total_pages": 21, "image_filename": "19930091993_p4.jpg", "text": "# National Advisory Committee for Aeronautics\n\n*Headquarters, 1724 F Street NW., Washington 25, D. C.*\n\nCreated by act of Congress approved March 3, 1915, for the supervision and direction of the scientific study of the problems of flight (U. S. Code, title 50, sec. 151). Its membership was increased from 12 to 15 by act approved March 2, 1929, and to 17 by act approved May 25, 1948. The members are appointed by the President, and serve as such without compensation.\n\n**JEROME C. HUNSAKER**, Sc. D., Cambridge, Mass., *Chairman*\n\n**ALEXANDER WETMORE**, Sc. D., Secretary, Smithsonian Institute, *Vice Chairman*\n\n**HON. JOHN R. ALISON**, Assistant Secretary of Commerce.\n**DUDLEY W. BROWN**, Ph. D., President, Johns Hopkins University.\n**KARL T. COMPTON**, Ph. D., Chairman, Research and Development Board, Department of Defense.\n**EDWARD U. CONDON**, Ph. D., Director, National Bureau of Standards.\n**JAMES H. DOOLITTLE**, Sc. D., Vice President, Shell Union Oil Corp.\n**R. M. HAZEN**, B. S., Director of Engineering, Allison Division, General Motors Corp.\n**WILLIAM LITTLEWOOD**, M. E., Vice President, Engineering, American Airlines, Inc.\n**THEODORE C. LONNQUIST**, Rear Admiral, United States Navy, Deputy and Assistant Chief of the Bureau of Aeronautics.\n\n**DONALD L. PUTT**, Major General, United States Air Force, Director of Research and Development, Office of the Chief of Staff, Materiel.\n**JOHN D. PRICE**, Vice Admiral, United States Navy, Vice Chief of Naval Operations.\n**ARTHUR E. RAYMOND**, Sc. D., Vice President, Engineering, Douglas Aircraft Co., Inc.\n**FRANCIS W. REICHELDERFER**, Sc. D., Chief, United States Weather Bureau.\n**HON. DELOS W. RENTZEL**, Administrator of Civil Aeronautics, Department of Commerce.\n**HOYT S. VANDENBERG**, General, Chief of Staff, United States Air Force.\n**THEODORE P. WRIGHT**, Sc. D., Vice President for Research, Cornell University.\n\n**HUGH L. DRYDEN**, Ph. D., *Director*\n**JOHN W. CROWLEY**, Jr., B. S., *Associate Director for Research*\n\n**JOHN F. VICTORY**, LL.M., *Executive Secretary*\n**E. H. CHAMBERLIN**, *Executive Officer*\n\n**HENRY J. E. REID**, D. Eng., Director, Langley Aeronautical Laboratory, Langley Field, Va.\n**SMITH J. DEFRANCE**, B. S., Director, Ames Aeronautical Laboratory, Moffett Field, Calif.\n**EDWARD R. SHARP**, Sc. D., Director, Lewis Flight Propulsion Laboratory, Cleveland Airport, Cleveland, Ohio\n\n## TECHNICAL COMMITTEES\n\n| | |\n| :--- | :--- |\n| AERODYNAMICS | OPERATING PROBLEMS |\n| POWER PLANTS FOR AIRCRAFT | INDUSTRY CONSULTING |\n| AIRCRAFT CONSTRUCTION | |\n\n*Coordination of Research Needs of Military and Civil Aviation*\n*Preparation of Research Programs*\n*Allocation of Problems*\n*Prevention of Duplication*\n*Consideration of Inventions*\n\n**LANGLEY AERONAUTICAL LABORATORY**, Langley Field, Va.\n**LEWIS FLIGHT PROPULSION LABORATORY**, Cleveland Airport, Cleveland, Ohio\n**AMES AERONAUTICAL LABORATORY**, Moffett Field, Calif.\n\n*Conduct, under unified control, for all agencies, of scientific research on the fundamental problems of flight*\n\n**OFFICE OF AERONAUTICAL INTELLIGENCE**\nWashington, D. C.\n\n*Collection, classification, compilation, and dissemination of scientific and technical information on aeronautics*\n\nII", "timestamp": "2026-07-22T05:46:14.722519+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 5, "total_pages": 39, "image_filename": "19930093769_p5.jpg", "text": "4 CONFIDENTIAL NACA RM No. E3L10a\n\nchamber were measured with standard instrumentation. The instrumentation for the measurement of turbine-outlet gas temperatures consisted of three rakes of nine unshielded, chromel-alumel thermocouples each installed about $2\\frac{1}{2}$ inches downstream of the turbine outlet.\n\nThe turbine-inlet gas-temperature limits of the engine were obtained from readings of thermocouples installed at the turbine outlet by the engine manufacturer. The thermocouples were calibrated in conformance with the engine manufacturer's specifications.\n\nThe values of combustion efficiency presented are defined as the ratio of the actual enthalpy rise of the gases in passing through the engine divided by the theoretical enthalpy rise, when complete combustion of the fuel is assumed. The enthalpy of the inlet-air and the exhaust gases were determined from the measured compressor-inlet temperature and the turbine-outlet gas temperature with the aid of thermodynamic charts.\n\nDATA CORRECTIONS\n\nAll data were corrected by the following factors:\n\n$\\delta = \\frac{\\text{static pressure at engine discharge, (lb/sq in.)}}{14.7}$\n\n$\\theta = \\frac{\\text{static temperature at engine inlet, (°R)}}{519}$\n\nThe corrected performance parameters are:\n\n| Parameter | Description |\n| :--- | :--- |\n| $F_n/\\delta$ | corrected net thrust, (lb) |\n| $f_n/\\sqrt{\\theta}$ | corrected specific fuel consumption based on net thrust, ((lb/hr)/lb net thrust) |\n| $N/\\sqrt{\\theta}$ | corrected engine speed, (rpm) |\n| $T_5/\\theta$ | corrected tail-pipe gas temperature, (°R) |\n| $W_f/\\delta\\sqrt{\\theta}$ | corrected fuel flow, (lb/sec) |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:46:17.439180+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 11, "total_pages": 36, "image_filename": "19930086097_p11.jpg", "text": "```markdown\nNACA RM A9H11 CONFIDENTIAL 9\n\n$$\n\\frac{(\\Delta c_d)_s}{c_{d_1}} = \\frac{-1 + \\frac{(2-\\eta)^3}{4(2-\\eta^3)} + \\frac{(-P_b \\sqrt{M_\\infty^2-1})}{4(t_1/c)} \\eta \\sqrt{\\frac{2-\\eta}{2-\\eta^3}}}{1 + \\frac{c_{d_f} \\sqrt{M_\\infty^2-1}}{4(t_1/c)^2}} \\quad (10)\n$$\n\nIt can be seen from equations (9) and (10) that the value of $\\eta$ which gives the greatest reduction in profile drag will depend only on the parameter $\\frac{(-P_b)\\sqrt{M_\\infty^2-1}}{t/c}$. Hence, the magnitude of the drag reduction will increase if the product $|P_b|\\sqrt{M_\\infty^2-1}$ is decreased, or if the thickness ratio $t/c$ is increased. The vacuum pressure coefficient $P_{b_v}$ and the product $P_{b_v}\\sqrt{M_\\infty^2-1}$ are shown in figure 2 as a function of the Mach number. The maximum drag reductions possible in comparison to a 10-percent-thick double-wedge airfoil have been calculated for $P_b/P_{b_v}=0$, $1/4$, $1/2$, $3/4$, and $1$, and for $c_{d_f}=0.0028$. This value of the skin-friction coefficient corresponds to laminar flow at a Reynolds number of 1 million. The results are shown in figures 3 and 4. Figure 3 represents the case of equal thickness ratio, and figure 4 the case of equal section modulus. For each curve in these figures the corresponding range of $\\eta$ is indicated. Beyond a Mach number of about 2 the value of $\\eta$ producing the greatest drag reduction increases as the Mach number is increased, and decreases as the base drag is increased. Moreover, for a fixed value of $\\eta$ the greatest drag reductions are obtained at relatively high Mach numbers. This is to be expected since the quantity $|P_{b_v}|\\sqrt{M_\\infty^2-1}$ decreases with increasing Mach number, as shown in figure 2. Thus, the greater drag reduction at high Mach numbers is explained qualitatively by the fact that under the assumed conditions the base drag coefficient decreases more rapidly with increasing Mach number than does the pressure drag coefficient of the airfoil contour forward of the base. It should be remembered that in an actual case the ratio $P_b/P_{b_v}$ probably will change considerably as the Mach number is changed.\n\nFrom the curves in figure 4 it is apparent that, for airfoils of approximately 10-percent-thickness ratio, significant reductions in drag can be achieved, provided the ratio $P_b/P_{b_v}$ is less than about one-half at Mach numbers near 1.5, or less than about three-fourths at Mach numbers near 2.5. At Mach numbers near and beyond about 4, appreciable drag reductions can be achieved for these relatively thick airfoils even if a vacuum exists at the base. The experimental measurements of references 1, 5, and 6 indicate a value of approximately 0.6 for the ratio $P_b/P_{b_v}$. These tests, however, were conducted in the low supersonic Mach number range on wedge airfoils with predominately laminar boundary layers which were relatively thin compared to the thickness of the trailing edge. If the boundary layer were very thick compared to the trailing-edge thickness, then the base drag would have been virtually zero. Hence, by using a moderate amount of bluntness it is\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:46:18.540009+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 24, "total_pages": 42, "image_filename": "19930085938_p24.jpg", "text": "NACA RM NO. L9B04\n\n[Figure: A black-and-white photograph of a model aircraft with twin booms, shown from a three-quarter bottom view. The aircraft has a central fuselage, two long wings, and two tail booms extending rearward with vertical stabilizers. A propeller is mounted on the nose. In the lower right corner of the image, there is a label reading “NACA L-55734.1”.]\n\n(a) Three-quarter bottom view.\n\nFigure 5.— Model with twin booms. Langley tank model 237-7TB.\n\n83", "timestamp": "2026-07-22T05:46:22.233565+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 21, "total_pages": 33, "image_filename": "19930085977_p21.jpg", "text": "CONFIDENTIAL\n\n$1.0 \\times 10^6$\n\nReynolds number, R\n\n.8\n\n.6\n\n.4\n\nMean\n\n.6 .7 .8 .9 1.0 1.1 1.2\n\nMach number, M\n\nCONFIDENTIAL\n\nFigure 7.- Variation of test Reynolds number with Mach number for a model with $0^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section.\n\nNACA RM L9E22", "timestamp": "2026-07-22T05:46:22.475205+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 6, "total_pages": 21, "image_filename": "19930082474_p6.jpg", "text": "4\nNACA TN No. 1799\n\nacceleration and pitching velocity in this type of maneuver apparently\nwould continue to increase for even greater periods of time were it not\nfor the stabilizing influence of the associated speed change.\n\nThe stick forces accompanying these maneuvers are undesirable. The\npilot's impressions are that after transient effects have disappeared,\nthe forces become somewhat unstable, that is, a push in pull-ups or a pull\nin push-downs, the magnitude of the forces depending upon blade charac-\nteristics. Of course stable forces are considered necessary for satis-\nfactory handling qualities.\n\nLongitudinal Oscillations\n\nStick-fixed longitudinal oscillations of the test helicopter were\nstudied to clarify the interaction of the stability with speed and insta-\nbility with angle of attack. Time histories of two attempted stick-fixed\noscillations have been prepared. For these cases the helicopter had a\nset of experimental blades of low solidity that were not production blades.\nLow solidity necessitates higher pitch at the same rotational speed and\nthus stalling was encountered at lower forward speeds for the low-solidity\nblades than for the production blades.\n\nFigure 2 shows an oscillation initiated from steady level flight at\n40 miles per hour by a momentary rearward motion of the stick. The type\nof motion shown resembles the airplane phugoid motion in that changes in\nairspeed and altitude occur, but the important difference is that definite\nchanges in angle of attack take place. The period of the motion is about\n14 seconds, which is long enough that the pilot does not have trouble\ncontrolling the oscillation. The motion about doubles in amplitude in\none cycle. During the third cycle the machine reaches 25° nose up from\nthe trim attitude and shows increments in acceleration, from the 1 g\ncondition, of about 0.4g and -0.3g. This maneuver was terminated when\nthe attitude and the rate of change of attitude, acceleration, and speed\nwere such as to cause the pilot to become apprehensive.\n\nFigure 3 shows an oscillation attempted from steady level flight at\n65 miles per hour. Again the helicopter was disturbed by an intentional\nstick motion, after which the stick was held fixed at the trim position.\nThe helicopter nosed up mildly and then nosed down. The helicopter was\nstill nosing down at an increasing rate, as the acceleration curve\nindicates, at about 9 1/2 seconds after the start of the maneuver, or about\n4 seconds after the 1 g axis was crossed, and the recovery had to be made by\ncontrol application. Immediate response to rearward control was obtained,\nbut as 1 g was reached, the pilot had not only moved the control back to\nthe trim position but was also moving it rapidly forward to check the\nacceleration which was building up at a high rate. The control reached", "timestamp": "2026-07-22T05:46:23.948421+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 5, "total_pages": 12, "image_filename": "19930091987_p5.jpg", "text": "```markdown\n# REPORT 922\n\n## CHARACTERISTICS OF LOW-ASPECT-RATIO WINGS AT SUPERCRITICAL MACH NUMBERS\n\nBy JOHN STACK and W. F. LINDSEY\n\n### SUMMARY\n\nThe separation of the flow over wings precipitated by the compression shock that forms as speeds are increased into the supercritical Mach number range has imposed serious difficulties in the improvement of aircraft performance. These difficulties arise principally as a consequence of the rapid drag rise and the loss of lift that causes serious stability changes when the wing shock-stalls. Favorable relieving effects due to the three-dimensional flow around the tips were obtained and these effects were of such magnitude that it is indicated that low-aspect-ratio wings offer a possible solution of the problems encountered.\n\n### INTRODUCTION\n\nFlight at supercritical Mach numbers has appeared extremely difficult because of rapid drag increase and marked stability and control changes. The change in stability has been found in many instances to be so great as to cause loss of normal control of the aircraft. Serious buffeting of the tail usually accompanies these adverse stability changes.\n\nThe adverse effects are shown in reference 1 and elsewhere to be directly connected with a change of flow over the wing.\n\nThis change of flow is precipitated by the formation of an essentially normal shock, which produces separation of the flow over the wing. The separated flow, as now seems clear, was indicated in reference 2 to be an outstanding contributing cause of the drag rise. The stability change encountered with airplanes is largely a consequence of either or both the change in angle of zero lift of the wing or the change in lift-curve slope of the wing when the separated flow occurs. Elimination of the separated flow could be expected to alleviate to a large extent the difficulties encountered.\n\nElimination of the separated flow could possibly be accomplished by boundary-layer control, though experiments made thus far indicate alleviation, though not elimination, of the separated condition. Other difficulties both aerodynamic and structural are, however, encountered. Because it appears clear that the normal-shock phenomenon produces the separation of the flow, some modification to reduce the shock losses could be expected to contribute markedly toward solution of the present difficulties.\n\nUnpublished results of experimental investigations of the flow around simulated propeller tips in the Langley 11-inch and 24-inch high-speed tunnels showed marked delay and alleviation in the adverse effects at supercritical Mach numbers as compared with results obtained in two-dimensional flows. It was likewise shown that the shocks formed at and near the tips were not normal to the stream. These results are substantiated by the work of reference 3 performed on actual rotating propellers. These results suggest the existence of marked three-dimensional relieving effects at tip sections of wings or propellers. Wings of low aspect ratio could therefore be expected to undergo much less adverse effects at supercritical Mach numbers than wings of present conventional aspect ratios.\n\nConsiderations of the effects of aspect-ratio reduction indicate that other effects may be expected. Thus, the slope of the lift curve is determined by the infinite aspect ratio or section characteristics plus the induced effects. If the induced angle, as with a low-aspect-ratio wing, is large, a given change in section characteristics should produce smaller relative change in lift-curve slope than would occur with a wing of moderate or high aspect ratio for which the induced angle is small. Further, reduction of wing aspect ratio through increasing the downwash angle at the tail reduces the stabilizing effect of the tail until finally a value of aspect ratio is reached for which the geometrical and the downwash angles are approximately equal. When this condition is reached, changes in the flow over the tail as a result of changes in wing characteristics may not produce large changes in stability. The stability will then depend primarily on the wing characteristics.\n\nAs a consequence of the foregoing considerations, experiments were conducted in 1944 in the Langley 24-inch high-speed tunnel to study the characteristics of low-aspect-ratio wings at supercritical Mach numbers. The experiments reported herein consisted of tests of wings of aspect ratios $A$ ranging from $\\infty$ to 2. All the wings were of NACA 0012 section. The tips were cut square, each aspect ratio being obtained by progressively cutting the tips off the original wing. The speed range extended to Mach numbers exceeding 0.9.\n\n### APPARATUS AND TESTS\n\nThe investigation was conducted in the Langley 24-inch high-speed tunnel, which is a nonreturn induction-type tunnel (reference 2). The induction nozzle, located downstream from the test section, induces the air to flow from the atmosphere through the tunnel. The length of air passage from the region of low-velocity air at the entrance section to the test section is small, approximately 4 feet (fig. 1). The absence of a return passage, the short entrance length, and the strong favorable pressure gradient along most of the entrance length provide a very thin boundary layer along the walls of the test section.\n\n585200-49\n1\n```", "timestamp": "2026-07-22T05:46:24.289126+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 34, "total_pages": 51, "image_filename": "19930085962_p34.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 33\n\nLift coefficient, $C_L$\n\nElevator deflection, $\\delta_e$, deg\n\n(c) M, 0.70.\n\nFigure 12. — Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:46:25.819134+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 16, "total_pages": 98, "image_filename": "19930086073_p16.jpg", "text": "14\nNACA RM A9H04\n\nTABLE III.- SUMMARY OF CONFIGURATIONS INVESTIGATED\n\n| Figure | Angle of sideslip (deg) | Deflection of split flaps (deg) | Deflection (deg) | | | Data presented |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | Left aileron | Right aileron | Rudder | |\n| **Wing alone** | | | | | | |\n| 4 | 0.0 | -22.0<br>0<br>21.2<br>44.0 | --- | --- | --- | $C_L$ vs $C_D$<br>$C_m$<br>$C_l$<br>$C_n$<br>$C_Y$ |\n| 5 | 12.1 | -22.0<br>0<br>21.2<br>44.0 | --- | --- | --- | |\n| 6 | 0.0<br>12.1 | --- | 11.7 | -11.3 | --- | |\n| **Wing + body** | | | | | | |\n| 7 | 0.0 | -20.7<br>-10.8<br>0<br>20.4<br>45.4 | --- | --- | --- | $C_L$ vs $C_D$<br>$C_m$ |\n| 8 | 0.0<br>6.0<br>12.0<br>15.9 | --- | --- | --- | --- | $C_L$ vs $C_D$<br>$C_m$<br>$C_l$<br>$C_n$<br>$C_Y$ |\n| 9 | 0.0<br>6.0<br>12.0<br>15.9 | -20.7 | --- | --- | --- | |\n| 10 | 0.0<br>6.0<br>12.0<br>15.9 | 20.4 | --- | --- | --- | |\n| 11 | 0.0<br>6.0<br>12.0<br>15.9 | 45.4 | --- | --- | --- | |\n| 12 | 0.0 | --- | 10.8<br>0<br>0.0 | 0<br>0<br>-10.8 | --- | |\n| 13 | 0.0 | --- | 10.8<br>0 | -10.8<br>0 | --- | |\n| 14 | 12.0 | --- | -10.8<br>0<br>10.8 | --- | --- | $C_L$ vs $C_D$<br>$C_m$<br>$C_l$<br>$C_n$<br>$C_Y$ |\n| 15 | 12.0 | --- | --- | -10.8<br>0<br>10.8 | --- | |\n| 16 | 12.0 | --- | 10.8<br>0 | -10.8<br>0 | --- | |\n| **Wing + body + vertical tail** | | | | | | |\n| 17 | 0.0<br>6.0<br>12.0<br>15.9 | --- | --- | --- | 0 | $C_L$ vs $C_D$<br>$C_m$<br>$C_l$<br>$C_n$<br>$C_Y$ |\n| 18 | 0.0<br>6.0<br>12.0<br>15.9 | --- | --- | --- | 10 | |\n| 19 | 0.0<br>12.0 | -20.7 | --- | --- | 10 | |\n\nNACA", "timestamp": "2026-07-22T05:46:26.122148+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 5, "total_pages": 34, "image_filename": "19930082472_p5.jpg", "text": "NACA TN No. 1797\n3\n\n$c_l$ section lift coefficient $\\left( \\frac{1}{c} \\int_{0}^{c} P dx \\cos \\alpha - \\frac{1}{c} \\int_{0}^{t} P dz \\sin \\alpha \\right)$\n\nc.p. center of pressure of section normal force, measured in percent chord aft of the leading edge\n\nP pressure coefficient $\\left( \\frac{p_l - p}{q} \\right)$\n\np free-stream static pressure, pounds per square foot\n\n$p_l$ local static pressure, pounds per square foot\n\nq free-stream dynamic pressure, pounds per square foot\n\n$\\alpha$ angle of attack, degrees\n\nS wing area, square feet\n\nb wing span, feet\n\n$\\bar{c}$ mean aerodynamic chord $\\left( \\frac{\\int_{0}^{b/2} c^2 dy}{\\int_{0}^{b/2} c \\ dy} \\right)$\n\nc local chord, feet\n\nt maximum thickness of local section, feet\n\ny spanwise coordinate perpendicular to plane of symmetry, feet\n\nx chordwise coordinate parallel to plane of symmetry, feet\n\nz vertical coordinate to airfoil contour perpendicular to chord line, feet\n\n$\\Lambda$ angle of sweep to the quarter-chord line, degrees\n\n$\\delta$ boundary-layer thickness, feet\n\nMODEL\n\nThe geometric characteristics and dimensions of the swept-forward wing are shown in figure 1. The wing had $45^\\circ$ of sweepforward of the quarter-chord line, an aspect ratio of 3.55, a taper ratio of", "timestamp": "2026-07-22T05:46:26.288276+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 15, "total_pages": 50, "image_filename": "19930086076_p15.jpg", "text": "NACA RM E9F09\n13\n\n[Figure: A photograph of a metal fuel injector assembly. An arrow labeled \"Air flow\" points to the left. A scale bar labeled \"INCHES\" is visible. A stamp in the bottom right corner reads \"NACA C-23186 3-23-49\".]\n\n(a) Simple-orifice fuel injector.\nFigure 2. - Fuel injectors for 4- by 8-inch combustor.", "timestamp": "2026-07-22T05:46:26.874779+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 25, "total_pages": 46, "image_filename": "19930085972_p25.jpg", "text": "NACA RM L9B18\n23\n\nPitching-moment coefficient, $C_m$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\nLongitudinal-force coefficient, $C_x$\n\n$i_t$ (deg)\nUnfaired cutout\ntail off\nLarge unfaired cutout\ntail off\n\n(a) Concluded.\nFigure 7.- Continued.", "timestamp": "2026-07-22T05:46:40.699949+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 25, "total_pages": 42, "image_filename": "19930085938_p25.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:46:45.636078+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 1, "total_pages": 78, "image_filename": "19930082483_p1.jpg", "text": "Y3.N21/5:6/1807\nGOVT. DOC.\nNACA TN No. 1807\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\n\nNo. 1807\n\nEFFECTS OF PARTIAL ADMISSION ON PERFORMANCE\nOF A GAS TURBINE\n\nBy Robert C. Kohl, Howard Z. Herzig\nand Warren J. Whitney\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\n[Figure: NACA logo]\n\nWashington\nFebruary 1949\n\nBUSINESS, SCIENCE\n& TECHNOLOGY DEPT.\nFEB 10 1949", "timestamp": "2026-07-22T05:46:46.049709+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 13, "total_pages": 28, "image_filename": "19930086092_p13.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 11\n\n6. Hopkins, Edward J.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.– Effects of Split Flaps, Elevons, and Leading-Edge Devices at Low Speed. NACA RM A9C21, 1949.\n\n7. Jones, Lloyd J.. and Demele, Fred A.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.– Characteristics Throughout the Subsonic Speed Range with the Wing Cambered and Twisted for a Uniform Load at a Lift Coefficient of 0.25. NACA RM A9D25, 1949.\n\n8. Mas, Newton: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.– Characteristics for Symmetrical Wing Sections at High Subsonic and Moderate Supersonic Mach Numbers. NACA RM A9E09, 1949.\n\n9. DeYoung, John: Theoretical Additional Span Loading Characteristics of Wings With Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA TN 1491, 1947.\n\n10. Betz, A.: Applied Airfoil Theory. Unsymmetrical and Non-Steady Types of Motion. Vol. IV of Aerodynamic Theory, div. J, ch. IV, sec. 4, W. F. Durand, ed., Julius Springer (Berlin), 1935, pp. 94–107.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:46:47.269191+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 35, "total_pages": 51, "image_filename": "19930085962_p35.jpg", "text": "34\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (94, 126, 854, 879) -->\n\nFigure 12. — Continued.\n(d) M, 0.80.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:46:51.821514+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 12, "total_pages": 36, "image_filename": "19930086097_p12.jpg", "text": "```markdown\n10 CONFIDENTIAL NACA RM A9H11\n\nto be expected that the ratio $P_b/P_{b_V}$ can be held considerably below 0.6. It is concluded, therefore, that the proper use of a blunt trailing edge on 10-percent-thick airfoils will enable drag reductions to be achieved in the low range as well as in the high range of supersonic Mach numbers.\n\nFor thinner airfoils the preceding conclusion is still valid, although the percentage drag reductions are less for a given base pressure coefficient. Thus, for airfoils of 5-percent-thickness ratio, $P_b/P_{b_V}$ would have to be a little less than one-fourth in order to achieve the same percentage drag reduction that is possible for 10-percent-thick airfoils with a $P_b/P_{b_V}$ of one-half. Experiments are needed to answer the question of whether or not this ratio can be held to values of approximately one-fourth at the Reynolds numbers encountered in practical applications.\n\nIn addition to illustrating the conditions under which the use of bluntness will decrease the profile drag, equations (9) and (10) also illustrate the conditions under which the improper use of bluntness may lead to an increase in drag. For example, with $P_b/P_{b_V} = 1/2$, a wedge airfoil of 10-percent-thickness ratio at a Mach number of 1.5 will have approximately 13-percent higher drag than a double-wedge airfoil of the same thickness ratio. In general, airfoils with excessive thickness at the trailing edge will have considerably higher drag at low supersonic Mach numbers than airfoils with sharp trailing edges.\n\nLift and Pitching Moment\n\nApart from the effect of trailing-edge bluntness on profile drag, the accompanying effect on the lift characteristics at supersonic velocities can also be of practical importance. The conventional considerations of two-dimensional perturbation theory applied to sharp-trailing-edge profiles show that the lift-curve slope is independent of the airfoil shape even if second-order terms are considered. This statement, however, must be modified in order to apply to blunt-trailing-edge airfoils. Although the foregoing analysis of drag characteristics was restricted to symmetrical profiles of straight-line segments, the subsequent analysis of lift and pitching-moment characteristics is not restricted to any particular airfoil contour.\n\nTo the second order in angular deflections the pressure coefficient, according to the Busemann second-order airfoil theory, is\n\n$$P = C_1\\theta + C_2\\theta^2 \\quad (11)$$\n\nwhere $\\theta$ and $C_1$ are as previously defined and\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:46:55.088276+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 16, "total_pages": 50, "image_filename": "19930086076_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:46:55.284749+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 5, "total_pages": 29, "image_filename": "19930093789_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM No. E8I21\n\nDesign values\nEntrance absolute total temperature, $T_1'$, $^\\circ$R . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T05:46:55.435460+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 6, "total_pages": 39, "image_filename": "19930093769_p6.jpg", "text": "NACA RM No. E8L10a CONFIDENTIAL 5\n\nPROCEDURE AND RESULTS\n\nAltitude Performance with AN-F-58 Fuel\n\nThe altitude performance of engine A using AN-F-58 fuel (NACA fuel numbers 48-206 and 48-210) was investigated and found satisfactory over a range of altitudes from 5000 to 50,000 feet at flight Mach numbers from 0.25 to 0.85. Engine-inlet pressures and temperatures were maintained at values corresponding to NACA standard air at simulated flight conditions and standard NACA altitude pressure was maintained at the engine outlet. The recorded and corrected values of engine net thrust, specific fuel consumption, tail-pipe gas temperature, and fuel flow for the range of engine speeds and flight conditions investigated, as well as the fuel used, are presented in table II. Although an engine failure during the investigation prevented the determination of comparative performance data on the same engine with gasoline, these data illustrate the range of conditions over which the engine satisfactorily operated with AN-F-58 fuel. At altitudes of 45,000 and 50,000 feet, the maximum operable engine speed was limited to the values indicated in table II by excessive tail-pipe gas temperatures, which were believed to be partly caused by burning of fuel through the turbine.\n\nEngine starts during this investigation of altitude performance of the engine were unusually hot and accompanied by longer time intervals between the introduction of the fuel and the start of combustion than usually experienced with gasoline. This difficulty was believed to be caused by operational technique rather than the characteristics of the fuel and was not again encountered in the investigation.\n\nComparison of Altitude Performance\n\nwith AN-F-58 Fuel and Gasoline\n\nThe performance with AN-F-58 fuel and with gasoline, for comparative purposes, was determined in engine B. This investigation was conducted with AN-F-58 fuel (NACA fuel number 48-210) for a range of engine speeds from 9000 to 12,500 rpm and at the various simulated flight conditions listed in the following table:\n\n| Altitude (ft.) | Flight Mach number |\n|----------------|--------------------|\n| 5,000          | 0                  |\n| 20,000         | .60, .85, 1.00     |\n| 35,000         | 1.00               |\n| 50,000         | .85                |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:46:55.485688+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 5, "total_pages": 21, "image_filename": "19930091993_p5.jpg", "text": "```markdown\n# REPORT 928\n\n## ANALYSIS OF PERFORMANCE OF JET ENGINE FROM CHARACTERISTICS OF COMPONENTS\n## II—INTERACTION OF COMPONENTS AS DETERMINED FROM ENGINE OPERATION\n\nBY ARTHUR W. GOLDSTEIN, SUMNER ALPERT, WILLIAM BEEDE, and KARL KOVACH\n\n### SUMMARY\n\nIn order to understand the operation and the interaction of jet-engine components during engine operation and to determine how component characteristics may be used to compute engine performance, a method to analyze and to estimate performance of such engines was devised and applied to the study of the characteristics of a research turbojet engine built for this investigation. An attempt was made to correlate turbine performance obtained from engine experiments with that obtained by the simpler procedure of separately calibrating the turbine with cold air as a driving fluid in order to investigate the applicability of component calibration. After correction for blade-tip leakage, the turbine-characteristic curves of weight flow and total-pressure ratio checked with the results from cold-air component calibration. Some discrepancies in efficiency were noted between the two sets of experiments. Despite such errors, turbine-compressor interaction may be accurately determined but some error in exhaust pressure may be involved.\n\nFrom analysis of the component calibrations, predictions that investigation of the engine without modifications would not cover an adequate range of turbine performance were verified by the engine performance. The range of turbine operation was extended by study of the engine with modifications to the compressor.\n\nThe system of analysis was also applied to prediction of the engine and component performance with assumed modifications of the burner and bearing characteristics, to prediction of component and engine operation during engine acceleration, and to estimates of the performance of the engine and the components when the exhaust gas was used to drive a power turbine.\n\n### INTRODUCTION\n\nIn order to understand better the performance characteristics of jet-propulsion engines, a research unit was designed and built at the NACA Cleveland laboratory under the direction of Eastman N. Jacobs. The operation and the interaction of the components of this unit under all possible engine operating conditions were studied. Preliminary studies of the compressor and turbine performance were first made and reported in references 1 to 4. An analysis of the component data (reference 4) showed that the constraint imposed on the turbine by operation with the compressor in the engine would limit obtainable data to a very small portion of the good efficiency range of the turbine. In fact, the range was expected to be so small that in view of the relatively large experimental errors expected in the engine data, no reliable indication of the location of the region for best turbine operation was anticipated. In order to circumvent this difficulty, two sets of engine experiments were made. After the initial investigations of the engine were completed, the compressor stator blades were reset for a lower air capacity and the engine was rerun. These two sets of data were sufficient to give a reasonably accurate picture of the turbine performance over the most important range.\n\nFrom the characteristics of the components, the operation of the engine components under all possible modes of engine operation were investigated during 1946; the study included the effect of bearing and burner modifications on engine performance, operation of the engine and components during acceleration, and operation of the gas turbine and components when a power turbine is driven by the jet-engine exhaust. The basis of the analysis, which was developed for this purpose, is a simultaneous graphical solution of the equations relating compressor and turbine speed, power and gas flow, and the curves expressing the performance characteristics of compressor, bearings, burner, and turbine.\n\n### DESCRIPTION OF ENGINE\n\nThe compressor component of the engine (fig. 1) was an eight-stage compressor described in reference 1 but modified for installation in the engine by replacing the discharge\n\n[Figure: Compressor - Diffuser - Combustion chamber - Turbine - C-9469 8-1-49]\nFIGURE 1.—NACA research jet engine.\n\nscroll collector with an axial-flow diffuser containing a row of straightening blades. The rear-compressor journal and tapered-land thrust bearings and the turbine journal bearing were contained in a cone supported by struts from the outside wall of the diffuser (fig. 2). An outer shell bolted to the compressor casing supported the turbine casing. Inside this shell, the outer wall of the annular combustion chamber was bolted to the outlet end of the diffuser and was connected to the turbine casing with an air-tight expansion joint. The diffuser casing and the combustion chamber with part of the outer wall and the supporting shell removed is shown in figure 3; some of the 27 individual burners are\n\n1\n```", "timestamp": "2026-07-22T05:46:56.139471+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 17, "total_pages": 98, "image_filename": "19930086073_p17.jpg", "text": "NACA RM A9H04\n15\n\n[Figure: Diagram of an aircraft showing force and moment coefficients. Axes X, Y, Z are shown. Forces $C_D$, $C_Y$, $C_Z$ and moments $C_l$, $C_n$, $C_m$ are indicated with arrows. Angles $\\alpha$ and $\\beta$ are shown at the nose. Velocity vector V is shown.]\n\nFigure 1.- Sign convention for force and moment coefficients. All forces, moments, angles, control-surface deflections, and axes are shown as positive.\n\nNACA", "timestamp": "2026-07-22T05:46:58.312344+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 6, "total_pages": 34, "image_filename": "19930082472_p6.jpg", "text": "4\nNACA TN No. 1797\n\n0.5, no twist, and no dihedral. The wing had an NACA 64A112 section (table I) perpendicular to the quarter-chord line. A photograph of the wing mounted in the tunnel is shown in figure 2.\n\nThe pressure orifices were positioned over the upper and lower surfaces of streamwise sections which were located at eight spanwise positions varying from 20.9- to 96.2-percent semispan. The spanwise and chordwise positions of the orifices are listed in table II.\n\nSurveys of the boundary layer were made by means of rakes attached to the wing surface. A typical rake used is shown in figure 3. The rakes consisted of a bank of total-head tubes parallel to the axis of the rake and two banks placed at an angle of 63° to the axis. The tubes parallel to the axis of the rake were used to measure the total-head variation through the boundary layer. The tubes at an angle to the axis were used to determine the variation of flow angle through the boundary layer.\n\nTESTS AND RESULTS\n\nForce and pressure-distribution measurements, tuft studies, and boundary-layer measurements were made through an angle-of-attack range at zero sideslip. Data were obtained at Reynolds numbers from $6.9 \\times 10^6$ to $14 \\times 10^6$ (based on the mean aerodynamic chord length of 10.41 ft). However, since these data indicated no appreciable Reynolds number effect, particularly within the purpose of this report, only the data obtained at a Reynolds number of $9.7 \\times 10^6$ (a tunnel speed of approximately 100 mph) are presented.\n\nStandard tunnel-wall corrections for a straight wing of the same area and span as the swept-forward wing have been applied to angle-of-attack and drag-coefficient data. This procedure was followed, since a brief analysis indicated that tunnel-wall corrections were approximately the same for straight and swept wings of the size under consideration. The corrections applied are as follows:\n\n$$\n\\Delta\\alpha = 0.74 \\text{ C}_L\n$$\n\n$$\n\\Delta\\text{C}_D = 0.013 \\text{ C}_L^2\n$$\n\nThe data were corrected for drag tares. Pitching-moment tares were not applied, since they were not known with sufficient accuracy to warrant application. The pitching-moment tares are felt to be quite small, however, and should not appreciably affect the results.", "timestamp": "2026-07-22T05:47:01.410564+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 7, "total_pages": 21, "image_filename": "19930082474_p7.jpg", "text": "NACA TN No. 1799\n\nthe forward stop about 2 seconds before the acceleration reached its peak of $1.7g$. The time history does not tell the whole story, however, for during this maneuver, as the stick approached the forward stop, the collective pitch was reduced to about $6^\\circ$ to reduce the acceleration and the associated blade stalling. The rotational speed went above the placard limit. In addition, the horizon disappeared from the pilot's view; thus a very high attitude was indicated, when the field of view from this aircraft is considered. A roll at the top of the maneuver as in a wing-over was necessary for recovery. The maneuver just described could be entered inadvertently should the pilot permit his attention to be briefly diverted. It is obviously extremely hazardous and the consequences should not be underestimated.\n\nComparison of these two time histories indicates the marked influence that speed has on the instability with angle of attack and therefore on the difficulty of controlling the aircraft. In order to bring out this trend with speed more clearly, additional oscillations, including some made in mildly gusty air, were made to provide more points in the speed range.\n\nIn order to obtain greater generality, different rotor blades were used on the same helicopter and a later model helicopter of basically similar design was also utilized. In all cases the time required with controls fixed to reach a dangerous flight condition following the first definite nose-down motion was noted. For the cases where relatively complete instrumentation was used, the increment in normal acceleration from the $1\\ g$ condition that had been reached, at the flight condition considered dangerous, was usually about $\\frac{1}{4}g$, regardless of forward speed. The acceleration increment appears to be a much better criterion for the flight condition at which recovery must be started than is the more commonly discussed attitude angle. The value of $\\frac{1}{4}g$ mentioned for this increment probably corresponds to the particular helicopter under test and may be expected to vary with size and other characteristics of the helicopter. The results of the measurements that have been made are summarized in figure 4. In this figure the increment in acceleration per unit time is shown plotted against airspeed. The ordinate values were obtained by taking the reciprocal of the values of time to reach a dangerous flight condition, which, as has been pointed out, corresponded to a reasonably fixed acceleration increment of about $\\frac{1}{4}g$. Thus, the higher the value shown, the earlier a dangerous condition would be reached and, hence, the more frequently the pilot has to apply control to maintain steady flight. In other words, if corrective control is applied at given intervals, then the higher the value shown, the greater the amount of corrective control required.\n\nFrom about 40 miles per hour to 50 or 60 miles per hour the values shown are relatively low. In this region the helicopter can actually be made stable by relatively simple means, and in any event it requires relatively little attention from the pilot. At the higher speeds the attentiveness required of the pilot rises rapidly. In like manner, many methods of improving the stability characteristics which could readily be", "timestamp": "2026-07-22T05:47:02.482559+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 26, "total_pages": 46, "image_filename": "19930085972_p26.jpg", "text": "24\nNACA RM L9B18\n\nPitching-moment coefficient, $C_m$\n.2\n0\n-.2\n-.4\n-.6\n\n(deg)\n$\\nabla$ 0\n$\\nabla$ 3\n$\\diamond$ tail off\n$\\square$ 0\n$\\square$ 3\n$\\triangle$ tail off\nNo cutout\nFaired cutout\n\n-.4\n-.3\n-.2\n-.1\n0\nLongitudinal-force coefficient, $C_x$\n\nAngle of attack, $\\alpha$, deg\n16\n8\n0\n-8\n\n-4\n0\n4\n8\n12\nLift coefficient, $C_L$\n\nNACA\n\n(b) Alternate tail position.\nFigure 7.- Continued.", "timestamp": "2026-07-22T05:47:08.189503+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 35, "total_pages": 42, "image_filename": "19930085914_p35.jpg", "text": "34\n\nLift-drag ratio, $L_D$\n\n16\n12\n8\n4\n0\n\n$R=.8 \\times 10^6$ $R=2.0 \\times 10^6$ $R=5.0 \\times 10^6$ $R=7.0 \\times 10^6$ $R=9.0 \\times 10^6$\n\n0 .2 .4 .6 .8 for $R=.8 \\times 10^6$\n\nLift coefficient, $C_L$\n\n[Figure: Graph showing multiple curves with different symbols (circles, squares, diamonds, triangles, inverted triangles) representing different Reynolds numbers. A NACA logo is present in the bottom right corner of the graph area.]\n\nFigure 13.— The variation of lift-drag ratio with lift coefficient of the wing-fuselage combination at several Reynolds numbers at a Mach number of 0.20.\n\nNACA RM A9D25", "timestamp": "2026-07-22T05:47:08.712239+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 2, "total_pages": 78, "image_filename": "19930082483_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:47:14.507061+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 17, "total_pages": 50, "image_filename": "19930086076_p17.jpg", "text": "NACA RM E9F09\n15\n\n[Figure: A photograph of a metal mechanical component, a hollow-spray-cone fuel injector. An arrow on the left is labeled \"Air flow\". A scale bar in the bottom left is labeled \"INCHES\". A NACA logo with the text \"C-22960\" and \"2-21-49\" is in the bottom right corner.]\n\n(b) Hollow-spray-cone fuel injector\nFigure 2. - Concluded. Fuel injectors for 4- by 8-inch combustor.", "timestamp": "2026-07-22T05:47:24.772116+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 36, "total_pages": 51, "image_filename": "19930085962_p36.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 35\n\n[Figure: Graph plotting Lift coefficient, $C_L$ vs. Elevator deflection, $\\delta_e$, deg. The y-axis ranges from -0.8 to 1.4. The x-axis ranges from -4 to 32. Multiple curves are plotted for angles of attack ($\\alpha_{approx.}$) ranging from $8^\\circ$ to $-8^\\circ$. A NACA logo is present in the bottom right corner of the graph area.]\n\nFigure 12. — Continued.\n(e) M, 0.85.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:47:28.701453+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 13, "total_pages": 36, "image_filename": "19930086097_p13.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 11\n\n$$\nC_2 = \\frac{(\\gamma+1)M_\\infty^4 - 4(M_\\infty^2-1)}{2(M_\\infty^2-1)^2}\n\\tag{12}\n$$\n\nThe coordinates used in the calculations are shown in figure 5. Integrating the pressure coefficient over the airfoil contour yields the lift coefficient\n\n$$\nc_l = \\int_0^1 (P_l - P_u) \\, d\\left(\\frac{x}{c}\\right)\n\\tag{13}\n$$\n\nIn this equation the small negative contribution of the base pressure to the lifting forces has been neglected, since it amounts to less than about 1 percent for $t/c=0.05$ and less than about 2 percent for $t/c=0.10$. The variable $x$ is measured along the chord line which is arbitrarily defined as passing through the leading edge and bisecting the base at the trailing edge. On the upper surface the local angle of inclination is $\\theta_u = (dy/dx)_u - \\alpha$, and on the lower surface it is $\\theta_l = -(dy/dx)_l + \\alpha$. Substituting equation (11) into (13) and carrying out the detailed integration yields\n\n$$\n\\begin{aligned}\nc_l &= 2C_1\\alpha + C_2 \\int_0^1 \\left\\{ \\left(\\frac{dy}{dx}\\right)_l^2 - \\left(\\frac{dy}{dx}\\right)_u^2 - 2\\alpha \\left[ \\left(\\frac{dy}{dx}\\right)_l - \\left(\\frac{dy}{dx}\\right)_u \\right] \\right\\} d\\left(\\frac{x}{c}\\right) \\\\\n&= 2C_1\\alpha + C_2 \\int_0^1 \\left[ \\left(\\frac{dy}{dx}\\right)_l^2 - \\left(\\frac{dy}{dx}\\right)_u^2 \\right] d\\left(\\frac{x}{c}\\right) + 2 \\frac{C_2\\alpha}{c} \\int_{x=0}^{x=c} (dy_u - dy_l) \\\\\n&= 2C_1\\alpha + 2C_2\\alpha \\frac{h}{c} + C_2 \\int_0^1 \\left[ \\left(\\frac{dy}{dx}\\right)_l^2 - \\left(\\frac{dy}{dx}\\right)_u^2 \\right] d\\left(\\frac{x}{c}\\right)\n\\end{aligned}\n\\tag{14}\n\\tag{15}\n$$\n\nSince the last term in this equation is independent of angle of attack,\n\n$$\n\\frac{dc_l}{d\\alpha} = 2C_1 \\left( 1 + \\frac{C_2}{C_1} \\frac{h}{c} \\right)\n\\tag{16}\n$$\n\nThe effect of trailing-edge bluntness for any airfoil contour, therefore, is simply to increase the section lift-curve slope by the factor $1+(C_2/C_1)(h/c)$. This expression was, in fact, given many years ago by Busemann (reference 1) for the case of a wedge airfoil. The above analysis simply brings to light a fact which is implicit in Busemann's equations, though not explicitly stated; namely, equation (16) expressing\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:47:28.853712+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 6, "total_pages": 29, "image_filename": "19930093789_p6.jpg", "text": "```markdown\nNACA RM No. E8I21 CONFIDENTIAL 5\n\ncontained 28 stator blades. The stator-blade height was about\n1/8 inch less than the blade height at the rotor entrance, or 0.92\nof the rotor-blade height. This stator was cast in one piece.\n\nA second stator (fig. 6), which was of welded construction,\nhad a $0^\\circ$ cone angle, and contained 36 stator blades. The stator-\nblade height was made equal to the blade height at the rotor\nentrance. If the assumptions used in designing the rotor blades\nare applied to this stator, the stator meets the design conditions\nat the rotor entrance.\n\nFour differences existed in the two stators; namely, the method\nof construction, the number of stator blades, the cone angle, and\nthe stator-blade height. The internal surfaces of the $70^\\circ$-cone-\nangle stator were hand-finished so that the surfaces of both stators\nwere of approximately equal roughness. Because the solidity is high\nand separation is unlikely to occur in either stator, the number of\nstator blades should have little effect. The cone angle and the\nstator-blade height are the significant differences. With each\nstator, the minimum axial clearance between the stator blades and\nthe rotor blades was between 0.120 and 0.140 inch.\n\n[Handwritten annotation: if thickness of outlet edge is the same]\n\nShrouds\n\nThe two shrouds used in this investigation are shown in fig-\nure 7. The labyrinth, no-leakage, stationary shroud (fig. 7(a))\nconsists of a circular ring with two circumferential grooves on the\ninner surface. The radial clearance between the labyrinth station-\nary shroud and the cylindrical rotating shroud formed by the blade\ncaps is 0.025 inch. Leakage between the labyrinth shroud and the\nblade caps was prevented by introducing air into the space down-\nstream of the labyrinth shroud. The amount of labyrinth-sealing\nair was so adjusted that the pressures in the two circumferential\ngrooves were equalized; when the pressures were equal, it was assumed\nthat no air leaked past the shroud.\n\nThe cylindrical stationary shroud (fig. 7(b)) is the same width\nas the blade caps. The radial clearance between the blade caps\nand the shroud was 0.025 inch, or 0.016 of the blade height.\n\nConfigurations\n\nThe three turbine configurations investigated are shown in\nfigure 8. Configuration 1 includes the $70^\\circ$-cone-angle stator and\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:47:32.306660+00:00"}