PDF/ocr_dataset_104.jsonl

{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 18, "total_pages": 21, "image_filename": "19930094534_p18.jpg", "text": "N.A.C.A. Technical Memorandum No. 882\n\nFigs. 5,10,11,12,13,27,28,29\n\nFigure 5.- Test cylinder.\n\nFigure 12.- Tubular duralumin push rod 35 x 1 with self-aligning bearing head, old and new design.\n\nFigure 13.- Flanged duralumin tubes.\n\nFigure 28.- Countersinking cutter.\n\nFigure 29.- Hand creasing tool.\n\nFig. 10 M 14X15\n\nFig. 11 M 12X15\n\nFigures 10,11.- Failures of a series of test tubes with rolled thread.\n\nFigure 27.- Multi spindle drill inserts for hand drill machine.", "timestamp": "2026-07-19T18:28:11.436672+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 34, "total_pages": 43, "image_filename": "19930094538_p34.jpg", "text": "N. A. C. A. Technical Memorandum No. 878 Figs. 2,3,4\n\n[Figure: Cylinder No. II after buckling of the skin.]\n\nFigure 2.- Cylinder No. II after buckling of the skin.\n\n[Figure: Cylinder No. III after buckling of the skin.]\n\nFigure 3.- Cylinder No. III after buckling of the skin.\n\n[Figure: Cylinder No. IV after buckling of the skin.]\n\nFigure 4.- Cylinder No. IV after buckling of the skin.", "timestamp": "2026-07-19T18:28:11.939626+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 11, "total_pages": 18, "image_filename": "19930094551_p11.jpg", "text": "10 N.A.C.A. Technical Memorandum No. 865\n\nfacilitates the finding of the coefficients relevant to the different A.\n\nThe right-hand scale in figure 9, containing the other three terms of (12), is in two parts. The part above the traced abscissa comprises the sum of the two terms to the left of the equals sign\n\n$$\n\\sin \\alpha_s \\cos^2 \\alpha_s + \\frac{C_2}{C_3} \\sin^2 \\alpha_s\n$$\n\nWhile the cited upper part of the ordinate depends, apart from $\\alpha_s$, only on the ratio $\\frac{C_2}{C_3} = \\frac{C}{C_1}$ i.e., on the ratio of the coefficients of lift and drag and consequently assumes other values simply as form factor for different tubings, the lower part carries the function $\\frac{2B}{C_3} = f(B)$. The different straights correspond to different $C_3$, i.e., different impact pressures for chosen thickness and length of the tubing elements.\n\nFigures 10 and 11 facilitate the determination of forces $A_n$ and $E_n$ in (8) and (9). They render the tracing of $C_2 \\sin^2 \\alpha_s$ and $C_3 \\sin^2 \\alpha_s \\cos \\alpha_s$ for different $\\alpha_s$ and C values, possible.\n\nFirst, the constants C and $C_1$ are ascertained from the wind-tunnel curves (equations (4) and (5)). Then, the choice of chord length, tubing diameter, and impact pressure is followed by the calculation of:\n\n$$\nC_2 = \\Delta \\ l \\ d \\ q \\ C \\tag{13}\n$$\n\n$$\nC_3 = \\Delta \\ l \\ d \\ q \\ C_1 \\tag{14}\n$$\n\nAs the construction of the tubing curve starts from the free end of the suspension tubing, the drag and weight of the air-speed head must be established first.\n\nThen the chord curve is built up with the above quantities as follows:", "timestamp": "2026-07-19T18:28:13.182307+00:00"}
{"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 16, "total_pages": 20, "image_filename": "19930091692_p16.jpg", "text": "12 REPORT NO. 617—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n5. Schäfer, D.: Neuere Anschauungen über motorische Entzündungs- und Verbrennungsvorgänge. Jahrbuch der Schiffbautechnischen Gesellschaft, Bd. 33, 1932, S. 181–211.\n\n6. Bird, A. L.: Experiments on Oil Jets and Their Ignition. Proc. Inst. Mech. Engrs., 1926, vol. II (Nov.), pp. 955–995.\n\n7. Bird, A. L.: Some Characteristics of Nozzles and Sprays for Oil Engines. Trans. Second World Power Conference, Bd. VIII, sec. 29, VDI-Verlag G. M. B. H. (Berlin), 1930, pp. 260–270.\n\n8. Hartner-Seberich: Der Zündverzug bei flüssigen Brennstoffen. Forschungsarbeiten auf dem Gebiete des Ingenieurwesens, Heft 290, VDI-Verlag G. M. B. H. (Berlin), 1928.\n\n9. Neumann, Kurt: Experiments on Self-Ignition of Liquid Fuels. T. M. No. 391, N. A. C. A., 1926.\n\n10. Wentzel, Wolfram: Der Zünd- und Verbrennungsvorgang im kompressorlosen Dieselmotor. VDI-Forschungsheft 366, VDI-Verlag (Berlin), 1934, S. 14–26.\n\n11. Michailova, M. N., and Neumann, M. B.: The Cetene Scale and the Induction Period Preceding the Spontaneous Ignition of Diesel Fuels in Bombs. T. M. No. 813, N. A. C. A., 1936.\n\n12. Hetzel, T. B.: The Development of Diesel Fuel Testing. Eng. Exp. Sta. Bull. No. 45, Penn. State Coll., 1936.\n\n13. MacGregor, J. R.: Diesel Fuels—Significance of Ignition Characteristics. S. A. E. Jour., vol. 38, no. 6, June 1936, pp. 217–223.\n\n14. Rose, R. A., Wilson, G. C., and Benedict, R. R.: Photo-Electric Combustion Analysis. S. A. E. Jour., vol. 39, no. 5, Nov. 1936, pp. 459–468.\n\n15. Tozier, R. E.: The N. A. C. A. Optical Engine Indicator. T. N. No. 634, N. A. C. A., 1938.\n\n16. Baxley, C. H., and Rendel, T. B.: Report of the Volunteer Group for Compression-Ignition Fuel Research. Paper presented at the May 4–9, 1937, meeting of the S. A. E., White Sulphur Springs, W. Va.\n\n17. Geschelin, Joseph: Fuels for Diesels. Auto Ind., vol. 75, no. 4, July 25, 1936, pp. 114–116.\n\n18. Selden, Robert F., and Spencer, Robert C.: Heat Transfer to Fuel Sprays Injected into Heated Gases. T. R. No. 580, N. A. C. A., 1937.\n\n19. Joachim, W. F., and Beardsley, Edward G.: The Effects of Fuel and Cylinder Gas Densities on the Characteristics of Fuel Sprays for Oil Engines. T. R. No. 281, N. A. C. A., 1927.\n\n20. LeMesurier, L. J., and Stansfield, R.: Combustion in Heavy Oil Engines. Trans. North-East Coast Institution of Engineers and Shipbuilders, England, vol. XLVIII, 1932, pp. 195–220.\n\n21. Judge, Arthur W.: High Speed Diesel Engines. Chapman and Hall, Ltd. (London), 1935, pp. 75–77.\n\n22. Moore, Charles S., and Collins, John H., Jr.: Compression-Ignition Engine Performance at Altitude. S. A. E. Jour., vol. 40, no. 6, June 1937, pp. 263–272.\n\n23. Moore, Charles S., and Collins, John H., Jr.: Compression-Ignition Engine Performance at Altitudes and at Various Air Pressures and Temperatures. T. N. No. 619, N. A. C. A., 1937.\n\n24. Wilson, G. C., and Rose, R. A.: Behavior of High- and Low-Cetane Diesel Fuels. S. A. E. Jour., vol. 41, no. 2, Aug. 1937, pp. 343–348.\n\n25. Moore, Charles S., and Foster, Hampton H.: Boosted Performance of a Compression-Ignition Engine with a Displacer Piston. T. N. No. 569, N. A. C. A., 1936.\n\n26. Rothrock, A. M.: Combustion in a High-Speed Compression-Ignition Engine. T. R. No. 401, N. A. C. A., 1931.\n\n27. Bone, William A., Newitt, Dudley M., and Townend, Donald T. A.: Gaseous Combustion at High Pressures. Longmans, Green and Co. (London), 1929, p. 337.\n\n28. Mucklow, G. F.: Experiments with a Supercharged Single-Cylinder Unit. R. & M. No. 1460, British A. R. C., 1932.\n\n29. Moore, C. S., and Foster, H. H.: Performance Tests of a Single-Cylinder Compression-Ignition Engine with a Displacer Piston. T. N. No. 518, N. A. C. A., 1935.\n\n30. MacGregor, J. R.: Influence of Humidity on Knock Ratings. S. A. E. Jour., vol. 40, no. 6, June 1937, pp. 243–249.", "timestamp": "2026-07-19T18:28:26.740425+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 75, "total_pages": 102, "image_filename": "19930094542_p75.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\nFigs.56,57\n\n[Figure: Graph of $\\frac{dc_a}{d\\alpha}$ vs $\\kappa$ for various $\\lambda$ values]\n\nFigure 56.- Lift slope $\\frac{dc_a}{d\\alpha}$ as a function of $\\kappa$.\n\n[Figure: Graph of $\\frac{dc_a}{d\\alpha}$ vs $\\lambda$ for various $\\kappa$ values]\n\nFigure 57.- Lift slope $\\frac{dc_a}{d\\alpha}$ as a function of $\\lambda$.", "timestamp": "2026-07-19T18:28:54.307279+00:00"}
{"citation_id": "19930091714", "source_url": "https://ntrs.nasa.gov/api/citations/19930091714/downloads/19930091714.pdf", "page_number": 34, "total_pages": 36, "image_filename": "19930091714_p34.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:28:58.964642+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 25, "total_pages": 43, "image_filename": "19930094544_p25.jpg", "text": "N.A.C.A. Technical Memorandum No. 872 23\n\n5. Weight Survey\n\nIn closing, a survey of the weights used for the framing and other parts of the dead weights of the airships mentioned should be given.\n\nIn figure 51 the weight ratios of framing, wiring, etc., to dead weight, as well as the ratio of the dead weight to static lift, are shown graphically to the same scale for the various airships. The square shown, representing 5 tons* serves as a measure of the actual weights. Primarily noteworthy in this drawing is the large ratio of the framing to dead weight in the two English airships R 100 and R 101. This probably lies, as is already mentioned, mainly in the close ring spacing as well as in the relatively high factors of safety chosen. The greater ratio of the wiring to dead weight in R 100 compared with R 101 is to be attributed to the greater ratio of the wiring area to the profile area of the hull in the case of R 100. The smaller weight ratio of outer cover and gas cells in the \"Akron,\" R 100, and R 101 in comparison with LZ 127 is to be attributed to the greater volume and the smaller slenderness ratio. The large ratio of the machinery installation in the LZ 127 and \"Akron\" in comparison with the R 100 probably lies largely in the relatively high unit weight of the Maybach engines chargeable to operating safety, and in comparison with the R 101 in the relatively low total power of the machinery installation of the R 101. Finally, in addition there is the large ratio of the crew and passenger spaces in the two English airships. This results from the fact that in the two English airships a relatively high weight has been expended for the furnishing of these spaces. The dashed lines in the case of R 101 show the ratio if approximately the same expenditure is made as in the case of the \"Graf Zeppelin.\"\n\nIn conclusion, it must be noted that in this comparison, in which all airships are assumed inflated with hydrogen, the \"Akron\" comes out somewhat too favorable, since with helium inflation the framing portion is more lightly stressed; however, offsetting this in the \"Akron\" is the additional weight of the water-recovery apparatus.\n\n---\n\n*Metric. 1 ton, metric = 2204.6 pounds.", "timestamp": "2026-07-19T18:29:02.732203+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 4, "total_pages": 28, "image_filename": "19930091697_p4.jpg", "text": "```markdown\n# NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nHEADQUARTERS, NAVY BUILDING, WASHINGTON, D. C.\nLABORATORIES, LANGLEY FIELD, VA.\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 to 15 by act approved March 2, 1929. The members are appointed by the President, and serve as such without compensation.\n\nJOSEPH S. AMES, Ph. D., *Chairman*,\nBaltimore, Md.\n\nDAVID W. TAYLOR, D. Eng., *Vice Chairman*,\nWashington, D. C.\n\nWILLIS RAY GREGG, Sc. D., *Chairman, Executive Committee*,\nChief, United States Weather Bureau.\n\nWILLIAM P. MACCRACKEN, J. D., *Vice Chairman, Executive Committee*,\nWashington, D. C.\n\nCHARLES G. ABBOT, Sc. D.,\nSecretary, Smithsonian Institution.\n\nLYMAN J. BRIGGS, Ph. D.,\nDirector, National Bureau of Standards.\n\nARTHUR B. COOK, Rear Admiral, United States Navy,\nChief, Bureau of Aeronautics, Navy Department.\n\nFRED D. FAGG, JR., J. D.,\nDirector of Air Commerce, Department of Commerce.\n\nHARRY F. GUGGENHEIM, M. A.,\nPort Washington, Long Island, N. Y.\n\nSYDNEY M. KRAUS, Captain, United States Navy,\nBureau of Aeronautics, Navy Department.\n\nCHARLES A. LINDBERGH, LL. D.,\nNew York City.\n\nAUGUSTINE W. ROBINS, Brigadier General, United States Army,\nChief Matériel Division, Air Corps, Wright Field,\nDayton, Ohio.\n\nEDWARD P. WARNER, M. S.,\nGreenwich, Conn.\n\nOSCAR WESTOVER, Major General, United States Army,\nChief of Air Corps, War Department.\n\nORVILLE WRIGHT, Sc. D.,\nDayton, Ohio.\n\nGEORGE W. LEWIS, *Director of Aeronautical Research*\n\nJOHN F. VICTORY, *Secretary*\n\nHENRY J. E. REID, *Engineer-in-Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.*\n\nJOHN J. IDE, *Technical Assistant in Europe, Paris, France*\n\n## TECHNICAL COMMITTEES\n\n| | |\n| :--- | :--- |\n| AERODYNAMICS | AIRCRAFT STRUCTURES |\n| POWER PLANTS FOR AIRCRAFT | AIRCRAFT ACCIDENTS |\n| AIRCRAFT MATERIALS | INVENTIONS AND DESIGNS |\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 MEMORIAL AERONAUTICAL LABORATORY**\nLANGLEY FIELD, VA.\n\nUnified conduct, for all agencies, of scientific research on the fundamental problems of flight.\n\n**OFFICE OF AERONAUTICAL INTELLIGENCE**\nWASHINGTON, D. C.\n\nCollection, classification, compilation, and dissemination of scientific and technical information on aeronautics.\n```", "timestamp": "2026-07-19T18:29:03.164595+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 13, "total_pages": 17, "image_filename": "19930094535_p13.jpg", "text": "12 N.A.C.A. Technical Memorandum No. 881\n\nTABLE I\nDaily and Monthly Effective Weathering Hours\n\n| Months | Average weathering hours | |\n| :--- | :---: | :---: |\n| | Daily | Monthly |\n| January | 0.8 | 25 |\n| February | 1.5 | 42 |\n| March | 2.5 | 78 |\n| April | 3.2 | 96 |\n| May | 4.4 | 136 |\n| June | 5.0 | 150 |\n| July | 4.8 | 149 |\n| August | 4.5 | 140 |\n| September | 3.8 | 114 |\n| October | 2.0 | 62 |\n| November | 1.1 | 33 |\n| December | .6 | 19 |", "timestamp": "2026-07-19T18:29:04.993325+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 2, "total_pages": 18, "image_filename": "19930091701_p2.jpg", "text": "# AERONAUTIC SYMBOLS\n\n## 1. FUNDAMENTAL AND DERIVED UNITS\n\n| | Symbol | Metric | | English | |\n|---|---|---|---|---|---|\n| | | Unit | Abbreviation | Unit | Abbreviation |\n| Length…… | $l$ | meter…… | m | foot (or mile)…… | ft. (or mi.) |\n| Time…… | $t$ | second…… | s | second (or hour)…… | sec. (or hr.) |\n| Force…… | $F$ | weight of 1 kilogram…… | kg | weight of 1 pound…… | lb. |\n| Power…… | $P$ | horsepower (metric)…… | k.p.h. | horsepower…… | hp. |\n| Speed…… | $V$ | kilometers per hour…… | m.p.s. | miles per hour…… | m.p.h. |\n| | | meters per second…… | | feet per second…… | f.p.s. |\n\n## 2. GENERAL SYMBOLS\n\n$W$, Weight=$mg$\n\n$g$, Standard acceleration of gravity=9.80665 m/s² or 32.1740 ft./sec.²\n\n$m$, Mass=$\\frac{W}{g}$\n\n$I$, Moment of inertia=$mk^2$. (Indicate axis of radius of gyration $k$ by proper subscript.)\n\n$\\mu$, Coefficient of viscosity\n\n$\\nu$, Kinematic viscosity\n\n$\\rho$, Density (mass per unit volume)  \nStandard density of dry air, 0.12497 kg·m⁻⁴·s² at 15° C. and 760 mm; or 0.002378 lb.-ft.⁻⁴ sec.²  \nSpecific weight of “standard” air, 1.2255 kg/m³ or 0.07651 lb./cu. ft.\n\n## 3. AERODYNAMIC SYMBOLS\n\n$S$, Area\n\n$S_w$, Area of wing\n\n$G$, Gap\n\n$b$, Span\n\n$c$, Chord\n\n$b^2$, Aspect ratio\n\n$S'$, True air speed\n\n$V$, Dynamic pressure=$\\frac{1}{2}\\rho V^2$\n\n$L$, Lift, absolute coefficient $C_L=\\frac{L}{qS}$\n\n$D$, Drag, absolute coefficient $C_D=\\frac{D}{qS}$\n\n$D_0$, Profile drag, absolute coefficient $C_{D_0}=\\frac{D_0}{qS}$\n\n$D_i$, Induced drag, absolute coefficient $C_{D_i}=\\frac{D_i}{qS}$\n\n$D_p$, Parasite drag, absolute coefficient $C_{D_p}=\\frac{D_p}{qS}$\n\n$C$, Cross-wind force, absolute coefficient $C_C=\\frac{C}{qS}$\n\n$R$, Resultant force\n\n$i_w$, Angle of setting of wings (relative to thrust line)\n\n$i_t$, Angle of stabilizer setting (relative to thrust line)\n\n$Q_t$, Resultant moment\n\n$\\Omega$, Resultant angular velocity\n\n$\\rho \\frac{Vl}{\\mu}$, Reynolds Number, where $l$ is a linear dimension  \n(e.g., for a model airfoil 3 in. chord, 100 m.p.h. normal pressure at 15° C., the corresponding number is 234,000; or for a model of 10 cm chord, 40 m.p.s., the corresponding number is 274,000)\n\n$C_p$, Center-of-pressure coefficient (ratio of distance of c.p. from leading edge to chord length)\n\n$\\alpha$, Angle of attack\n\n$\\epsilon$, Angle of downwash\n\n$\\alpha_0$, Angle of attack, infinite aspect ratio\n\n$\\alpha_i$, Angle of attack, induced\n\n$\\alpha_a$, Angle of attack, absolute (measured from zero-lift position)\n\n$\\gamma$, Flight-path angle", "timestamp": "2026-07-19T18:29:11.117955+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 4, "total_pages": 13, "image_filename": "19930091693_p4.jpg", "text": "# NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nHEADQUARTERS, NAVY BUILDING, WASHINGTON, D. C.  \nLABORATORIES, LANGLEY FIELD, VA.\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 to 15 by act approved March 2, 1929. The members are appointed by the President, and serve as such without compensation.\n\nJOSEPH S. AMES, Ph. D., Chairman,  \nBaltimore, Md.\n\nDAVID W. TAYLOR, D. Eng., Vice Chairman,  \nWashington, D. C.\n\nWILLIS RAY GREGG, Sc. D., Chairman, Executive Committee,  \nChief, United States Weather Bureau.\n\nWILLIAM P. MACCRACKEN, J. D., Vice Chairman, Executive Committee,  \nWashington, D. C.\n\nCHARLES G. ABBOT, Sc. D.,  \nSecretary, Smithsonian Institution.\n\nLYMAN J. BRIGGS, Ph. D.,  \nDirector, National Bureau of Standards.\n\nARTHUR B. COOK, Rear Admiral, United States Navy,  \nChief, Bureau of Aeronautics, Navy Department.\n\nFRED D. FAGE, JR., J. D.,  \nDirector of Air Commerce, Department of Commerce.\n\nHARRY F. GUGGENHEIM, M. A.,  \nPort Washington, Long Island, N. Y.\n\nSYDNEY M. KRAUS, Captain, United States Navy,  \nBureau of Aeronautics, Navy Department.\n\nCHARLES A. LINDBERGH, LL. D.,  \nNew York City.\n\nAUGUSTINE W. ROBINS, Brigadier General, United States Army,  \nChief Matériel Division, Air Corps, Wright Field,  \nDayton, Ohio.\n\nEDWARD P. WARNER, M. S.,  \nGreenwich, Conn.\n\nOSCAR WESTOVER, Major General, United States Army,  \nChief of Air Corps, War Department.\n\nORVILLE WRIGHT, Sc. D.,  \nDayton, Ohio.\n\nGEORGE W. LEWIS, Director of Aeronautical Research\n\nJOHN F. VICTORY, Secretary\n\nHENRY J. E. REID, Engineer-in-Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.\n\nJOHN J. IDE, Technical Assistant in Europe, Paris, France\n\n---\n\n## TECHNICAL COMMITTEES\n\n| AERODYNAMICS                  | AIRCRAFT STRUCTURES         |\n|-------------------------------|-----------------------------|\n| POWER PLANTS FOR AIRCRAFT     | AIRCRAFT ACCIDENTS          |\n| AIRCRAFT MATERIALS            | INVENTIONS AND DESIGNS      |\n\nCoordination of Research Needs of Military and Civil Aviation  \nPreparation of Research Programs  \nAllocation of Problems  \nPrevention of Duplication  \nConsideration of Inventions\n\n---\n\n## LANGLEY MEMORIAL AERONAUTICAL LABORATORY  \nLANGLEY FIELD, VA.\n\nUnified conduct, for all agencies, of  \nscientific research on the fundamental  \nproblems of flight.\n\n## OFFICE OF AERONAUTICAL INTELLIGENCE  \nWASHINGTON, D. C.\n\nCollection, classification, compilation,  \nand dissemination of scientific and technical information on aeronautics.", "timestamp": "2026-07-19T18:29:15.191474+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 45, "total_pages": 51, "image_filename": "19930094533_p45.jpg", "text": "N.A.C.A. Technical Memorandum No. 863\nFigs. 26,27,28\n\n[Figure: Graph showing Temperatures (y-axis, 0 to 1,0) vs Angle of attack (x-axis, -10 to 15). Data points plotted with a downward sloping trend line.]\nFigure 26.- Temperature curves of tap 4 plotted against angle of attack. (model)\n\n[Figure: Graph showing Temperatures (y-axis, 0 to 1,0) vs Angle of attack (x-axis, -10 to 15). Data points plotted with a downward sloping trend line.]\nFigure 27.- Temperature curves of tap 5 plotted against angle of attack. (model)\n\n[Figure: Graph showing Temperatures (y-axis, 0 to 1,0) vs Angle of attack (x-axis, -10 to 15). Data points plotted with a nearly flat trend line.]\nFigure 28.- Temperature curves of tap 6 plotted against angle of attack. (model)", "timestamp": "2026-07-19T18:29:16.766708+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 15, "total_pages": 76, "image_filename": "19930094549_p15.jpg", "text": "N.A.C.A. Technical Memorandum No. 867 13\n\n$$\nT = \\frac{2\\pi}{b} = 4\n$$\n\n$$\n\\tau_{\\frac{1}{2}} = \\frac{0.692}{|a|} = 0.185\n$$\n\nwhich means that after about 1/20 of a period the amplitude is again affected by a factor 0.5. In the case where $\\varphi = \\frac{\\pi}{2}$ the amplitudes start out from $\\rho$, but decrease more rapidly than according to the sinusoidal curve, the damped motion being that shown in figure 4. The airplane is, in fact, very energetically brought back to the angle of attack which corresponds to a zero moment about the center of gravity.\n\nThe same does not apply at the large angles of attack. At the maximum lift $\\frac{dC_z}{di} = 0$, and tests on numerous models show that $dC'_z/di'$ is also very small at this instant, whereas $dC_M/di$, on the contrary, maintains its former value. The theory thus predicts that the rapid oscillation may not be damped at the large angles of attack.\n\nb) Slow oscillation.- The slow oscillation may be studied separately by examining the modifications undergone by the path when the velocity is subject to a disturbance. It will readily be found that the flight path will be of an oscillatory character. These oscillations are necessarily accompanied by disturbances of the other variables. Taking account of the fact that the airplane strongly responds to the applied moments M and of the tendency, through the short-period oscillations, to assume rapidly the angle of attack of equilibrium, it might be supposed that these long-period oscillations are effected at a strictly constant angle of attack. This conclusion would, however, be premature. The moment M is, in fact, a function of two variables w and q, and we may write:\n\n$$\ndM = \\frac{\\partial M}{\\partial w} dw + \\frac{\\partial M}{\\partial q} dq\n$$\n\nThe rapid oscillation does not permit the continued existence of moments M and therefore, dM = 0. The irregularity of the flight path, however, corresponds to the existence of angular pitching velocities dq, so that the condition dM = 0 implies the existence of disturb-", "timestamp": "2026-07-19T18:29:18.654826+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 76, "total_pages": 102, "image_filename": "19930094542_p76.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:29:19.190974+00:00"}
{"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 17, "total_pages": 20, "image_filename": "19930091692_p17.jpg", "text": "TABLE I\n\nIGNITION LAGS ($\\times 10^3$) CORRESPONDING TO FIGURES 5 AND 7\n\n[Record numbers given in parentheses]\n\n| Air-fuel ratio | 15 | 20 | 23 | 25 | 27 | 30 | 33 | 35 | 37 | 40 | 45 | 50 | 60 | 75 | 80 | 90 | 100 | 120 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Density (lb./cu. ft.)** | | | | | | | | | | | | | | | | | | |\n| **Bomb temperature, 870° F.** | | | | | | | | | | | | | | | | | | |\n| 0.59 | | | | | | 7.0(186) | | | | 7.0(185) | | | | | | | | |\n| .89 | | 5.0(59) | | | | | | 5.0(194) | | 5.0(183) | | 5.0(192) | | | | | | |\n| 1.18 | | | | | | 5.0(181) | | | 4.3(190) | | 4.0(189) | 4.0(198) | | | 4.0(187) | | | |\n| 1.48 | | | | | | | 4.1(199) | 4.3(198) | 4.0(197) | | 4.1(196) | 4.1(195) | 4.3(194) | | | | 4.0(193) | |\n| 1.77 | | | | | | | | | 4.0(206) | 3.7(205) | 3.7(204) | 4.0(203) | 3.7(202) | | | 3.7(201) | | 4.0(200) |\n| 1.77 | | | | | | | | | 4.6(207) | | | | | | | | | |\n| **Bomb temperature, 1,000° F.** | | | | | | | | | | | | | | | | | | |\n| 0.59 | | 2.6(53) | | 2.8(103) | | 2.1(102) | | | | 2.0(101) | | | | | | | | |\n| .89 | | | 1.7(91) | 2.0(77) | | 2.0(76) | | | | | 2.0(75) | | 1.8(74) | | | | | |\n| .89 | | | | | | 1.7(90) | | | | | 1.7(89) | | 1.7(88) | | | | | |\n| 1.18 | | | | | 1.4(76) | 1.5(68) | 1.5(67) | | | 1.6(66) | | | 1.3(65) | | 1.4(64) | | | |\n| 1.18 | | | | | 1.5(69) | | | | | | | | | | | | | |\n| **Bomb temperature, 1,155° F.** | | | | | | | | | | | | | | | | | | |\n| 0.59 | 1.8(41) | 1.8(40) | | 1.8(173) | | 1.8(39) | | | | | | | | | | | | |\n| .59 | | 1.8(43) | | | | 1.8(167) | | | | | | | | | | | | |\n| .59 | | 1.8(168) | | | | | | | | | | | | | | | | |\n| .89 | | 1.3(38) | 1.2(37) | 1.3(151) | | 1.3(36) | | | | | 1.3(35) | | 1.3(148) | | | | | |\n| .89 | | 1.3(154) | 1.3(152) | 1.3(153) | | 1.3(150) | | | | | 1.3(149) | | | | | | | |\n| 1.18 | | | | | 1.2(134) | 1.2(133) | | | | 1.2(132) | | | 1.2(131) | | 1.2(130) | | | |\n| **Bomb temperature, 1,255° F.** | | | | | | | | | | | | | | | | | | |\n| 0.59 | | | | | | 1.7(219) | | | | 1.7(226) | | | | | | | | |\n| .59 | | | | | | 1.6(227) | | | | | | | | | | | | |\n| .89 | | | 1.0(232) | | | 1.2(229) | | | | | 1.3(230) | | 1.0(229) | | | | | |\n| .89 | | | | | | 1.2(231) | | | | | | | | | | | | |\n| 1.18 | | | 1.0(232) | | | 1.0(221) | | | | | 1.3(230) | | | | 1.0(222) | | | |\n| 1.48 | | | | | | | | | | | 0.8(223) | | | | | | | |\n\nU. S. GOVERNMENT PRINTING OFFICE: 1952\n\nAUTO-IGNITION AND COMBUSTION OF DIESEL FUEL IN A CONSTANT-VOLUME BOMB\n\n13", "timestamp": "2026-07-19T18:29:52.502545+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 26, "total_pages": 43, "image_filename": "19930094544_p26.jpg", "text": "24 N.A.C.A. Technical Memorandum No. 872\n\nIV. CONCLUSION\n\nIn covering the matters relating to the present position of airship construction it was possible only to a limited degree to go into them thoroughly. Particularly, only a part of the materials graciously made available by domestic and foreign airship authorities could be introduced. The foregoing discussion is intended primarily to give an idea as to what mental and material media have been used in airship construction up to the present time, and what guiding influence German airship construction has exerted on the previous development.\n\nTranslation by Ray E. Brown,  \nBureau of Aeronautics,  \nNavy Department.", "timestamp": "2026-07-19T18:29:52.763888+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 77, "total_pages": 102, "image_filename": "19930094542_p77.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\nFigs.58,59\n\n<!-- Image (224, 199, 808, 483) -->\n\nFigure 58.- Polars of wing in presence of propeller, $\\kappa = 9^\\circ$.\n\n<!-- Image (224, 563, 816, 846) -->\n\nFigure 59.- Polars of wing in presence of propeller. $\\kappa = 4^\\circ$.", "timestamp": "2026-07-19T18:29:57.059544+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 3, "total_pages": 18, "image_filename": "19930091701_p3.jpg", "text": "REPORT No. 626\n\nTHE TRANSITION PHASE IN THE TAKE-OFF  \nOF AN AIRPLANE\n\nBy J. W. WETMORE  \nLangley Memorial Aeronautical Laboratory\n\n61908—38—1", "timestamp": "2026-07-19T18:30:01.464129+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 19, "total_pages": 21, "image_filename": "19930094534_p19.jpg", "text": "N.A.C.A. Technical Memorandum No. 882\n\ncm²\nVolume of tube section\n1.5\n1.3\n1.1\n0.9\n0.7\n0 5 10 15 20 25 30 mm\nReduction of diameter.\nTube 50x1\nTube 45x1\nTube 40x1\nTube 35x1\n\n50x1\n45x1\n40x1\n35x1\noriginal\ndiameter\n20φ\n20φ\n20φ\n20φ\n\nFigure 9.- Cross-sectional change on contracting duralumin tubes\nof different original diameter to 20 mm diameter.\n\nFIG.9", "timestamp": "2026-07-19T18:30:05.069845+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 35, "total_pages": 43, "image_filename": "19930094538_p35.jpg", "text": "N. A. C. A. Technical Memorandum No. 878\nFigs. 8,9,12\n\n<!-- Image (224, 89, 823, 940) -->\n\nFigure 8.- Normal stress distribution $\\sigma_n$ across panel width.\n\nLoad through\nlongitudinal\nsections\n(individual\nloads)\n\nBulkhead stress\nFigure 9.- Loading of stiffener system.\n\nOrigin\nof transverse\nload p\n\nLoading of\nstringers.\n\nFigure 12.- Experimental setup for twisting tests.", "timestamp": "2026-07-19T18:30:09.126631+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 30, "total_pages": 32, "image_filename": "19930094552_p30.jpg", "text": "N.A.C.A. Technical Memorandum No. 864 Figs. 19,20\n\n<!-- Image (206, 155, 785, 412) -->\n\n$$R = \\frac{Pa}{a'}$$\n\n$$\\tau s x = R - P = P \\left( \\frac{a-a'}{a'} \\right)$$\n\nFigure 19.- Stress distribution in a wall at force application side.\n\n<!-- Image (206, 488, 805, 772) -->\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-19T18:30:13.732907+00:00"}
{"citation_id": "19930091714", "source_url": "https://ntrs.nasa.gov/api/citations/19930091714/downloads/19930091714.pdf", "page_number": 35, "total_pages": 36, "image_filename": "19930091714_p35.jpg", "text": "Positive directions of axes and angles (forces and moments) are shown by arrows\n\n| Axis | | Force (parallel to axis) symbol | Moment about axis | | | Angle | | Velocities | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Designation | Symbol | | Designation | Symbol | Positive direction | Designation | Symbol | Linear (component along axis) | Angular |\n| Longitudinal<br>Lateral<br>Normal | $X$<br>$Y$<br>$Z$ | $X$<br>$Y$<br>$Z$ | Rolling<br>Pitching<br>Yawing | $L$<br>$M$<br>$N$ | $Y \\longrightarrow Z$<br>$Z \\longrightarrow X$<br>$X \\longrightarrow Y$ | Roll<br>Pitch<br>Yaw | $\\phi$<br>$\\theta$<br>$\\psi$ | $u$<br>$v$<br>$w$ | $p$<br>$q$<br>$r$ |\n\nAbsolute coefficients of moment\n$C_l = \\frac{L}{q b S}$ (rolling)\n$C_m = \\frac{M}{q c S}$ (pitching)\n$C_n = \\frac{N}{q b S}$ (yawing)\n\nAngle of set of control surface (relative to neutral position), $\\delta$. (Indicate surface by proper subscript.)\n\n**4. PROPELLER SYMBOLS**\n\n$D$, Diameter\n$p$, Geometric pitch\n$p/D$, Pitch ratio\n$V$, Inflow velocity\n$V_s$, Slipstream velocity\n$T$, Thrust, absolute coefficient $C_T = \\frac{T}{\\rho n^2 D^4}$\n$Q$, Torque, absolute coefficient $C_Q = \\frac{Q}{\\rho n^2 D^5}$\n$P$, Power, absolute coefficient $C_P = \\frac{P}{\\rho n^3 D^5}$\n$C_s$, Speed-power coefficient $= \\sqrt[5]{\\frac{\\rho V^5}{P n^2}}$\n$\\eta$, Efficiency\n$n$, Revolutions per second, r.p.s.\n$\\Phi$, Effective helix angle $= \\tan^{-1} \\left( \\frac{V}{2 \\pi r n} \\right)$\n\n**5. NUMERICAL RELATIONS**\n\n1 hp. = 76.04 kg-m/s = 550 ft-lb./sec.\n1 metric horsepower = 1.0132 hp.\n1 m.p.h. = 0.4470 m.p.s.\n1 m.p.s. = 2.2369 m.p.h.\n1 lb. = 0.4536 kg.\n1 kg = 2.2046 lb.\n1 mi. = 1,609.35 m = 5,280 ft.\n1 m = 3.2808 ft.", "timestamp": "2026-07-19T18:30:16.559884+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 46, "total_pages": 51, "image_filename": "19930094533_p46.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\nFigs. 29,30,31\n\n[Figure: Graph plotting Temperatures (0 to 1.0) against Angle of attack (-10 to 10). Data points are plotted with circles and a trend line is drawn.]\nFigure 29.- Temperature curves of tap 7 plotted against the angle of attack(model).\n\n[Figure: Graph plotting Temperatures (0 to 1.0) against Angle of attack (-10 to 15). Data points are plotted with circles and a trend line is drawn.]\nFigure 30.- Temperature curves of tap 8 plotted against the angle of attack(model).\n\n[Figure: Graph plotting Temperatures (0 to 1.0) against Angle of attack (-10 to 15). Data points are plotted with circles and a trend line is drawn.]\nFigure 31.- Temperature curves of tap 9 plotted against the angle of attack(model).", "timestamp": "2026-07-19T18:30:18.059163+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 5, "total_pages": 13, "image_filename": "19930091693_p5.jpg", "text": "REPORT No. 618\n\nCOMPARATIVE FLIGHT AND FULL-SCALE WIND-TUNNEL MEASUREMENTS OF THE MAXIMUM LIFT OF AN AIRPLANE\n\nBy ABE SILVERSTEIN, S. KATZOFF, and JAMES A. HOOTMAN\n\nSUMMARY\n\nDeterminations of the power-off maximum lift of a Fairchild 22 airplane were made in the N. A. C. A. full-scale wind tunnel and in flight. The results from the two types of test were in satisfactory agreement. It was found that, when the airplane was rotated positively in pitch through the angle of stall at rates of the order of $0.1^\\circ$ per second, the maximum lift coefficient was considerably higher than that obtained in the standard tests, in which the forces are measured with the angles of attack fixed. Scale effect on the maximum lift coefficient was also investigated.\n\nINTRODUCTION\n\nThe purpose of the present investigation was to obtain a direct comparison between flight and full-scale wind-tunnel measurements of the maximum lift coefficient of a Fairchild 22 airplane. The comparison was desirable in order to indicate the extent to which the various wind-tunnel effects and both wind-tunnel and flight techniques might influence maximum-lift determinations. The turbulence in the full-scale tunnel (reference 1) was of particular concern.\n\nObviously, a high order of accuracy must exist in both flight and wind-tunnel measurements if the comparison is to be significant. The many possibilities for experimental error in both series of tests required that great care be exercised in obtaining the test data. Previous comparisons between flight and full-scale wind-tunnel results (references 2 and 3) were incidental to other studies and unsuited for the accuracy here desired.\n\nInasmuch as, in the flight determinations of maximum lift, the airplane was rotated through the angle of stall, some wind-tunnel tests were made with the airplane rotating at corresponding angular velocities in order to investigate the effect of this technique on the results.\n\nWind-tunnel tests to determine the Reynolds Number effects on the maximum lift coefficient and on the minimum drag coefficient were also made.\n\nFULL-SCALE WIND-TUNNEL INVESTIGATION\n\nAPPARATUS AND TESTS\n\nThe N. A. C. A. full-scale wind tunnel and its equipment are described in reference 2. Figure 1 is a 3-view drawing of the Fairchild 22 parasol monoplane. Two positions of the center of gravity are indicated, corresponding to two airplane loadings used in the flight tests. The airplane was equipped for these tests with a specially surfaced wing of N. A. C. A. 2R₁₂ section. A paint filler was applied over the forward 15 percent\n\n[Figure: Fairchild 22 airplane with wing of N. A. C. A. 2R₁₂ section.]\n\nWeight of  \nairplane  \n(lb.)  \n1,613  \n2,219  \n\na  \n(in.)  \n69¾  \n71⅝  \n\nb  \n(in.)  \n0⅞  \n2⅛  \n\nof the wing surface and waxed to a reflecting finish, the polish being maintained throughout both wind-tunnel and flight tests. The purpose of the polish was not only to provide a reproducible surface but also to increase any differences between the wind-tunnel and flight results due to turbulence in the wind tunnel. All the tests were made with the airplane at $0^\\circ$ yaw and", "timestamp": "2026-07-19T18:30:20.582645+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 12, "total_pages": 18, "image_filename": "19930094551_p12.jpg", "text": "N.A.C.A. Technical Memorandum No. 865 11\n\n1) Determine value $\\frac{2 A + G_s}{C_3}$ from figure 9a (for start $A_{n-1} = G$) and ascertain in figure 9.\n\n2) Determine $\\frac{2 E_{n-1}}{C_3}$ in figure 9 (for start $E_{n-1} = W$).\n\n3) Connect both points and shift parallel upward, until equal angles appear on the ordinates (plot the 1st chord element).\n\n4) Define the new A from the preceding A by subtracting $G_s$ and adding the amount of lift which corresponds to the just found $\\alpha_s$ in figure 11.\n\n5) Determine the new B from the previous B, by adding the amount of the drag which corresponds to the just found $\\alpha_s$ in figure 10.\n\n6) Proceed with A and B as under 1, 2, 3.\n\nConcerning point 3, it should be noted that the parallel shift merely fulfills the condition (equation 12), which postulates that the sum of the terms on the left-hand side of the equation is equal to the term on the right-hand side.\n\nThe tubing curves I, II, III in figure 12 were obtained by this method. Curves I and III are for the heavy air-speed head at 210 and 360 km/h, while curve II indicates the position of the small air-speed head at 210 km/h.\n\nThe dots designating the various speeds on which the calculation was based, are actual positions of the air-speed head in space as measured by bearing in flight. The comparison of the measured and the computed values manifests ample agreement.\n\nCurve IV in figure 12, also established by nomographic method, is for a flying speed of 600 km/h and 25 m suspension length for an air-speed head of 9 kg.\n\nTranslation by J. Vanier,\nNational Advisory Committee\nfor Aeronautics.", "timestamp": "2026-07-19T18:30:24.451388+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 11, "total_pages": 16, "image_filename": "19930094564_p11.jpg", "text": "10 N.A.C.A. Technical Memorandum No. 852\n\nVI. RECAPITULATION OF THE RESULTS\n\nFigure 12 shows the maximum lift for the airfoil series with and without split flap for the average value of the practical range of Reynolds Numbers (Reffective ≅ 4 x 10⁶) plotted against the airfoil thickness.\n\nAs regards $c_{a_{max}}$ the findings are:\n\n1) Without split flap, airfoil series 230 is superior to the other series in thickness range of between 10 to 21 percent.\n\n2) With split flap, airfoil series 24 gives the best results between 9-and 17-percent thickness, but for still greater thickness the symmetrical airfoils of the series 00 are superior.\n\nA survey of the rating factor $c_{a_{max}}/c_{wp}(c_a = 0.1)$ is afforded in figure 13, where this factor has been plotted against the airfoil thickness with and without split flap. The mean value of the practical range of Reynolds Number was assumed at 4 x 10⁶ for the $c_{a_{max}}$ values, and the $c_{wp}(c_a = 0.1)$ values referred to their mean value of the practical range R ≅ 20 x 10⁶.\n\nRegarding $c_{a_{max}}/c_{wp}(c_a = 0.1)$ figure 13 discloses the following:\n\n1) Without split flap, airfoil series 230 (2-percent camber at 15-percent chord) is superior throughout the explored thickness range (9 to 21 percent). On series 24 the optimum $c_{a_{max}}/c_{wp}(c_a = 0.1)$ is reached at around 9-percent thickness.\n\n2) With split flap, airfoil series 24 (2-percent camber at 40-percent chord) excels below 15-percent thickness, while the symmetrical airfoil series (0-percent camber) gives the best results when the thickness exceeds 15 percent. The optimum $c_{a_{max}}/c_{wp}(c_a = 0.1)$ for airfoil 24 is reached with 332 at approximately 12-percent thickness. A 50-percent thickness", "timestamp": "2026-07-19T18:30:49.791539+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 36, "total_pages": 43, "image_filename": "19930094538_p36.jpg", "text": "N. A. C. A. Technical Memorandum No. 878\n\nCylinder No. II\n\n[Figure: Schematic of cylinder with sections I to VI, and three circular cross-sections labeled II, III, IV with angular markings and “800°” annotations]\n\nRivet 5ᵈ  \n74.5  \n12  \n3.5  \n76.5  \nFₗ = 54 mm²\n\nRivet 2ᵈ  \n21  \n8  \n12  \n8  \n4  \n15  \nØ7  \nFₗ = 57 mm²\n\nRivet 5ᵈ  \n15.5  \n0.5  \n20  \nØ7.5  \n10  \n14  \n10  \nFₗ = 58 mm²\n\nRivet 3ᵈ  \n20  \n19  \nProfile  \nRivet 4ᵈ  \n0.8  \n16  \n2”  \n5.5  \nFᵧ = 68 mm²\n\nFigure 11.—  \nSchematic presentation of the test specimens.\n\n[Graph: Curve rising from origin, x-axis labeled 0 to 0.5 in steps of 0.1, y-axis labeled 0 to 1.0 in steps of 0.2]\n\nFigure 10.— Curve of integral  \n$$ J(\\omega) = \\frac{2}{\\pi} \\int_{0}^{\\pi/2} \\sin^{1/\\omega} \\chi \\, d\\chi. $$\n\nFigs. 10, 11", "timestamp": "2026-07-19T18:31:04.244215+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 13, "total_pages": 18, "image_filename": "19930094551_p13.jpg", "text": "12 N.A.C.A. Technical Memorandum No. 865\n\nREFERENCES\n\n1. Kiel, Georg: Fehlerabschätzung bei Standruckeichungen mittels unter dem Flugzeug geschleppter Sonden. Luftfahrtforschung, vol. 14, no. 6, June 20, 1937, pp. 310-313.\n\n2. Glauert, H.: Heavy Flexible Cable for Towing a Heavy Body Below an Aeroplane. R. & M. No. 1592, British A.R.C., 1934.", "timestamp": "2026-07-19T18:31:08.128039+00:00"}
{"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 48, "total_pages": 50, "image_filename": "19930094543_p48.jpg", "text": "N.A.C.A. Technical Memorandum No. 873\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-19T18:31:09.902126+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 14, "total_pages": 17, "image_filename": "19930094535_p14.jpg", "text": "N.A.C.A. Technical Memorandum No. 881\nFigs.1,2\n\n[Figure: View through gunner's cockpit of a fighter plane provided with artificial glass of acrylate plastic.]\n\nView through gunner's cockpit of a fighter plane provided with artificial glass of acrylate plastic.\n\n[Figure: Pieces of sheet silicate glass after testing for resistance against excess pressure on one side.]\n[Figure: Fragments of hard (heat treated) glass.]\n\nFigure 1.- Pieces of sheet silicate glass after testing for resistance against excess pressure on one side.\nFigure 2.- Fragments of hard (heat treated) glass.", "timestamp": "2026-07-19T18:31:10.134788+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 16, "total_pages": 76, "image_filename": "19930094549_p16.jpg", "text": "14 N.A.C.A. Technical Memorandum No. 867\n\nances of angle of attack $\\delta w$ or $\\delta i$ connected with $\\delta q$ by\n\n$$\n\\frac{\\partial C_M}{\\partial w} \\delta w + \\frac{\\partial C_M}{\\partial q} \\delta q = 0\n$$\n\nWe thus have:\n\n$$\n\\delta q \\left( - \\frac{l}{V} \\frac{S l \\cdot i^2}{S l \\cdot i^2} \\frac{d C'_M}{d i'} \\right) = - \\delta w \\frac{\\partial C_M}{\\partial w}\n$$\n\n$$\n= \\delta w \\frac{\\partial C_M}{\\partial i} \\frac{1}{V}\n$$\n\nwhence\n\n$$\n\\frac{\\delta q}{\\delta w} = - \\frac{1}{l} \\frac{S l^2}{S l \\cdot i^2} \\frac{\\frac{\\partial C_M}{\\partial i}}{\\frac{\\partial C'_M}{\\partial i'}}\n$$\n\nThis condition will only be true provided that:\n\n1) $\\rho_3$ and $\\rho_2$ are in the previously given ratio, and\n\n2) $\\delta q$ and $\\delta w$ corresponding to the short-period oscillation are out of phase by $\\pi$.\n\nIn the case of the numerical examples, we should have:\n\n$\\rho_3 = 0.01385 \\rho_2$ for the airplane for which $\\mu = 0.002$\n\n$\\rho_3 = 0.0352 \\rho_2$ $\\mu = 0.008$\n\nIt may be seen from the numerical tables computed, that this relation is practically verified and that the phase displacements are by $\\pi \\pm 3^\\circ$.\n\nDETERMINATION OF THE MOTION AFTER INITIAL DISTURBANCE\n\nIn the general motion each of the elementary motions is involved to a degree that depends on the nature of the initial disturbance. Preliminary studies on automatic stabilizers have led us to investigate how a given initial", "timestamp": "2026-07-19T18:31:10.543799+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 10, "total_pages": 22, "image_filename": "19930091655_p10.jpg", "text": "6\nREPORT NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nThe results obtained with benzene are presented in table III and figure 7. These tests were limited to the smaller weights because of the greater pressure changes per unit weight of liquid.\n\nAn inspection of the experimental records reveals two time intervals more or less clearly defined on all the start of injection corresponds, with one exception, to the vertical line appearing on some of the records and is coincident with the A points; i. e., the decrease in the gas pressure began immediately after the first part of the fuel charge entered the bomb. The one exception (record 295, fig. 5) was due to improper synchronization of the injection start and the timing spark. The A-B interval and the pressure drop associated with it are indicative of processes occurring immediately after the injection starts. This interval is therefore of primary interest with respect to compression-ignition\n\n<!-- Image (142, 182, 896, 769) -->\n\n(a) Gas density, 4.73 grams per liter; gas-fuel ratio, 10.\nFIGURE 4.—Variation of pressure drop with gas temperature. Diesel fuel; fuel weight, 0.284 gram.\n\npressure-time curves. Three characteristic points are designated on all the records reproduced: A, the point at which the pressure drop begins; B, the end of the initial pressure drop for which the rate was essentially constant; and C, the minimum pressure point. The", "timestamp": "2026-07-19T18:31:11.135389+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 47, "total_pages": 51, "image_filename": "19930094533_p47.jpg", "text": "N.A.C.A. Technical Memorandum No. 683\nFigs. 32,33,34\n\n[Figure: Graph showing Temperatures vs Angle of attack. Y-axis labeled \"Temperatures\" with values 0, 0.5, 0.1. X-axis labeled \"Angle of attack\" with values -10, -5, 0, 5, 10, 15. Data points plotted with a trend line.]\nFigure 32.- Temperature curves of tap 10 plotted against angle of attack. (model)\n\n[Figure: Graph showing Temperatures vs Angle of attack. Y-axis labeled \"Temperatures\" with values 0, 0.5, 1.0. X-axis labeled \"Angle of attack\" with values -10, -5, 0, 5, 10, 15. Data points plotted with a trend line.]\nFigure 33.- Temperature curves of tap 11 plotted against angle of attack. (model)\n\n[Figure: Graph showing Temperatures vs Angle of attack. Y-axis labeled \"Temperatures\" with values 0, 0.5, 1.0. X-axis labeled \"Angle of attack\" with values -10, -5, 0, 5, 10, 15. Data points plotted with a trend line.]\nFigure 34.- Temperature curves of tap 12 plotted against angle of attack. (model)", "timestamp": "2026-07-19T18:31:16.974837+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 78, "total_pages": 102, "image_filename": "19930094542_p78.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:31:18.626267+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 27, "total_pages": 43, "image_filename": "19930094544_p27.jpg", "text": "N.A.C.A. Technical Memorandum No. 872 25\n\nREFERENCES\n\n1. Moedebeck, H.: Die Luftschiffahrt in ihrer neuesten Entwicklung (Airship Transportation in Its Latest Development). Verlag Ernst Mittler & Sohn, Berlin, 1887.\n\n2. Roeser: Aus der Geschichte des deutschen Starrluftschiffes (From the History of the German Rigid Airship). Z.F.M., vol. 16, no. 13, 1925, pp. 261-272.\n\n3. Dürr, L.: 25 Jahre Luftschiffbau Zeppelin (25 Years of Luftschiffbau Zeppelin). VDI-Verlag, Berlin, 1924.\n\n4. Luftschiffbau Schütte-Lanz (The Schütte-Lanz Airship Construction), published by J. Schütte. Berlin 1926. Verlag von R. Oldenbourg.\n\n5. Müller-Breslau, H.: Zur Geschichte des Zeppelin-Luftschiffes. (On the History of the Zeppelin Airship). Reprint of Transaction of the Verein zur Beförderung des Gewerbefleisses, 1914.\n\n6. Basenach, R.: Bau und Betrieb von Prallluftschiffen (Construction and Operation of Pressure Airships). Verlag von F. B. Auffahrth. Leipzig, 1911-12.\n\n7. V. Parseval, A.: Der Parseval-Ballon (The Parseval Airship). Z.F.M., vol. 1, no. 7, 1910, pp. 76-80.\n\n8. Krell, O.: Das Luftschiff der Siemens-Schuckert-Werke und seine Halle (The Airship of the Siemens-Schuckert-Werke and Its Hangar). Z.F.M., vol. 2, no. 5, 1911, pp. 61-64, no. 6, pp. 72-76, and no. 19, p. 280.\n\n9. Lehmann, E. A.: Das Luftschiff \"Graf Zeppelin\" (The Airship \"Graf Zeppelin\"). Yearbook of the Schiffbautechnische Gesellschaft, vol. 32, 1931, pp. 65-92.\n\n10. Scherz, W.: Das Luftschiff \"Graf Zeppelin\" (The Airship \"Graf Zeppelin\"). Illustr. Flugwoche, no. 9, 1928, pp. 253-258.", "timestamp": "2026-07-19T18:31:18.791806+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 6, "total_pages": 13, "image_filename": "19930091693_p6.jpg", "text": "```markdown\n2\nREPORT NO. 618—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nroll and with the ailerons locked in the neutral position.\nThe propeller was locked in the vertical position except\nwhere otherwise noted. Figure 2 shows the airplane\nmounted on the wind-tunnel balance supports.\n\nThe tests were of two types. The tests designated\n“standard” were similar to those normally made in the\nwind tunnel, in which the force readings are not taken\nuntil a number of seconds after the airplane has been\nbrought to rest at the desired attitude. In the other\ntype of test, force readings were taken at regular inter-\nvals while the angle of attack was being changed at a\nconstant rate. The rates of change of angle of attack\nin these runs were varied between $0.025^\\circ$ and $0.2^\\circ$ per\nsecond, this range including that used in the flight\ntests.\n\nExcept in the tests to determine minimum drag, all\nmeasurements were made in the region of maximum lift,\nthe angle-of-attack range being usually from $12^\\circ$ to $20^\\circ$.\n\nScale effect.—Figure 4 shows the lift curves obtained\nat five different air speeds for the airplane with the\nhorizontal tail removed. It will be observed that,\nwith increasing air speed, the maximum lift coefficient\nreaches higher values and the entire lift curve is slightly\nraised, even over the linear range. The break in the\nlift curve at the peak becomes sharper with increasing\nair speed, indicating a variation in the mechanism of\nstalling. It may be noted that beyond the stall the\nlift curve represents only a rough average, for there is\nwide scatter of the points in this region.\n\nThe variation of the maximum lift coefficient with\nair speed is shown in figure 5 for three different test\n\n<!-- Image (119, 386, 490, 586) -->\nFIGURE 2.—Fairchild 22 airplane in full-scale wind tunnel.\n\nAir speeds ranged from 29 to 63 miles per hour, except\nfor the minimum-drag tests, in which air speeds up to\n119 miles per hour were used. Much of the work was\ndone at an air speed of 56 miles per hour, which is\napproximately flight speed at maximum lift.\n\n<!-- Image (521, 297, 919, 673) -->\nFIGURE 3.—Characteristics of Fairchild 22 airplane with the horizontal tail removed.\n$C_m$ for airplane based on weight of 1,613 pounds. Air speed, 56 m. p. h.\n\nRESULTS AND DISCUSSION\n\nThe results of the wind-tunnel tests are summarized\nin figures 3 to 19. Except where otherwise noted, the\nfigures refer to tests of the standard type. All measure-\nments were corrected for jet-boundary effect at the\nwing, balance-support tare values, and blocking, as\ndescribed in reference 4.\n\nIn figure 3 the lift, the drag, and the pitching-\nmoment coefficients, and the lift-drag ratio are plotted\nagainst the angle of attack of the thrust axis, $\\alpha_T$, for\nthe airplane with the horizontal tail removed. The\ntest data were obtained at an air speed of 56 miles per\nhour. The pitching-moment coefficient was based on\na center-of-gravity position as determined for a gross\nweight of 1,613 pounds. (See fig. 1.)\n\nconditions, namely, tail removed, tail on, and tail on\nwith the angle changing at the rate of $0.1^\\circ$ per second.\nThe indicated stabilizer angle ($\\delta_s$) and elevator angle\n($\\delta_e$) correspond approximately to trim at maximum lift.\nAll the tests show essentially the same variation of\nmaximum lift coefficient with air speed. Results from\nthe tests in the variable-density tunnel of the plain\nairfoil are also shown in the figure, and it will be seen\nthat, except for a vertical displacement due to difference\nin plan form and to the effect of the fuselage, the agree-\nment is very good.\n\nExperiments to determine whether the presence of\nthe propeller fixed in the vertical or the horizontal\nposition materially influenced the maximum lift showed\nthat the propeller in either position had a negligible\neffect (fig. 6).\n```", "timestamp": "2026-07-19T18:31:26.449240+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 19, "total_pages": 47, "image_filename": "19930093641_p19.jpg", "text": "L-452.\n\nThe results indicate that the enclosed-engine arrangements with tractor propellers 0.26c and 0.39c ahead of the wing and with pusher propellers are of about equal merit; the 0.13c tractor-propeller position, however, shows a definitely lower over-all efficiency.\n\nLangley Memorial Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nLangley Field, Va., February 17, 1938.\n\nREFERENCES\n\n1. DeFrance, Smith J.: The N.A.C.A. Full-Scale Wind Tunnel. NACA Rep. No. 459, 1933.\n\n2. Silverstein, Abe, and Nickle, F. R.: Preliminary Full-Scale Wind-Tunnel Investigation of Wing Ducts for Radiators. NACA ACR, March 1938.", "timestamp": "2026-07-19T18:31:30.501320+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 20, "total_pages": 21, "image_filename": "19930094534_p20.jpg", "text": "```markdown\n| Construction sizes | | Rivet diameter | | | Type of riveting | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | **2** | **2.6** | **3** | | | | |\n| Maximum sheet thickness | $s_x$ | 1 | 1.2 | 1.2 | [Figure: Rivet diagram labeled p] | [Figure: Rivet diagram labeled FS₁/F] | | Dimpled sheets (x) |\n| Maximum gripping length | $S_{x1}$ | 1.4 | 1.8 | 2.1 | | | | |\n| Maximum gripping length | $S_{x2}$ | 2 | 2.6 | 3 | | | | |\n| Normal sheet thickness | $s_{y1}$ | 1.2 | 1.5 | 1.8 | [Figure: Rivet diagram labeled R/FSy] | [Figure: Rivet diagram labeled FS₁/F] | Thickest sheet counter-sunk | Counter-sunk sheets (y) and flat counter-sunk sheets |\n| Minimum sheet thickness | | 1 | 1.2 | 1.5 | | | | |\n| Countersink diameter | $D_1$ | 4.8 | 5.7 | 6.5 | [Figure: Rivet diagram labeled R/FSy] | [Figure: Rivet diagram labeled FS₁/F] | Counter-sink start 3rd sheet | |\n| Least sheet thickness for second sheet in flap riveting, (hinged plates) | $s_{y2}$ | 0.6 | 0.8 | 1 | [Figure: Rivet diagram labeled FS₁/F] | [Figure: Rivet diagram labeled R/FS] | Outside sheet counter-sunk | |\n| Related countersink diameter | $D_2$ | 3.4 | 4.4 | 5.2 | | | | |\n| Maximum sheet thickness | $s_z$ | 0.6 | 0.8 | 1 | [Figure: Rivet diagram labeled R/FSz] | [Figure: Rivet diagram labeled FR/FSz] | | Dimpled sheets (z) |\n| Maximum gripping length | $S_{z1}$ | 1.4 | 1.8 | 2.1 | | | | |\n| Maximum gripping length | $S_{z2}$ | 1 | 1.3 | 1.5 | | | | |\n\nBy unrestricted accessibility:\n(wings and fuselages)\nFlat countersunk riveting\n$FS_x/F$ or $FS_y/F$ ;\nMushroom riveting or $R/FS_y$\n\nOn closed sections,\nadjoining walls, etc.\nFlat countersunk riveting,\n$FS_x/F$ or $FS_y/F$ ; by hand\nriveting also mushroom\nriveting, $R/FS_y$ , $R/FS_z$\n\nFor low clearance\n$PS_x/F$ for H $\\ge$ 12 ;\n$R/FS_z$ \" H $\\le S_{z1}$\n\nEdge riveting\nof thin sheets\n$R/FS_z$\n\nFor flaps\nin air\nstream\n$FS/FS$\nor\n$FR/FS_z$\n\nFigure 14.- Skin riveting methods.\n\nN.A.C.A. Technical Memorandum No. 382\nFIG. 14\n```", "timestamp": "2026-07-19T18:31:30.715229+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 37, "total_pages": 43, "image_filename": "19930094538_p37.jpg", "text": "Test stations on\n\nCylinder II\n\nS are section centroids\n\nTensiometer\n\nSection\n\nTest stations at bulkhead\n\nBulkhead\n\nTest points for tension diagonal stresses\n\nCenter of section\n\nCut:a-a\n\nValley\n\nHill\n\nValley\n\nHill\n\nValley\n\na\n\na\n\nCenter of section\n\nAngle of wrinkle\n\nTest lengths\n\nFigure 13.- Summary of test stations.\n\nMeasuring bridge\n\nPressure by rubber string\n\nPointer\n\nTest station panel 3\n\nRivet 1\n\nPointer\n\nSection A-B\n\nRoller bearing\n\nClamping ring\n\nPanel 1\n\n2\n\n3\n\n4\n\n5\n\n6\n\nExperimental cylinder\n\nLoading ring\n\nPointer\n\nKnife edge bearing\n\nFigure 14.- Mechanical bridge.\n\nN. A. C. A. Technical Memorandum No. 878\n\nFigs. 13, 14.", "timestamp": "2026-07-19T18:31:47.659036+00:00"}
{"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 49, "total_pages": 50, "image_filename": "19930094543_p49.jpg", "text": "N.A.C.A. Technical Memorandum No. 873\n\n[Figure: Diagram 1 - optimum advance - hot spot not heated (temperature 530°)]\n\nFigure 22.- Diagram 1 - optimum advance - hot spot not heated (temperature 530°).\n\n[Figure: Diagram 2 advance 0° - hot-spot temperature: 500° for lower, 970° for upper diagram; no auto-ignition.]\n\nFigure 23.- Diagram 2 advance 0° - hot-spot temperature: 500° for lower, 970° for upper diagram; no auto-ignition.\n\nFigs. 22,23", "timestamp": "2026-07-19T18:31:49.788657+00:00"}
{"citation_id": "19930091714", "source_url": "https://ntrs.nasa.gov/api/citations/19930091714/downloads/19930091714.pdf", "page_number": 36, "total_pages": 36, "image_filename": "19930091714_p36.jpg", "text": "FEB 25 1974", "timestamp": "2026-07-19T18:31:56.154241+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 4, "total_pages": 18, "image_filename": "19930091701_p4.jpg", "text": "# NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nHEADQUARTERS, NAVY BUILDING, WASHINGTON, D. C.  \nLABORATORIES, LANGLEY FIELD, VA.\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 to 15 by act approved March 2, 1929. The members are appointed by the President, and serve as such without compensation.\n\nJOSEPH S. AMES, Ph. D., Chairman,  \nBaltimore, Md.\n\nDAVID W. TAYLOR, D. Eng., Vice Chairman,  \nWashington, D. C.\n\nWILLIS RAY GREGG, Sc. D., Chairman, Executive Committee,  \nChief, United States Weather Bureau.\n\nWILLIAM P. MACCRACKEN, J. D., Vice Chairman, Executive Committee,  \nWashington, D. C.\n\nCHARLES G. ABBOT, Sc. D.,  \nSecretary, Smithsonian Institution.\n\nLYMAN J. BRIGGS, Ph. D.,  \nDirector, National Bureau of Standards.\n\nARTHUR B. COOK, Rear Admiral, United States Navy,  \nChief, Bureau of Aeronautics, Navy Department.\n\nFRED D. FAGG, JR., J. D.,  \nDirector of Air Commerce, Department of Commerce.\n\nHARRY F. GUGGENHEIM, M. A.,  \nPort Washington, Long Island, N. Y.\n\nSYDNEY M. KRAUS, Captain, United States Navy,  \nBureau of Aeronautics, Navy Department.\n\nCHARLES A. LINDBERGH, LL. D.,  \nNew York City.\n\nAUGUSTINE W. ROBINS, Brigadier General, United States Army,  \nChief Matériel Division, Air Corps, Wright Field, Dayton, Ohio.\n\nEDWARD P. WARNER, M. S.,  \nGreenwich, Conn.\n\nOSCAR WESTOVER, Major General, United States Army,  \nChief of Air Corps, War Department.\n\nORVILLE WRIGHT, Sc. D.,  \nDayton, Ohio.\n\nGEORGE W. LEWIS, Director of Aeronautical Research\n\nJOHN F. VICTORY, Secretary\n\nHENRY J. E. REID, Engineer-in-Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.\n\nJOHN J. IDE, Technical Assistant in Europe, Paris, France\n\n## TECHNICAL COMMITTEES\n\n| AERODYNAMICS               | AIRCRAFT STRUCTURES         |\n|----------------------------|-----------------------------|\n| POWER PLANTS FOR AIRCRAFT  | AIRCRAFT ACCIDENTS          |\n| AIRCRAFT MATERIALS         | INVENTIONS AND DESIGNS      |\n\nCoordination of Research Needs of Military and Civil Aviation  \nPreparation of Research Programs  \nAllocation of Problems  \nPrevention of Duplication  \nConsideration of Inventions\n\n## LANGLEY MEMORIAL AERONAUTICAL LABORATORY  \nLANGLEY FIELD, VA.\n\nUnified conduct, for all agencies, of scientific research on the fundamental problems of flight.\n\n## OFFICE OF AERONAUTICAL INTELLIGENCE  \nWASHINGTON, D. C.\n\nCollection, classification, compilation, and dissemination of scientific and technical information on aeronautics.", "timestamp": "2026-07-19T18:32:08.541067+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 14, "total_pages": 18, "image_filename": "19930094551_p14.jpg", "text": "N.A.C.A. Technical Memorandum No. 865\nFigs. 1,2,2a\n\n[Figure: Diagram of a Heinkel He 70 aircraft with annotations $H_1$, $H_2$, $N_1$, $N_2$, $a$, and $x_2$ showing an experimental arrangement for ascertaining depth and trail.]\n\nFigure 1.- Heinkel He 70 with static pressure head for high speeds (up to 400 km/h.). Experimental arrangement for ascertaining the depth and trail.\n\n[Figure: Side view of a static pressure head with annotations $N$, $H$, $N_2$.]\n\nFigure 2.- Static pressure head for high speeds; total length; 625 mm; weight 6.5 kg; range up to 400 km/h; tubing diameter: 7 mm; H: thrust; N-nipple, $N_2$- pressure head nipple.\n\n[Figure: Side view of a low-speed pressure head.]\n\nFigure 2a.- Low-speed pressure head; total length: 450 mm; weight: 2.4 kg; range; to 250 km/h.", "timestamp": "2026-07-19T18:32:11.743145+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 20, "total_pages": 47, "image_filename": "19930093641_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:32:14.406671+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 48, "total_pages": 51, "image_filename": "19930094533_p48.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\nFigs. 35,36,37\n\n[Figure: Diagram of a wing cross-section with measurement points and a scale indicator reading \"Scale 0,2°\"]\n\nFigure 35.- Temperature record at -6,1° incidence. (wing)\n\n[Figure: Diagram of a wing cross-section with measurement points and a scale indicator reading \"Scale 0,2°\"]\n\nFigure 36.- Temperature record at -3,1° incidence. (wing)\n\n[Figure: Diagram of a wing cross-section with measurement points and a scale indicator reading \"Scale 0,2°\"]\n\nFigure 37.- Temperature record at 0° incidence. (wing)", "timestamp": "2026-07-19T18:32:15.143057+00:00"}
{"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 19, "total_pages": 20, "image_filename": "19930091692_p19.jpg", "text": "Positive directions of axes and angles (forces and moments) are shown by arrows\n\n| Axis | | | Moment about axis | | | Angle | | Velocities | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Designation | Symbol | Force (parallel to axis) symbol | Designation | Symbol | Positive direction | Designation | Symbol | Linear (component along axis) | Angular |\n| Longitudinal<br>Lateral<br>Normal | $X$<br>$Y$<br>$Z$ | $X$<br>$Y$<br>$Z$ | Rolling<br>Pitching<br>Yawing | $L$<br>$M$<br>$N$ | $Y \\rightarrow Z$<br>$Z \\rightarrow X$<br>$X \\rightarrow Y$ | Roll<br>Pitch<br>Yaw | $\\phi$<br>$\\theta$<br>$\\psi$ | $u$<br>$v$<br>$w$ | $p$<br>$q$<br>$r$ |\n\nAbsolute coefficients of moment\n$C_l = \\frac{L}{qbS}$ (rolling)\n$C_m = \\frac{M}{qcS}$ (pitching)\n$C_n = \\frac{N}{qbS}$ (yawing)\n\nAngle of set of control surface (relative to neutral position), $\\delta$. (Indicate surface by proper subscript.)\n\n**4. PROPELLER SYMBOLS**\n\n$D$, Diameter\n$p$, Geometric pitch\n$p/D$, Pitch ratio\n$V$, Inflow velocity\n$V_s$, Slipstream velocity\n$T$, Thrust, absolute coefficient $C_T = \\frac{T}{\\rho n^2 D^4}$\n$Q$, Torque, absolute coefficient $C_Q = \\frac{Q}{\\rho n^2 D^5}$\n\n$P$, Power, absolute coefficient $C_P = \\frac{P}{\\rho n^3 D^5}$\n$C_s$, Speed-power coefficient $= \\sqrt[5]{\\frac{\\rho V^4}{P n^2}}$\n$\\eta$, Efficiency\n$n$, Revolutions per second, r.p.s.\n$\\Phi$, Effective helix angle $= \\tan^{-1} \\left( \\frac{V}{2\\pi r n} \\right)$\n\n**5. NUMERICAL RELATIONS**\n\n1 hp. = 76.04 kg-m/s = 550 ft-lb./sec.\n1 metric horsepower = 1.0132 hp.\n1 m.p.h. = 0.4470 m.p.s.\n1 m.p.s. = 2.2369 m.p.h.\n\n1 lb. = 0.4536 kg.\n1 kg = 2.2046 lb.\n1 mi. = 1,609.35 m = 5,280 ft.\n1 m = 3.2808 ft.", "timestamp": "2026-07-19T18:32:22.784171+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 6, "total_pages": 28, "image_filename": "19930091697_p6.jpg", "text": "2\nREPORT NO. 622—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nThe fuel quantity is regulated by changing the length of the plunger travel. During most of the tests, a seven-orifice injection nozzle (fig. 2) was used. This nozzle is similar to the one used by Schey and Young in the work reported in reference 5. The injection valve was mounted opposite the intake valves so that the fuel was injected counter to the intake air flow. Injection was timed to start 20° after top center on the intake stroke and the injection period was about 120°.\n\nThe engine-jacket temperature was maintained at 250° F. throughout the tests. The compression ratio was 7.0. The injection started at 20° A. T. C. on the intake stroke and, for the full-load fuel quantity, lasted straight-line relationship for the mixtures having been assumed.\n\nThe pressure indicator has been described in reference 6. A steel blank, which fits into the window opening, is used for mounting the indicator directly in the combustion-chamber wall.\n\nThree different types of photograph were taken: (1) High-speed 16 mm motion pictures; (2) streak schlieren photographs using a continuous light source and moving film, recording the combustion travel along a narrow slit across the chamber; and (3) spark schlieren motion pictures. The last method uses the same film-drum camera as the second, but the light for the schlieren photo-\n\n<!-- Image (224, 274, 839, 703) -->\n\nFIGURE 1.—Diagrammatic sketch of combustion apparatus.\n\nfor about 0.015 second. For the tests at an engine speed of 500 r. p. m., the spark advance was 30° B. T. C. and, at 1,500 r. p. m., it was 20° B. T. C. The earlier spark was used at the lower speed to increase the tendency to knock.\n\nFour fuels having different octane ratings were used in the tests: A commercial iso-octane with an octane number of 100, containing approximately 90 percent 2, 2, 4 trimethyl pentane; aviation gasoline to Army Specification Y-3557-6, having an octane number of 87; ordinary automobile gasoline having an octane number of about 65; and a special fuel having an octane number of 18.\n\nIn addition, blends of the 18-octane fuel with the 65-octane gasoline were used, having estimated octane ratings of approximately 30, 40, and 50, a graphs is furnished by a series of spark discharges and the entire window is photographed.\n\nThe high-speed motion-picture camera (reference 7) was mounted above the combustion apparatus with the camera lens on the center line of the window and parallel to the plane of the window. The N. A. C. A. spark-photography apparatus has been described in reference 8.\n\nFor these tests, the apparatus was used in conjunction with the schlieren optical arrangement (reference 9), by which slight differences in index of refraction of a gaseous medium may be made visible or photographed. The differences may be caused by air flow, by waves traveling through the medium, or by temperature differences.", "timestamp": "2026-07-19T18:32:25.855705+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 7, "total_pages": 13, "image_filename": "19930091693_p7.jpg", "text": "```markdown\nTHE MAXIMUM LIFT OF AN AIRPLANE\n3\n\n<!-- Image (106, 73, 870, 226) -->\n\nFIGURE 4.—Scale effect, Fairchild 22 airplane. Variation of maximum lift coefficient with air speed. Horizontal tail removed.\n\n<!-- Image (106, 256, 870, 463) -->\n\nA. From V. D. T. tests of N. A. C. A. 2R12 airfoil (plotted against equivalent Reynolds Number)\nFIGURE 5.—Scale effect, Fairchild 22 airplane. Variation with air speed of $C_{Lmax}$ at maximum lift and of the maximum lift coefficient, for three test conditions.\n\n<!-- Image (106, 488, 481, 703) -->\n\nFIGURE 6.—Propeller effect on maximum lift. Fairchild 22 airplane; $i_h = -1.6^\\circ$; $i_h = -25^\\circ$.\n\n<!-- Image (498, 488, 870, 703) -->\n\nFIGURE 7.—Comparison of maximum-lift determinations made 6 months apart. Fairchild 22 airplane; $i_h = -1.6^\\circ$; $i_h = -25^\\circ$.\n\nSix months after the completion of the wind-tunnel tests, some of the measurements were repeated in order to test for a suspected deterioration of the wing. The repeat tests failed to show any appreciable effect of deterioration, the results being in satisfactory agreement with the earlier measurements (fig. 7).\n\nThe scale effect on minimum drag is shown in figure 8. The minimum drag coefficient decreases from 0.058 to 0.042 as the speed increases from 30 to 119 miles per hour. This decrease in the drag coefficient is many times greater than that to be expected from the wing alone and may be attributed to a large scale effect on the junctures, struts, and smaller parts of the airplane.\n\nThe effects of scale on the pitching-moment coefficient and on the angle of attack for trim are shown in figure 9 to be negligible, except where reduction in air speed causes stalling, as at the first point on the $14^\\circ$ curve. The elevator angle for trim at maximum lift is plotted against air speed in figure 10; the variation is due to the scale effect on maximum lift.\n\n36572—38—2\n```", "timestamp": "2026-07-19T18:32:31.005992+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 20, "total_pages": 24, "image_filename": "19930091716_p20.jpg", "text": "16 REPORT NO. 641—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nOwing to the reduced velocity in the slipstream of the braking propeller, $q_1 < q_0$ so that the upsetting-moment and the righting-moment slopes, as expressed in the foregoing formulas, cannot be directly compared.\n\nThe ratios of these two values of $q$ may be obtained from the following slipstream-velocity formula:\n\n$$\n\\frac{q_1}{q_0} = \\frac{w^2}{V^2} = 1 + 2.545T_c\n$$\n\nwhere $w$ is the air velocity in the slipstream and $V$ is the velocity of the airplane, which in this example is 285 miles per hour.\n\nFor this airplane, then,\n\n$$\n\\frac{q_1}{q_0} = 1 + (2.545 \\times -0.106) = 0.73\n$$\n\nTherefore\n\n$$\n\\left( \\frac{dN}{d\\psi} \\right)_r = 7.9 \\times 0.73 \\times q_0 = 5.77q_0\n$$\n\nand the resulting slope of the yawing-moment curve with the propeller operating as a brake is\n\n$$\n\\frac{dN}{d\\psi} = \\left( \\frac{dN}{d\\psi} \\right)_r + \\left( \\frac{dN}{d\\psi} \\right)_a = 5.77q_0 - 2.4q_0 = 3.37q_0\n$$\n\nand the airplane is found to be stable by a good margin. From this brief analysis, one should say that little trouble from directional instability will be encountered in a dive where the propeller is being used as a brake unless the airplane was originally designed with little lateral stability.\n\nThis conclusion seems to agree with the evidence obtained in the dive tests reported in reference 2, where no instability was noted even when the propeller was producing its maximum braking effect. This evidence, however, must not be taken as entirely conclusive, for it represents only two specific examples. It is conceivable that conditions of airplane weight, propeller-braking effect, and the basic stability of the airplane could be such that trouble from directional instability might be difficult to avoid.\n\nGLIDE CONTROL AND REDUCTION IN LANDING RUN\n\nThere are a number of situations in the flight range of every airplane where an air brake could be used to advantage. With the great increase in functional flexibility given to the propeller through recently acquired control mechanism, the propeller now provides an ideal air brake; if the propeller is set at negative blade angles and engine power is applied, it becomes a powerful, though nicely controlled, power brake, which need not rely on the speed of the airplane for braking power.\n\nOne situation in which a power brake could be used to advantage on an airplane is in the glide to a landing and also during the landing ground run. In these situations it is conceivable that the use of the propellers as brakes would find its best application in multiengine airplanes. Consider, for example, the new 4-engine transport and bombing airplanes. The landing distance of such airplanes is excessive and only a certain few fields throughout the country are large enough to accommodate them. It is entirely possible that their landing distances (glide over 50-foot obstacle + ground run) could be reduced one-third to one-half by the use of propeller braking power. The blade angles of the two outboard propellers could be set to negative values of 15° or 20° before the landing glide started and the blade angles of the inboard propellers left in their normal take-off position so that in an emergency the airplane could fly off again on these two engines. By a differential use of the throttles for the inboard (thrust-producing) and outboard (braking) engines a very nice control of the glide path could be obtained, thus making spot landings possible and making the best use of landing-field size.\n\nOnce on the ground the pilot could open his braking engines wide and reduce the ground run by a large amount. The sum of both of these maneuvers would reduce the landing distance by 25 to 50 percent. There is a possibility, of course, that the reduced velocity in the slipstream of the braking propellers might have a bad effect on the wing lift or on the cooling of the engines.\n\nThe following table presents the results of landing-run calculations that were made to show the effect on the total distance to land over a 50-foot obstacle of using the propellers as power brakes. The example airplane was a fictitious 4-engine transport having a gross weight of 32,000 pounds and a landing speed of 100 feet per second. Calculations were made for two conditions of landing, as follows: (1) where the propellers were producing no thrust or drag during the glide and ground run; and (2) where the four propellers were producing thrust as indicated in figure 18, i. e., the inboard propellers producing zero thrust and the outboard propellers 2,000 pounds negative thrust each. The value of $L/D=8$ was assumed for landing, with flaps and landing gear down and no propeller thrust.\n\n| Case | $\\gamma$ | $l_1$ (ft.) | $l_2$ (ft.) | $l_1+l_2$ (ft.) | Ratio |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| 1 | 7° 8' | 400 | 1,900 | 2,300 | 1.00 |\n| 2 | 14° 2' | 200 | 1,260 | 1,460 | .63 |\n\nIn the foregoing table $\\gamma$ is the gliding angle; $l_1$ is the distance from the obstacle to the point of contact with the ground; and $l_2$ is the ground run. In the last column, the ratio of the total distance to the total distance for case 1 is shown. It is observed that in this case the glide angle has been nearly doubled and the total landing distance reduced by 37 percent by the use of propellers as power brakes. If all four propellers had been used as brakes, the landing distance could have been reduced still more.", "timestamp": "2026-07-19T18:32:48.765575+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 21, "total_pages": 21, "image_filename": "19930094534_p21.jpg", "text": "N.A.C.A. Technical Memorandum No. 882\nFigs.15 to 26\n\nHammer side for\nhand operation.\n\nHammer side for\nmachine riveting\nInsertion Dimp- Driving\nling\n\nFigure 15.- Mushroom-head\nriveting.\n\nFigure 16.- Mushroom riveting tool.\n(machine operation).\n\nFigure 17.- Dolly with elastic\nmass and curved set\nfor hand driving.\n\nHammer side\nInsertion Dimp- Driving\nling\n\nFigure 18.- Flat countersunk riveting.\n\nSection Section\nA-B C-D\n\nFigures 19-21.- Flat countersunk\nfor limited accessibility.\n\nAir hammer\n(machine only)\n\nFigures 22,23.- Countersunk riveting\nwith flat counter-\nsunk rivets.\n\nHand\nhammer\n\nAir hammer\n(machine riveting)\nDrill Insert Dimple Drive\nCounter-\nsink\n\nFigures 24,25.- Countersunk\nriveting with roundhead rivets.\n\nHammer\nside\n\nInsertion Tight- Up- flat\nening set- ham-\nting mer-\ning\n\nFigure 26.- Riveting with forced-\nin sheets.", "timestamp": "2026-07-19T18:32:56.595658+00:00"}