PDF/ocr_dataset_105.jsonl

{"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 50, "total_pages": 50, "image_filename": "19930094543_p50.jpg", "text": "N.A.C.A. Technical Memorandum No. 873\n\n[Figure: Diagram 3 - advance 0°; hot-spot temperature; T₁ 1015° for lower, T₂ 1045° for upper diagram.]\n\nFigure 24.- Diagram 3 - advance 0°; hot-spot temperature; T₁ 1015° for lower, T₂ 1045° for upper diagram.\n\n[Figure: Diagram 4 - advance 0°; hot-spot temperature; 1190; very advanced auto-ignition.]\n\nFigure 25.- Diagram 4 - advance 0°; hot-spot temperature; 1190; very advanced auto-ignition.\n\nFigs. 24,25", "timestamp": "2026-07-19T18:33:07.716650+00:00"}
{"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 20, "total_pages": 20, "image_filename": "19930091692_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:33:08.923671+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 49, "total_pages": 51, "image_filename": "19930094533_p49.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\nFigs. 38,39,40\n\n[Figure: Diagram of an airfoil with temperature measurement points indicated by circles and lines. A scale bar labeled \"Scale 0,2°\" is present to the right.]\n\nFigure 38.- Temperature record at 3,1° incidence. (wing)\n\n[Figure: Diagram of an airfoil with temperature measurement points indicated by circles and lines. A scale bar labeled \"Scale 0,2°\" is present to the right.]\n\nFigure 39.- Temperature record at 6,1° incidence. (wing)\n\n[Figure: Diagram of an airfoil with temperature measurement points indicated by circles and lines. A scale bar labeled \"Scale 0,2°\" is present to the right.]\n\nFigure 40.- Temperature record at 9,1° incidence. (wing)", "timestamp": "2026-07-19T18:33:15.920074+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 31, "total_pages": 32, "image_filename": "19930094552_p31.jpg", "text": "N.A.C.A. Technical Memorandum No. 864\nFig. 21\n\nPeripherial stresses $\\sigma_y$\n$\\frac{kg}{cm^2}$\n+20\n0\n-20\n10 11 12 13 14 15 16 17 18\nCylinder perimeter\nLongitudinal stiffeners\nLoading by arching forces,\nbulkheads e and f riveted.\nCross section\nEB\nED\nEa\nYB\nYf\n36\n\nPeripherial stresses $\\sigma_y$\n$\\frac{kg}{cm^2}$\n+120\n+100\n+80\n+60\n+40\n+20\n0\n-20\n-40\n-60\n-80\n-100\n-120\n1/18 2/17 3/16 4/15 5/14 6/13 7/12 8/11 9/10\nCylinder perimeter\nCross section\nEB\nED\nYB\nYf\n36\nLongitudinal stiffeners\nLoading by bending forces\n\nPeripherial stresses $\\sigma_y$\n$\\frac{kg}{cm^2}$\n+20\n0\n-20\n36 f YB a. 72 YD YB YD\nStrips\nE'\nD'\nC'\nB'\nA'\nA\nB\nC\nD\nE\nPeripherial stresses $\\sigma_y$\n$\\frac{kg}{cm^2}$\n+120\n+100\n+80\n+60\n+40\n+20\n0\n-20\n36 f YB 72 YD YB YD\nThe peripheral stresses in the strips A B C D E\nare obtained by reflection of the curves A'B'C'D'E'\nat x axis\n\nPeripherial stresses $\\sigma_y$\n$\\frac{kg}{cm^2}$\nE\n10\nD\n11\nC\n12\nB\n13\nA\n14\nA'\n15\nB'\n16\nC'\n17\nD'\n18\nE'\n1\n2\n3\n4\n5\n6\n7\n8\n9\nCross Section\nEB\nED\nYB\n\nFigure 21.- Plot of the\ntransverse\nstresses for P= 1,000 kg\n(for the bending loading\ncondition and for the\narching loading condition\nwith attached bulkheads)\n\nPeripherial stresses $\\sigma_y$\n$\\frac{kg}{cm^2}$\nE\n10\nD\n11\nC\n12\nB\n13\nA\n14\nA'\n15\nB'\n16\nC'\n17\nD'\n18\nE'\n1\n2\n3\n4\n5\n6\n7\n8\n9\nCross section\n36\n72", "timestamp": "2026-07-19T18:33:23.385201+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 7, "total_pages": 28, "image_filename": "19930091697_p7.jpg", "text": "A PHOTOGRAPHIC STUDY OF COMBUSTION AND KNOCK IN A SPARK-IGNITION ENGINE 3\n\nDuring most of the tests in which the schlieren method was used, the film-drum camera of the spark-photography apparatus was used in conjunction with a high-intensity arc light. The optical arrangement is shown diagrammatically in figure 3. In this set-up, light from the arc is brought to a focus on the round hole in the metal plate by the first lens. This round hole, placed at the principal focus of the second lens, serves as the source of light. Light passing through the second lens is rendered parallel and is directed into the combustion chamber by the mirror and reflected back slightly offset from its original path by the mirror on the piston. The settings of the mirrors are such that the light is brought back through the second lens and to a focus slightly to one side of the original source. Thus it is possible to insert a small mirror in the optical path just before the light comes to a focus and to reflect the entire beam at a right angle. The round stop to obtain the schlieren effect is placed in the plane of the image, allowing only an annular ring of light to pass. The film drum of the camera is placed at the image of the combustion chamber formed by the second lens. A stop with a ⅛-inch slit is placed in front of the film drum, so that the image of only a ⅛-inch strip across the combustion chamber is photographed. An electromagnetic shutter is synchronized with the engine, so as to expose the film for only the part of the cycle that is of interest.\n\nWhen sparks are used as the light source for the schlieren pictures, the round-hole light source is replaced by a horizontal spark gap enclosed in a glass tube to confine the spark to a straight path, and the stop to obtain the schlieren effect consists of a slit in a plate. The condensers and distributor of the spark-photography apparatus give 13 sparks at a rate of about 1,000 per second.\n\nWith the schlieren set-up, most of the light from the combustion flame is eliminated in the optical train and does not register on the film. The combustion front is accompanied by a marked temperature increase and is therefore recorded on the film. Sound waves or compression waves, which are accompanied by local changes in density, also are shown by the schlieren method. A ½-inch space on the side of the chamber nearest the usual position of the spark plug was not covered by the mirror on the piston.\n\n[Figure 2: The seven-orifice nozzle. Dimensions labeled: .007\", .006\", .010\", .007\", 1/2\", 15/16\", 12 1/2°]\n\n[Figure 3: Diagrammatic sketch of schlieren set-up. Components labeled: Camera, Stop with annular opening, Mirror, Lens, Round hole in plate, Lens, Arc light, Mirror on piston]\n\nRESULTS\n\nThe high-speed motion pictures reproduced in figure 4 show the effect of air-fuel ratio on nonknocking flame propagation with spark plugs at E and F (fig. 1). These tests were made at an engine speed of 1,500 r. p. m. The irregular flame fronts are characteristic of all the flame photographs that have been taken. The rate of flame propagation was somewhat slower for the lean mixtures, and two distinctly different types of afterburning were present in the rich and lean mixtures.", "timestamp": "2026-07-19T18:33:32.734355+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 8, "total_pages": 13, "image_filename": "19930091693_p8.jpg", "text": "4\nREPORT NO. 618—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (111, 72, 888, 244) -->\n\nFIGURE 8.—Scale effect, Fairchild 22 airplane. Variation of $C_D$ and $C_{D_{min}}$ with air speed. Horizontal tail removed.\n\n**Trim lift curves.**—In order to obtain lift curves corresponding to flight, the wind-tunnel results were adjusted to the trim condition at all angles of attack. Trim lift curves were determined from the wind-tunnel data for two positions of the center of gravity, corre-\n\n<!-- Image (111, 338, 497, 511) -->\n\nFIGURE 9.—Scale effect, Fairchild 22 airplane. Variation with air speed of angle of attack for trim and of pitching-moment coefficient at two angles of attack. Weight, 1,613 pounds; $i_w = -1.6^\\circ$; $i_h = -25^\\circ$.\n\nsponding to the two airplane loadings used in the flight tests. The plots used in deriving the trim lift curves are shown in figures 11, 12, and 13. In figure 11 the effect of elevator deflection on the pitching-moment coefficient is shown. At each elevator setting, the angle of\n\n<!-- Image (111, 582, 497, 755) -->\n\nFIGURE 10.—Scale effect, Fairchild 22 airplane. Variation with air speed of elevator setting for trim at maximum lift. $i_w = -3.6^\\circ$.\n\nattack for trim is found where the curve crosses the axis. Cross plots of these data are shown in figure 12 where the elevator angle for trim is plotted against angle of attack for both airplane weights and for two stabilizer settings. In figure 13 the variation of lift coefficient with elevator angle is shown for several\n\nangles of attack. From figures 12 and 13 the trim lift curves are constructed (fig. 14) for a constant speed of 56 miles per hour.\n\nThese trim lift curves, however, do not actually represent the conditions that would be found in flight tests, for in flight (1) the air speed varies with the lift coefficient, thus giving rise to a small scale effect, and (2) the changing of the angle of attack has an effect,\n\n<!-- Image (546, 383, 888, 728) -->\n\nFIGURE 11.—Pitching-moment curves at different elevator angles. Fairchild 22 airplane; weight, 1,613 pounds; $i_w = -3.6^\\circ$; air speed, 56 m. p. h.\n\napparent mainly as increased maximum lift. Both of these corrections were applied in the construction of the \"flight-speed\" curves of figure 14. The variation with Reynolds Number of the maximum lift coefficient, determined at $d\\alpha/dt=0.1^\\circ$ per second and corrected to the trim condition, is compared with flight results in figure 15.\n\n**The effect of angular velocity on maximum lift.**—During the course of the investigation it was observed, as previously mentioned, that the maximum lift coefficients obtained in the tests with changing angle of", "timestamp": "2026-07-19T18:33:33.764625+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 5, "total_pages": 18, "image_filename": "19930091701_p5.jpg", "text": "REPORT No. 626\n\nTHE TRANSITION PHASE IN THE TAKE-OFF OF AN AIRPLANE\n\nBy J. W. WETMORE\n\nSUMMARY\n\nAn investigation was undertaken to determine the character and importance of the transition phase between the ground run and steady climb in the take-off of an airplane and the effects of various factors on this phase and on the air-borne part of the take-off as a whole. The information was obtained from a series of step-by-step integrations, which defined the motion of the airplane during the transition and which were based on data derived from actual take-off tests of a Verville AT airplane. Both normal and zoom take-offs under several loading and take-off speed conditions were considered. The effects of a moderate wind with a corresponding wind gradient and the effect of proximity of the ground were also investigated.\n\nThe results show that, for normal take-offs, the best transition was realized at the lowest possible take-off speed. Moreover, this speed gave the shortest over-all take-off distance for normal take-offs. Zoom take-offs required a shorter over-all take-off run than normal take-offs, particularly with a heavy loading, if the obstacle to be cleared was sufficiently high, e. g., greater than 50 feet; no advantage was indicated for the airplane with a light loading if the height to be cleared was less. The error that would result from the neglect of the transition in the calculation of the air-borne distance of take-off was found to vary from 4 percent with the heaviest loading considered to -4 percent with the lightest loading for normal take-offs over a 100-foot obstacle; the percentage error was twice as great for a 50-foot obstacle. For zoom take-offs the error attained much greater values. The average wind gradient corresponding to a 5-mile-per-hour surface wind reduced the air-borne distance required to clear a 50-foot obstacle by about 9 percent with the lightest loading and 16 percent with the heaviest loading; for a 100-foot obstacle, the reduction was about 10 percent in both cases. The over-all reduction due to this wind was approximately twice that resulting from the wind gradient alone. A simple expression for the reduction of observed take-off performance to no-wind conditions is presented. Ground effect is shown to reduce the air-borne distance to attain a height of 50 feet by 10 percent with the lightest loading and 16 percent with the heaviest loading; for a 100-foot obstacle, the percentage reduction was about one-half as great.\n\nINTRODUCTION\n\nIn the process of taking off, the course of an airplane consists of three phases: a run along the ground to attain flying speed, a transition curve in which the flight path changes from the horizontal direction of the ground run to an inclination suitable for climbing, and a more or less steady climb to a height at which any obstacles at the edge of the airport will be surmounted. The motion of an airplane in the ground-run and steady-climb stages is relatively simple and therefore can be predicted for prescribed conditions with reasonable accuracy, presupposing an adequate knowledge of the airplane characteristics. The transition, on the other hand, can be accurately defined only by very complex relations; hence, common practice in calculating take-off performance has been to regard this phase as negligible or to account for it with approximations of uncertain validity.\n\nThe investigation described herein was undertaken to provide an indication of the character and relative importance of the transition and of the effects of various factors on the transition itself and on the air-borne portion of the take-off as a whole. For this purpose a series of take-off tests was conducted with a conventional biplane. The tests included both normal take-offs, wherein the air speed was maintained as nearly constant as possible from the instant of leaving the ground, and zoom take-offs, in which the speed was reduced after leaving the ground. The test conditions for each type of take-off covered two loadings and several take-off speeds. The motion of the airplane in the take-offs was measured with a recording phototheodolite.\n\nThe results of these tests were not used directly, as originally intended, inasmuch as they were found to be confused by rather wide variations in piloting procedure and wind condition. Instead, the force relations pertaining to the airplane under take-off conditions were derived from data provided by the tests and served as the basis for a series of step-by-step integrations whereby the motion of the airplane during take-off was determined for various conditions without the effects of piloting and wind. The calculations covered the range of loading and speed conditions included by the actual", "timestamp": "2026-07-19T18:33:34.079558+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 13, "total_pages": 16, "image_filename": "19930094564_p13.jpg", "text": "```markdown\n12 N.A.C.A. Technical Memorandum No. 852\n\nREFERENCES\n\n1. Kramer, M.: The 5- by 7-Meter Wind Tunnel of the DVL. T.M. No. 788, N.A.C.A., 1936.\n\n2. Jacobs, Eastman N., Ward, Kenneth E., and Pinkerton, Robert M.: The Characteristics of 78 Related Airfoil Sections from Tests in the Variable-Density Wind Tunnel. T.R. No. 460, N.A.C.A., 1933.\n\n3. Jacobs, Eastman N. and Pinkerton, Robert M.: Tests of N.A.C.A. Airfoils in the Variable-Density Wind Tunnel. Series 230. T.N. No. 567, N.A.C.A., 1936.\n\n4. Hoerner, S.: Tests of Spheres with Reference to Reynolds Number, Turbulence, and Surface Roughness. T.M. No. 777, N.A.C.A., 1935.\n\n5. Platt, R. C.: Turbulence Factors of N.A.C.A. Wind Tunnels as Determined by Sphere Tests. T.R. No. 558, N.A.C.A., 1936.\n\n6. Doetsch, H.: Profilwiderstandsmessungen im grossen Windkanal der DVL. Luftfahrtforschung, vol. 14, no. 4/5, 1937, and DVL Jahrbuch, 1937.\n\n7. Millikan, Clark B. and Klein, A. L.: The Effect of Turbulence. An Investigation of Maximum Lift Coefficient and Turbulence in Wind Tunnels and in Flight. Aircraft Engineering, August 1933, pp.169-74.\n\n8. Stack, John: Tests in the Variable Density Wind Tunnel to Investigate the Effects of Scale and Turbulence on Airfoil Characteristics. T.N. No. 364, N.A.C.A., 1931.\n\n9. Pearson, H. A.: A Method of Estimating the Aerodynamic Effects of Ordinary and Split Flaps of Airfoils Similar to the Clark Y. T.N. No. 571, N.A.C.A., 1936.\n```", "timestamp": "2026-07-19T18:33:38.986609+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 21, "total_pages": 47, "image_filename": "19930093641_p21.jpg", "text": "L-456.\n\nTABLE I. COMPARISON OF PRINCIPAL AERODYNAMIC CHARACTERISTICS OF MODEL WITH DIFFERENT MOTOR-PROPELLER ARRANGEMENTS\n\n| Airplane | $C_{D_{min}}$ | $C_D$ at $C_L = 0.25$ | $C_{L_{max}}$ $\\delta_f = 0^\\circ$ | $C_{L_{max}}$ $\\delta_f = 60^\\circ$ | $(L/D)_{max}$ | Maximum propulsive efficiency$^2$ $C_L = 0.25$ | Maximum propulsive efficiency$^2$ $C_L = 0.70$ | Maximum over-all efficiency$^3$ $C_L = 0.25$ | Maximum over-all efficiency$^3$ $C_L = 0.70$ |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Wing alone | 0.0088 | 0.0098 | 1.26 | 1.77 | 24.5 | - | - | - | - |\n| Bare wing | 0.0164 | 0.0173 | 1.29 | - | 19.8 | 1.00 | 1.00 | 1.00 | 1.00 |\n| | 0.0155 | 0.0168 | | | | | | | |\n| Conventional nacelle tractor with radiators | .0208 | .0223 | 1.16 | $^5$1.69 | 16.6 | .78 | .75 | .60 | .68 |\n| Conventional nacelle tractor without radiators | .0179 | .0188 | - | - | - | - | - | $^6$.72 | - |\n| Pusher; spinners on | .0155 | .0168 | 1.34 | 1.82 | 19.6 | .80 | .80 | .79 | .77 |\n| Pusher; spinners removed | .0155 | .0168 | 1.34 | 1.79 | 19.4 | .79 | .80 | .79 | .77 |\n| Tractor position 1; diameter of extension-shaft housing, 4 in. | .0158 | .0172 | 1.28 | 1.77 | 19.0 | .79 | .83 | .78 | .78 |\n| Tractor position 2; diameter of extension-shaft housing, 4 in. | .0157 | .0170 | 1.27 | 1.75 | 19.0 | .79 | .79 | .78 | .74 |\n| Tractor position 3; diameter of extension-shaft housing, 4 in. | .0162 | .0175 | 1.30 | 1.74 | 18.8 | .79 | .77 | .75 | .72 |\n| Tractor position 2; diameter of cowling, 8 in. | .0161 | .0177 | 1.30 | - | 19.3 | .79 | .80 | .76 | .75 |\n| Tractor position 3; diameter of cowling, 8 in. | .0163 | .0175 | 1.30 | 1.75 | 18.7 | .78 | .78 | .76 | .72 |\n| Tractor position 3; diameter of cowling, 8 in.; spinners removed | .0163 | .0177 | - | - | - | .78 | .77 | .76 | .72 |\n\n$^1$Drag coefficients given are for 100 m.p.h. tunnel speed.\n$^2$Blade angle, 18-1/2$^\\circ$.\n$^3$Reference value for conventional nacelle tractor.\n$^4$Reference value for enclosed-engine arrangements.\n$^5$Landing gear extended; all others, landing gear retracted.\n$^6$Based on propulsive efficiency from tests with radiators.\n\n18", "timestamp": "2026-07-19T18:33:41.214311+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 11, "total_pages": 16, "image_filename": "19930094559_p11.jpg", "text": "| | | Figure 3 | Figure 4 | Figure 5 | Figure 6 |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| 1. Start of injection compressor crank angle | $^\\circ$KW | 330 | 340 | 350 | 357 |\n| 2. Air velocity in connecting passage | m/s | m $\\approx$ 63 | m = m$_{\\text{max}}$ $\\approx$ 97 | m $\\approx$ 67 | m $\\approx$ 35 |\n| 3. Pressure | atm. abs. | p $\\approx$ 11.5 | p $\\approx$ 14.5 | p $\\approx$ 17.5 | p $\\approx$ 19 |\n| 4. Temperature | $^\\circ$C | t = 800 | t $\\approx$ 790 | t $\\approx$ 780 | t $\\approx$ 780 |\n\n| | | Figure 13 | Figure 14 | Figure 15 | Figure 16 |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| 1. Fuel quantity | mg | B $\\approx$ 35 | B $\\approx$ 60 | B $\\approx$ 35 | B $\\approx$ 60 |\n| 2. Start of injection compressor crank angle | $^\\circ$KW | 340 | | 357 | |\n| 3. Air velocity | m/s | m = m$_{\\text{max}}$ $\\approx$ 190 | | m $\\approx$ 25 | |\n| 4. Pressure | atm. abs. | p $\\approx$ 23.5 | | p $\\approx$ 26.5 | |\n| 5. Temperature | $^\\circ$C | t $\\approx$ 550 | | t $\\approx$ 550 | |\n\n(These tables are part of figures 3, 4, 5, 6 and 13, 14, 15, 16.)\n\nN.A.C.A. Technical Memorandum No. 857\n\n9", "timestamp": "2026-07-19T18:33:41.872977+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 28, "total_pages": 43, "image_filename": "19930094544_p28.jpg", "text": "26 N.A.C.A. Technical Memorandum No. 872\n\n11. Engberding: Die englischen Luftschiffneubauten R 100 und R 101 (The New English Airships R 100 and R 101). Z.F.M., vol. 19, no. 9, 1928, pp. 194-198, and no. 10, p. 219.\n\n12. Wallis, E. N.: Rigid Airship Design and Construction (R 100). Aircraft Engineering, vol. 2, no. 11, 1930, pp. 7-10.\n\n13. Richmond, V. C.: \"R 101.\" Journal of the Royal Aeronautical Society, vol. 33, no. 224, 1929, pp. 686-723.\n\n14. Luftwacht: Das englische Luftschiff R101 (The English Airship R 101), vol. 3, no. 11, 1929, pp. 507-512; Das englische Luftschiff R 100 (The English Airship R 100), vol. 4, no. 3, 1930, pp. 137-140; Das Luftschiff Goodyear-Zeppelin \"Akron\" (The Goodyear-Zeppelin Airship \"Akron\"), vol. 5, no. 11, 1931, pp. 506-512.\n\n15. Arnstein, K.: Über einige Luftschiffprobleme. (Concerning Some Airship Problems). Z.F.M., vol. 25, no. 1, 1932, pp. 1-13.\n\n16. Ebner, H.: Das amerikanische Starrluftschiff \"Akron\" (The American Rigid Airship \"Akron\"), Z. VDI, vol. 76, no. 2, 1932, pp. 37-40.\n\n17. Arnstein, K.: The Development of Large Commercial Airships. Trans., Amer. Soc. Mech. Eng., vol. 8, 1928, pp. 1-14.\n\n18. Schwengler, J.: Das Grossluftschiff der Gegenwart und seine tatsächlichen Leistungen (The Large Airship of the Present and Its Actual Performance). Yearbook of the Wissenschaftliche Gesellschaft für Luftfahrt, 1927, pp. 65-90.\n\n19. Erfordernisse und Anregung für wirtschaftliche Luftschiffbauten (Requisites and Reasons for Economical Airship Construction). Z.F.M., vol. 23, no. 14, 1932, pp. 419-422.\n\n20. Fritsche, C. B.: Some Economic Aspects of the Rigid Airship. Aeronautical Engineering, Trans., A.S.M.E., vol. 3, no. 1, 1931, pp. 25-40.", "timestamp": "2026-07-19T18:33:45.600680+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 15, "total_pages": 18, "image_filename": "19930094551_p15.jpg", "text": "N.A.C.A. Technical Memorandum No. 465\nFigs. 3,4,5\n\n<!-- Image (292, 152, 727, 322) -->\nFigure 3.- Relation of depth to flying speed with different suspension lengths.\n\n<!-- Image (283, 417, 727, 641) -->\nFigure 4.- Relation of trail to flying speed with different suspension lengths.\n\n<!-- Image (301, 666, 587, 883) -->\nFigure 5.- Tensile stress P and direction $K_3$ of suspension tubing at airplane body against flying speed with 16 m suspension length.", "timestamp": "2026-07-19T18:33:51.300338+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 18, "total_pages": 76, "image_filename": "19930094549_p18.jpg", "text": "16 N.A.C.A. Technical Memorandum No. 867\n\nturbance in altitude $\\delta\\theta = -0.2$ radian or $-11.46^\\circ$, and a disturbance in the angle of attack $\\delta i = +0.2$ radian, corresponding to $\\delta w = -0.2$ V = $-8$ m/s (18 m.p.h.).\n\nThe numerical values found for $\\rho$, $\\varphi$, etc. are given in table I, and the diagram giving the disturbances $\\delta u$ of the velocity $u$, (which may be assumed to be the same as V), $\\delta\\theta$ of the inclination $\\theta$, and $\\delta i$ of the angle of attack $i$, is shown on figure 5. These curves, as well as the figures of the numerical table, show how the initial disturbance affecting the angle of attack and the attitude, is distributed between the two oscillations.*\n\nThe airplane is suddenly raised as a result of the excess of lift and tends to nose down if stable - these two phenomena decreasing the angle of attack.\n\nThe most stable airplane noses down energetically by the action of the rapid oscillation. After a fraction of a second the disturbance of the angle of attack is practically annulled by the joint effect of this diving action and the vertical acceleration which turns the flight path upward.\n\nThe strong curvature of the $\\delta\\theta$ curve corresponds to the vanishing of the rapid oscillations. At the instant when these vanish the airplane is at an angle of attack differing little from the normal angle of attack but stalled by $6^\\circ$ and on a rising flight path. A lack of equilibrium will be felt in the forces and the long-period oscillation will arise from this fact.\n\nThe airplane only slightly stable likewise noses down through the effect of the short-period oscillation but the action is less accentuated. The flight path meanwhile curves upward. At the instant when the rapid oscillation ceases to be felt, the airplane has nosed down only $2^\\circ$, but its flight path has been raised and the airplane will be found on a rising flight path of about $9^\\circ$, the disturbance in the angle of attack always remaining small. The nonequilibrium of the forces is greater than in the preceding case, and the resulting long-period oscillation will be of greater amplitude.\n\n---\n\n*On this figure, as well as on those following, the negative values of $\\delta\\theta$ are plotted above the x-axis so as to facilitate reading the diagrams. Upwardly inclined motions will then correspond to rising curves and vice versa.", "timestamp": "2026-07-19T18:34:06.092921+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 32, "total_pages": 32, "image_filename": "19930094552_p32.jpg", "text": "N.A.C.A. Technical Memorandum No. 864\nFig. 22\n\nPeripheral stresses\n$\\sigma_y$\n$kg/cm^2$\n+400\n+300\n+200\n+100\n0\n-100\n\nStrips A'\nB'\nC'\nD'\n\n36 I I' II$\\alpha$ II$\\beta$ I$\\beta$ I$\\beta$\n\nThe peripheral stresses in the strips A B C D\nare obtained by reflection of the curves A'B'C'D'\n\n[Figure: Polar plot of peripheral stresses $\\sigma_y$ with cross section labels]\n\nFigure 22.- Plot\nof\nthe transverse\nstresses for\nP=1,000 kg\n(for the arching\nloading condi-\ntion with un-\nattached bulk-\nheads).\n\nCross section\n$V\\beta'$\n$V\\alpha'$\n$I\\beta$\n\nPeripheral stresses $\\sigma_y$\n$kg/cm^2$\n+400\n+300\n+200\n+100\n0\n-100\n-200\n-300\n-400\n\nCross section\n$I\\beta$\n$II\\beta$\n$I\\alpha'$\n$I\\gamma'$\n$I\\gamma$\n36\n\nCylinder perimeter\n-y\n\n10 11 a b 12 13 14 15 a b 16 17 18\nLongitudinal stiffener.", "timestamp": "2026-07-19T18:34:12.918321+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 11, "total_pages": 22, "image_filename": "19930091655_p11.jpg", "text": "HEAT TRANSFER TO FUEL SPRAYS INJECTED INTO HEATED GASES 7\n\nengines. The whole A–C interval, on the contrary, is of no immediate interest in this respect and corresponds to the period within which heat is being abstracted from the gas phase at a rate greater than the rate of transfer from the bomb wall. The A–C interval is partial pressure of the vaporized fuel. Records representing this condition were not obtained as the relatively low rate of heat transfer from the bomb wall would have necessitated an extended deflection period for the diaphragm.\n\n[Figure: Graph showing pressure drop (atmosphere) vs. time (second) for five different gas temperatures (150°C, 200°C, 250°C, 300°C, 350°C). Each subplot shows curves labeled A, B, C with numerical values (46, 230, 197, 408, 441) near point A. Vertical axis labeled “PRESSURE DROP, ATMOSPHERE” on left; horizontal axis labeled “TIME, SECOND” at bottom; right vertical axis labeled “GAS TEMPERATURE, °C.”]\n\n(b) Gas density, 14.19 grams per liter; gas-fuel ratio, 30.\n\nFIGURE 4.—Continued. Variation of pressure drop with gas temperature. Diesel fuel; fuel weight, 0.0284 gram.\n\ninfluenced so little by most of the available variables that no theoretical basis for its approximate constancy is at present evident. Eventually, upon reestablishing thermal equilibrium, the pressure should increase beyond its initial value to an extent represented by the Spray photographs shown in figure 8 illustrate the manner in which sprays from the 13-orifice and the 2-impinging-jets nozzles penetrate air at room temperature and a density of 14.19 grams per liter for an intermediate injection pressure.\n\n103475—37——2", "timestamp": "2026-07-19T18:34:16.037181+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 80, "total_pages": 102, "image_filename": "19930094542_p80.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:34:16.285937+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 50, "total_pages": 51, "image_filename": "19930094533_p50.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\nFigs. 41,42\n\n[Figure: Diagram of a wing cross-section with multiple measurement points indicated by circles and lines. A scale bar labeled \"Scale\" and \"0.2°\" is shown on the right side.]\n\nFigure 41.-- Temperature record at 12.2° incidence(wing).\n\n[Figure: Diagram of a wing cross-section with multiple measurement points indicated by circles and lines. A scale bar labeled \"Scale\" and \"0.2°\" is shown on the right side.]\n\nFigure 42.-- Temperature record at 16.4° incidence(wing).", "timestamp": "2026-07-19T18:34:20.198497+00:00"}
{"citation_id": "19930094568", "source_url": "https://ntrs.nasa.gov/api/citations/19930094568/downloads/19930094568.pdf", "page_number": 1, "total_pages": 28, "image_filename": "19930094568_p1.jpg", "text": "848\n\nTECHNICAL MEMORANDUMS\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nNo. 848\n\nPLANING-SURFACE TESTS AT LARGE FROUDE\nNUMBERS - AIRFOIL COMPARISON\n\nBy A. Sambrus\n\nLuftfahrtforschung\nVol. 13, No. 8, August 20, 1936\nVerlag von R. Oldenbourg, München und Berlin\n\nWashington\nFebruary 1938\n\n[Stamp: FAIRCHILD AIRCRAFT CORPORATION HAGERSTOWN, MD.]", "timestamp": "2026-07-19T18:34:20.538653+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 39, "total_pages": 43, "image_filename": "19930094538_p39.jpg", "text": "```markdown\nN. A. C. A. Technical Memorandum No. 873\n\n36 cm\nx\n0\n$\\sigma_y$\n-100\n-200\n$kg/cm^2$\n-300\nBulkhead\ncenter\nBulkhead\ncenter\nCylinder III\n$\\left\\{ \\begin{array}{l} \\circ \\text{ Test points in panel II} \\\\ \\circ \\text{ \" \" \" III} \\\\ \\circ \\text{ \" \" \" IV} \\end{array} \\right.$\nComputed for complete tension field\n\" \" incomplete \" \"\n\" \"(Ebner - Heck method)\n\ny\n0\n$\\sigma_y$\n-100\n$kg/cm^2$\n-200\n-300\nTest points\nCenter\nof section 7\nCenter\nof section 8\nCenter\nof section 9\nCenter\nof section 10\n140\nCylinder III\nExperimental value\nTheoretical value complete and\nincomplete tension field.\nTheoretical value\n(Ebner - Heck method)\n\n36 cm\nx\n0\n$\\sigma_y$\n-100\n-200\n$kg/cm^2$\n-300\n-400\n-500\n-600\nBulkhead\ncenter\nBulkhead\ncenter\nCylinder IV\n$\\left\\{ \\begin{array}{l} \\circ \\text{ Test points in panel II} \\\\ \\circ \\text{ \" \" \" III} \\\\ \\circ \\text{ \" \" \" IV} \\end{array} \\right.$\nComputed for complete tension field\n\" \" incomplete \" \"\n\" \"(Ebner - Heck method)\n\ny\n0\n$\\sigma_y$\n-100\n$kg/cm^2$\n-200\nCenter\nof section 5\nCenter\nof section 6\nTest points\nCenter\nof section 7\nCenter\nof section 8\n210\nCylinder IV\nMean value from experiments\nTheoretical value complete and\nincomplete tension field.\nTheoretical value\n(Ebner - Heck method)\n\nFigure 17.- Compressive stresses $\\sigma_x$ in the stringers.\nFigure 18.- Compressive stresses $\\sigma_y$ in the bulkheads.\n\nFigs. 17,18\n```", "timestamp": "2026-07-19T18:34:24.566931+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 9, "total_pages": 13, "image_filename": "19930091693_p9.jpg", "text": "THE MAXIMUM LIFT OF AN AIRPLANE 5\n\nattack were considerably higher than those obtained in the standard tests. In figure 16 are given lift curves showing the manner in which the peak of the lift curve rises with increasing rate of change of angle. For the standard tests the maximum value is 1.405, whereas, with the angle changing at the rate of 0.2° per second, it is 1.480. In figure 17 the maximum lift coefficients\n\nin the figure; the variation parallels that of the lift coefficient.\n\nIn order to establish the validity of these results, particularly as regards the possibility of error due to balance characteristics, it was ascertained that, on the one hand, the damping was too low and, on the other hand, the natural frequency of the balance was too\n\n<!-- Image (179, 197, 406, 524) -->\n\nFIGURE 12.—Elevator settings for trim. Fairchild 22 airplane; air speed, 56 m. p. h.\n\n<!-- Image (588, 197, 815, 524) -->\n\nFIGURE 14.—Trim lift curves. Fairchild 22 airplane.\n\nfor two different airplane conditions are plotted against the nondimensional parameter $c \\frac{d\\alpha}{V dt}$, in which $c$ and $V$ are the chord and the velocity, respectively. The upper curve is for the tail-removed condition, while the lower curve represents a tail-on condition with the elevator set approximately for trim at maximum lift. The angle of attack at maximum lift is also plotted\n\nhigh to cause any appreciable discrepancy between the indicated and the actual forces. As a further check on the work, a small airfoil of N. A. C. A. 2R,12 section was tested at corresponding rates of change of angle of attack in the N. A. C. A. variable-density tunnel. The results of these tests were in very good agreement with the results just discussed.\n\n<!-- Image (108, 658, 575, 908) -->\n\nFIGURE 13.—Effect of tail setting on lift. Fairchild 22 airplane; air speed, 56 m. p. h.\n\n<!-- Image (609, 658, 876, 908) -->\n\nFIGURE 15.—Comparison of wind-tunnel with flight determination of the maximum lift coefficient. Variation with air speed of the maximum lift coefficient at trim. Fairchild 22 airplane.", "timestamp": "2026-07-19T18:34:24.769099+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 8, "total_pages": 28, "image_filename": "19930091697_p8.jpg", "text": "4\nREPORT NO. 622—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nDetails of the photographs are shown to better advantage in the enlargements in figures 5 and 6. In figure 5, at air-fuel ratios of 10 and 12.5, after the flame apparently passed through part of the chamber, an area of very bright illumination appeared behind the flame front as in the frames marked C, and in the succeeding frames this area spread rapidly across the window. In the last frames of the enlargements this bright area is the most prominent feature of the photograph. These areas, which appear suddenly and spread rapidly, do not behave like other regions of brightness, such as the region in the lower part of the window at an air-fuel ratio of 10 and the other bright spots that appear in many of the pictures. Withrow and Rassweiler observed a somewhat similar effect in their tests (reference 2) and attributed it to burning lubricating oil. This explanation appeared reasonable in their case because the brightness appeared as soon as the flame reached the edge of the cylinder. The piston and cylinder of the N. A. C. A. combustion apparatus, however, are lubricated by graphite and the apparatus has no oil in the crankcase. Hence, some other explanation must be sought. The effect appeared only at the richer mixtures and is probably associated with incomplete combustion. For the ratios of 18.5 and 21.6 the flame completely crossed the chamber; then, after the charge had apparently been burned, “afterburning” began and continued until long after the exhaust valves opened. To the eye the exhaust appeared a brilliant violet color.\n\nEnlargements of the burning at the air-fuel ratio of 18.5 are shown in figure 6, and it can be seen that the flame had traversed the chamber by 30° after top center. The afterburning, as shown in the figure, always originates near the spray nozzle and is probably caused by the sudden addition of a small amount of fuel to the hot combustion gases, which still contain oxygen. Records of the pressure in the injection system indicated that a secondary injection of fuel might occur more than one crankshaft revolution after the start of the main spray. One other possible source of fuel is the well between the nozzle seat and the spray orifices. Approximately 0.003 gram of fuel is trapped in this well. It is not known whether such a small quantity of fuel vaporizing into the combustion gases could cause the intense illumination shown in the figures. Further tests are being conducted on the effects of air-fuel ratio.\n\nIn most of the photographs, small local areas of brighter illumination appear throughout the flame. These areas are not believed to be caused by uneven distribution of the fuel inasmuch as 100-octane fuel, differing little in volatility from the other fuels, burned with very uniform illumination.\n\n[Figure: Effect of air-fuel ratio on flame propagation. White marks on upper edge of each film correspond to 75°, C, and 90° A. T. C. Engine speed, 1,500 r. p. m.; two spark plugs; 8°-octane fuel.]", "timestamp": "2026-07-19T18:34:27.538196+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 6, "total_pages": 18, "image_filename": "19930091701_p6.jpg", "text": "tests, and an additional loading condition was also considered.\n\nA measure of the effect of ground proximity on the airplane characteristics was obtained from the test data and, with this information, the influence of ground effect on the take-off was investigated for each of two loading conditions. For the same conditions the effects of a wind increasing in velocity with altitude were also evaluated.\n\nAPPARATUS\n\nA Verville AT airplane (fig. 1) was used for the take-off tests. The pertinent characteristics of this airplane are given in table I. The following standard N. A. C. A. recording instruments were mounted in the airplane: an air-speed recorder; an accelerometer located near the center of gravity and recording accelerations along the normal, or Z, axis of the airplane; an inclinometer recording the direction of the resultant of the external forces imposed on the airplane; a recording engine tachometer; and a control-position recorder connected\n\n[Figure: The Verville AT airplane.]\n\nto the elevators. Half-second intervals of time were recorded by all the instruments from impulses produced by a standard timer.\n\nAn N. A. C. A. recording phototheodolite, essentially a combination of a motion-picture camera and a recording theodolite, provided records from which the horizontal and vertical displacements of the airplane relative to the ground and its attitude angle could be determined at intervals of $\\frac{1}{16}$ second. A timer was also used in conjunction with this instrument.\n\nSynchronization of the phototheodolite records with those of the airplane instruments was accomplished by means of an electrically operated device mounted on the landing gear of the airplane and connected through the instrument switch so that, at the instant the pilot threw the switch to start the instruments, a quantity of white powder was discharged and formed a cloud that was readily discernible in the photographs.\n\nThe wind speed at the ground was measured with an indicating vane anemometer.\n\nTEST PROCEDURE\n\nA series of eight take-offs was made with each of two loading conditions: 2,060 pounds and 2,378 pounds gross weight. For four of the take-offs of each series,\n\nwhich will be designated “normal” take-offs, the pilot was requested to leave the ground at speeds ranging from 3 to 15 miles per hour in excess of the minimum level-flight speed and to climb at the same speeds, attaining steady climbing conditions as quickly as possible. For the four remaining runs, given the designation of “zoom” take-offs, the speeds at the instant of take-off were in the same range but were reduced after the airplane left the ground, the climbs in all cases being made at a speed slightly in excess of the minimum. In all the take-offs the airplane was headed directly into the wind. The engine was operated at full throttle throughout each run.\n\nThe phototheodolite was set up on the ground at a suitable distance from the course of the airplane and recorded its motion during the latter third of the ground run and throughout the transition and climb to a height of about 100 feet. The procedure followed in the operation of the phototheodolite and in the evaluation of the data obtained therefrom was substantially the same as that described for the landing tests of reference 1, although the instrument used for the present tests is of a later and improved design.\n\nCOMPUTATIONS\n\nThe results of the foregoing tests gave evidence of sufficiently great irregularities in the wind conditions and piloting to obscure completely the effects that the tests were expected to disclose; hence, the purpose of the investigation was not directly accomplished by the tests alone. The data obtained from the take-off tests, however, made possible the derivation of the force relations required as the basis for a series of step-by-step integrations defining the motion of the airplane during take-off for various conditions. In this way the troublesome factors of wind and piloting were eliminated.\n\nDerivation of force relations.—Synchronized readings of the data recorded by the airplane instruments and the phototheodolite during the take-offs were made at frequent intervals throughout the records, thus covering a considerable range of flight conditions. Values of lift and excess thrust were obtained for each set of readings according to the following procedure. The normal and longitudinal components of the aerodynamic forces acting on the airplane $F_z$ and $F_x$, respectively, were determined from the relations\n\n$$\nF_z = \\frac{W}{g} a_z\n$$\n\nand\n\n$$\nF_x = \\frac{W}{g} a_x \\tan \\theta\n$$\n\nwhere $W$ is the gross weight of the airplane.\n\n$g$, the acceleration of gravity.\n\n$a_z$, the normal acceleration as recorded by the accelerometer.\n\n$\\theta$, the angle of the inclinometer pendulum relative to the normal axis of the airplane.", "timestamp": "2026-07-19T18:34:37.220855+00:00"}
{"citation_id": "19930091719", "source_url": "https://ntrs.nasa.gov/api/citations/19930091719/downloads/19930091719.pdf", "page_number": 1, "total_pages": 33, "image_filename": "19930091719_p1.jpg", "text": "~~Hatswell Doc.~~\n~~Engines & Props.~~\n\nAERO. & ASTRO. LIBRARY\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nREPORT No. 642\n0.3\n\nMASS. INST. OF TECHNOLOGY\nENGINEERING LIBRARY\n\nTESTS OF FIVE FULL-SCALE PROPELLERS\nIN THE PRESENCE OF A RADIAL AND A LIQUID-\nCOOLED ENGINE NACELLE, INCLUDING TESTS\nOF TWO SPINNERS\n\nBy DAVID BIERMANN and EDWIN P. HARTMAN\n\n[Figure: Seal of the United States]\n\n1938\n\nFor sale by the Superintendent of Documents, Washington, D. C. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-19T18:34:43.222843+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 16, "total_pages": 18, "image_filename": "19930094551_p16.jpg", "text": "N.A.C.A. Technical Memorandum No. 865\nFigs. 6,7,8\n\n0.45\n$c_{\\alpha}$\n0.40\n0.35\n0.30\n0.25\n0.20\n0.15\n0.10\n0.05\n0\n\n1.1\n$c_w$\n1.0\n0.9\n0.8\n0.7\n0.6\n0.5\n0.4\n0.3\n0.2\n0.1\n\n$v=60m \\ s^{-1}$\n\n$c_s=90^{\\circ}$\n$c_s=0^{\\circ}$\n\nWind\nTunnel\n\n$c_w=$\n$1.0 \\sin^2 \\alpha_s$\n\nWind\nTunnel\n\n$c_{\\alpha}=0.86 \\sin^2 \\alpha_s \\cos \\alpha_s$\n\n20 40 60 80\n$\\alpha_s$\n\nFigure 6.- 7 mm diameter.\n\n0.45\n$c_{\\alpha}$\n0.40\n0.35\n0.30\n0.25\n0.20\n0.15\n0.10\n0.05\n0\n\n1.1\n$c_w$\n1.0\n0.9\n0.8\n0.7\n0.6\n0.5\n0.4\n0.3\n0.2\n0.1\n\n$v=65m \\ s^{-1}$\n\nWind\nTunnel\n\n$c_w=$\n$1.05 \\sin^2 \\alpha_s$\n\nWind\nTunnel\n\n$c_{\\alpha}=$\n$1.11 \\sin^2 \\alpha_s \\cos \\alpha_s$\n\n20 40 60 80\n$\\alpha_s$\n\nFigure 7.- 8 mm diameter.\n\nRelation of lift and drag coefficients to angle of incidence for suspension tube element of 400 mm length.\n\n[Figure: Force distribution diagram]\n\n$A_s=C_3 \\sin^2 \\alpha_s \\cos \\alpha_s$\n$\\Delta l$\n$A_n$\n$B_n$\n$W_s=C_2 \\sin^2 \\alpha_s$\n$G_s$\n$B_{n-1}$\n$\\Delta F$\n$A_{n-1}$\n\nFigure 8.- Force distribution on an element of the suspension tubing.", "timestamp": "2026-07-19T18:34:47.455000+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 19, "total_pages": 76, "image_filename": "19930094549_p19.jpg", "text": "N.A.C.A. Technical Memorandum No. 867 17\n\nThe long-period oscillation is very clearly brought out by the disturbances in the velocity V, to which it corresponds. The disturbances in the angle of attack which accompany the long-period oscillation are greater on a less stable than on a very stable machine. The case of an airplane statically neutral requires no remarks. In the case of the statically unstable airplane the disturbance of angle of attack decreases at first as a result of the upward curvature of the path; the latter is not stable, however. The airplane enters a condition of constantly increasing angle of attack, leading inevitably to such disturbances that the method of small motions, after a certain time, ceases to apply.\n\nDISTURBANCE OF ANGULAR VELOCITY\n\nThe disturbance above considered, suddenly increasing the angle of attack and the orientation in space of an important angle, is not actually realized in practice. It may, of course, be imagined that a localized gust strikes the rear of the airplane, but the latter cannot instantly attain the final angle of attack assumed. The airplane will pass through all intermediate angles of attack and the effects of excess lift and of the static stability will make themselves felt during these states.\n\nIn order to analyze the case, let us imagine that the disturbance applied is an impulsive angular velocity tending to turn the nose up $\\delta q < 0$. From the diagrams of V, $\\theta$, i, we have found that for $\\delta q = -0.96$ radian per second (which is a considerable value), the ultimate effects of the disturbance were the same as those of the initial disturbances of attitude and angle of attack studied above. The immediate effects of the disturbance evidently differ. The angle of attack starts from zero and tends to increase but the increase is rapidly stopped and the long-period phenomena are the only ones that subsist after about two seconds. It does not appear necessary to comment any further on the diagrams of figure 6, constructed for the initial disturbance $\\delta q = -0.96$ radian.\n\nDISTURBANCES OF THE SURROUNDING MEDIUM\n\nThe effect of the disturbances in the surrounding medium may be studied by the following device. Let us as-", "timestamp": "2026-07-19T18:34:50.509371+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 22, "total_pages": 47, "image_filename": "19930093641_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:34:59.353458+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 14, "total_pages": 16, "image_filename": "19930094564_p14.jpg", "text": "N.A.C.A. Technical Memorandum No. 852\n\nFigs. 1,6,7,8,9\n\n[Figure: Airfoil of 4 m span and 0.8 m chord in the 5×7 m wind tunnel of the D.V.L.]\n\nFigure 1.- Airfoil of 4 m span and  \n0.8 m chord in the 5×7 m  \nwind tunnel of the D.V.L.\n\n[Figure: Dimensions of experimental split flap.]\n\nFigure 6.- Dimensions of  \nexperimental  \nsplit flap.\n\n[Graph: Wing chord for 100 km/h, landing speed. Figure 7.- cₐₘₐₓ of N.A.C.A. airfoil series 2409 to 2421 with split flap.]\n\nWing chord for 100 km/h, landing speed.  \nFigure 7.- cₐₘₐₓ of N.A.C.A. airfoil  \nseries 2409 to 2421 with  \nsplit flap.\n\n[Graph: Wing chord for 100 km/h, landing speed. Figure 8.- cₐₘₐₓ of N.A.C.A. airfoil series 0009 to 0021 with split flap.]\n\nWing chord for 100 km/h, landing speed.  \nFigure 8.- cₐₘₐₓ of N.A.C.A.  \nairfoil series  \n0009 to 0021 with split flap.\n\n[Graph: Wing chord for 100 km/h, landing speed. Figure 9.- cₐₘₐₓ of N.A.C.A. airfoil series 23009 to 23018 with split flap.]\n\nFigure 9.- cₐₘₐₓ of N.A.C.A.  \nairfoil series  \n23009 to 23018 with split  \nflap.  \nWing chord for 100 km/h, landing speed.", "timestamp": "2026-07-19T18:35:01.104173+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 29, "total_pages": 43, "image_filename": "19930094544_p29.jpg", "text": "N.A.C.A. Technical Memorandum No. 872 27\n\n21. Blakemore, Thos. L., and Fagon, W. Watters: Pressure Airships. The Ronald Press, New York, 1927.\n\n22. Naatz, H.: Die neuesten Fortschritte im Prallluftschiffbau (The Latest Improvements in Pressure Airship Construction). Yearbook of the Wissenschaftliche Gesellschaft fur Luftfahrt, 1929, pp. 105-109.\n\n23. Wiesinger, K.: Das Luftschiff Bauart Wiesinger (The Wiesinger Type Airship). Z.F.M., vol. 21, no. 13, 1930, pp. 335-338.\n\n24. Bleistein, W.: Metallluftschiffe (Metal Airships). Z.F.M., vol. 21, no. 24, 1930, pp. 626-630.\n\n25. Fritsche, C. B.: The Metalclad Airship. Aeronautical Engineering, Trans., A.S.M.E., vol. 1, no. 4, 1929, pp. 245-266. Jour., Royal Aeronautical Society, vol. 35, no. 249, 1931, pp. 818-883.\n\n26. Abraham, M.: Drähte, Seile, und Litzen im Flugzeugbau (Wires, Cables, and Strands in Airplane Construction). DVL Yearbook, 1930, pp. 347-410. DVL Report No. 177.\n\n27. Klemperer, W.: Windkanalversuche an einem Zeppelin-Luftschiff-Modell (Wind Tunnel Tests on a Zeppelin Airship Model). Transactions, Aerodynamische Institut der Technischen Hochschule Aachen, no. 12, pp. 3-56.\n\n28. Freeman, H. B.: Pressure Distribution Measurements on the Hull and Fins of a 1/40-Scale Model of the U.S. Airship \"Akron.\" T.R. No. 443, N.A.C.A., 1933.\n\n29. De France, S. J.: Flight Tests on U.S.S. \"Los Angeles.\" Part I. Full Scale Pressure Distribution Investigation. T.R. No. 324, N.A.C.A., 1929.\nBurgess, C. P.: Flight Tests on U.S.S. \"Los Angeles.\" Part II. Stress and Strength Determination. T.R. No. 325, N.A.C.A., 1929.\n\n30. Fuhrmann, G.: Theoretische und experimentelle Untersuchungen an Ballonmodellen (Theoretical and Experimental Studies of Balloon Models). Z.F.M., vol. 2, 1911, pp. 165-166. Diss. Berlin 1912. Verlag Julius Springer. Yearbook of the Motorluftschiff-Studiengesellschaft, vol. 5, 1911-12, pp. 65-123.", "timestamp": "2026-07-19T18:35:07.864331+00:00"}
{"citation_id": "19930091697", "source_url": "https://ntrs.nasa.gov/api/citations/19930091697/downloads/19930091697.pdf", "page_number": 9, "total_pages": 28, "image_filename": "19930091697_p9.jpg", "text": "A PHOTOGRAPHIC STUDY OF COMBUSTION AND KNOCK IN A SPARK-IGNITION ENGINE\n\n| Air-fuel ratio | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 10.0 | 433 | | | | | | | | | C |\n| 12.5 | 436 | | | | | | | | | C |\n| 14.0 | 441 | | | | | | | | | |\n| 16.0 | 442 | | | | | | | | | |\n\n| | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | T.C. | | 10° | | 20° | | 30° | | A.T.C. | |\n\nFIGURE 5.—Effect of air-fuel ratio on flame propagation. Enlargements of high-speed motion pictures of figure 4 for air-fuel ratios of 10, 12.5, 14, and 16. Engine speed, 1,500 r. p. m.; two spark plugs; 87-octane fuel.\n\n5", "timestamp": "2026-07-19T18:35:13.068605+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 40, "total_pages": 43, "image_filename": "19930094538_p40.jpg", "text": "N. A. C. A. Technical Memorandum No. 879\n\nstringer\nCylinder : III\nSections\na-a\nb-b\ne-e\npanel III\nstringer\nPanel III\na\nb\ne\n\nstringer\nstringer\nArc\nBase line of substitute surface\nChord\n\nstringer\nCylinder : IV\nSections\na-a\nb-b\ne-e\npanel III\nSections\ne-e\nf-f\ng-g\npanel IV\nd-d d. bulkhead: d\nstringer\ne\ng\nd\nIV\ne\n\nstringer\nstringer\nArc\nBase line of substitute surface\nChord\n\nFigure 19.- Cross sections of buckling patterns.\n\nFig. 19", "timestamp": "2026-07-19T18:35:18.440265+00:00"}
{"citation_id": "19930094533", "source_url": "https://ntrs.nasa.gov/api/citations/19930094533/downloads/19930094533.pdf", "page_number": 51, "total_pages": 51, "image_filename": "19930094533_p51.jpg", "text": "N.A.C.A. Technical Memorandum No. 883\nFigs. 43,44\n\n<!-- Image (174, 87, 872, 438) -->\n\nFig. 43.- Comparison of results obtained on model and on wing at pressure tap No. 3.\n\n<!-- Image (174, 504, 889, 917) -->\n\nFig. 44.- Comparison of results obtained with the model and the wing at pressure tap No. 12.", "timestamp": "2026-07-19T18:35:18.730729+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 81, "total_pages": 102, "image_filename": "19930094542_p81.jpg", "text": "N.A.C.A. Technical Memorandum No. 974\nFigs.65,66,67,68\n\n<!-- Image (163, 109, 462, 456) -->\n\nFigure 65. $\\kappa=4^\\circ$\n\n<!-- Image (535, 109, 905, 456) -->\n\nFigure 66. $\\kappa=-1^\\circ$\n\n<!-- Image (163, 529, 462, 876) -->\n\nFigure 67. $\\kappa=-6^\\circ$\n\n<!-- Image (535, 529, 905, 876) -->\n\nFigure 68. $\\lambda=0.13$\n\nMoment curves of wing in presence of propeller.", "timestamp": "2026-07-19T18:35:24.725513+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 12, "total_pages": 16, "image_filename": "19930094559_p12.jpg", "text": "N.A.C.A. Technical Memorandum No. 857\nFigs.1,2,11,12\n\n[Figure: Diagram of a turbulence chamber model with labeled parts a through i]\n\nFigure 1.- Turbulence chamber model built into the bomb.(after Ricardo).\n\na Bomb combustion space.\nb Window cut out in bomb.\nc Discharge valve of air compressor.\nd Air passage from compressor.\nf Connecting passage to turbulence chamber.\ng Turbulence chamber.\nh Fuel valve.\ne Air inlet from compressor.\n\n[Figure: Graph showing Pressure p, temperature t, and velocity w as functions of crank angle]\n\nFigure 2.- Pressure p, temperature t and velocity w of air in connecting passage and temperature of turbulence chamber wall $t_{wall}$ as a function of crank angle of compressor.\n\n[Figure: Diagrams of an air storage model (Henschel-Lanova type) with labeled parts a through l]\n\nFigures 11,12.- Air storage model (Henschel-Lanova type) built into bomb.\n\na Combustion bomb.\nb Window of bomb.\nc Bosch nozzle.\nd Air storage chamber.\ne Main combustion space.\nf 6 orifices on one side for incoming air.\ng I. Throttle location.\nh I. Air cell.\ni II Throttle location.\nk II Air cell.\nl Observation window.", "timestamp": "2026-07-19T18:35:24.914885+00:00"}
{"citation_id": "19930091731", "source_url": "https://ntrs.nasa.gov/api/citations/19930091731/downloads/19930091731.pdf", "page_number": 1, "total_pages": 32, "image_filename": "19930091731_p1.jpg", "text": "NATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nREPORT No. 656\n\nTHE COLUMN STRENGTH OF TWO EXTRUDED\nALUMINUM-ALLOY H-SECTIONS\n\nBy WILLIAM R. OSGOOD and MARSHALL HOLT\n\n[Figure: Seal of the National Advisory Committee for Aeronautics]\n\n1939\n\nFor sale by the Superintendent of Documents, Washington, D. C. - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 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- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -", "timestamp": "2026-07-19T18:35:39.839578+00:00"}
{"citation_id": "19930091693", "source_url": "https://ntrs.nasa.gov/api/citations/19930091693/downloads/19930091693.pdf", "page_number": 10, "total_pages": 13, "image_filename": "19930091693_p10.jpg", "text": "6\nREPORT NO. 618—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (252, 83, 762, 283) -->\n\nFIGURE 16.—Lift curves for different angular velocities. Fairchild 22 airplane; horizontal tail removed; air speed, 56 m. p. h.\n\nThe effect upon the maximum lift of the rate of change of angle of attack is well known, but the magnitude observed here was much higher than had been anticipated on the basis of previous investigations. Thus, Kramer's formula (reference 5), which has been approximately confirmed both at low Reynolds Numbers (reference 6) and in flight (reference 7), predicts only 1/20 of the observed increase. The failure of Kramer's formula in this case may be due to the fact\n\nto investigate the general character of the time variation of the lift at fixed angles of attack in this range, a few tests were made in which the angle was increased at a rate of 0.1° per second up to a certain value and then held constant while observations of lift force were made. The results are shown in figure 18. The highest angle of attack at which the lift is maintained indefinitely is about 16.6°. When the angle of attack is increased above this value and then fixed, the flow breaks down within a few seconds and wide fluctuations occur in the lift.\n\n<!-- Image (151, 457, 466, 657) -->\n\nFIGURE 17.—Variation with $\\frac{c}{V} \\frac{d\\alpha}{dt}$ of $\\alpha_c$ at maximum lift and of the maximum lift coefficient. Fairchild 22 airplane; air speed, 56 m. p. h.\n\nthat the formula was based on experimental results obtained at very high values of $\\frac{c}{V} \\frac{d\\alpha}{dt}$. It is also possible that the phenomenon is not so independent of the wing section characteristics as has been heretofore supposed.\n\nThe influence of air speed on the phenomenon is shown in figure 5, where a set of standard runs at different air speeds is compared with a corresponding set in which the angle of attack was changed at the rate of 0.1° per second.\n\nIt is clear that the value of the lift coefficient in the neighborhood of and beyond the maximum is not uniquely determined by the angle of attack. In order\n\n<!-- Image (516, 457, 927, 657) -->\n\nFIGURE 18.—Decay with time of added lift due to angular velocity, and fluctuations in lift beyond the stall. Angle of attack increased at 0.1° per second up to the values shown, and then fixed. Fairchild 22 airplane; $h_0 = -1.6^\\circ$; $h_v = -25^\\circ$; air speed, 56 m. p. h.\n\nFigure 19 further illustrates the vagaries in the behavior of the lift coefficient near and beyond the angle of maximum lift. Three separate lift curves are shown, obtained under apparently identical conditions. Each is fairly smooth, yet different from the other two.\n\nFLIGHT TESTS\n\nThe flight tests consisted in recording in flight sufficient data to obtain the acceleration normal to the flight path, the angle of attack, and the dynamic pressure while the angle of attack was being slowly increased over a range of several degrees below and", "timestamp": "2026-07-19T18:35:40.430942+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 20, "total_pages": 76, "image_filename": "19930094549_p20.jpg", "text": "18 N.A.C.A. Technical Memorandum No. 867\n\nsume that the ambient air, instead of being undisturbed, is subject to a velocity U', having the components u' and w'. Let U of components u and w be the absolute velocity of the airplane. The relative velocity may be written:\n\n$$V = \\sqrt{(u - u')^2 + (w - w')^2}$$\n\nand the angle of attack as\n\n$$i = - \\frac{w - w'}{u - u'}$$\n\nIt is always possible, in the case of a given airplane, to investigate the motion corresponding to an equilibrium of forces in an atmosphere having any absolute velocity whatever, since the relative velocity and the angle of attack are independent of the disturbance velocity. It is therefore simple to investigate the motions which the airplane would assume during disturbances of the surrounding medium.\n\nLet R be the condition of the airplane for which $u'_1 = 0$ $w'_1 = 0$. This condition corresponds to the values $u_1$ and $w_1$ of the absolute velocity of the airplane. In a disturbed atmosphere $A_2$ for which $u'_2 \\neq 0$ $w'_2 \\neq 0$ the airplane, flying at the same angle of attack and same relative velocity, will be in a state $R_2$ characterized by different absolute velocities $u_2$ and $w_2$ determined by $u_2 = u_1 + u'_2$ $w_2 = w_1 + w'_2$. If the airplane flying in state $R_1$ suddenly passes from the atmosphere $A_1$ to the atmosphere $A_2$, it will cease to be in equilibrium. It will not be able to maintain its state $R_1$ and will tend to assume state $R_2$. The motions of the airplane could be calculated by considering the state $R_2$ as that of equilibrium, and the airplane as deviating from it by\n\n$$\\delta u = - u'_2$$\n$$\\delta w = - w'_2$$\n\nThis artifice enables us to study the effect of sudden gusts on the airplane.", "timestamp": "2026-07-19T18:35:44.304516+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 7, "total_pages": 18, "image_filename": "19930091701_p7.jpg", "text": "THE TRANSITION PHASE IN THE TAKE-OFF OF AN AIRPLANE 3\n\nThe flight-path angle $\\gamma$, referred to wind axes, was given by\n\n$$\n\\gamma = \\sin^{-1} \\frac{V_y}{V}\n$$\n\nwhere $V_y$ is the vertical velocity, determined by differentiation of the time-distance curves derived from the phototheodolite records.\n\n$V$, the air speed along the flight path.\n\nIt was necessary, of course, to assume here that the wind had no vertical component, apparently a reasonable assumption for average conditions according to the information of reference 1.\n\nThe angle of attack $\\alpha$ was then obtained from\n\n$$\n\\alpha = \\lambda - \\gamma\n$$\n\nwhere $\\lambda$ is the attitude angle of the airplane, provided\n\nof two variables, angle of attack and air speed. It would consequently be difficult to plot these data directly. For this reason the effective propeller thrust $T$, shown in figure 3, was calculated by means of the information provided in references 2 and 3. The drag $D$ could then be determined from the equation\n\n$$\nD = T - T_{ex}\n$$\n\nand thence the drag coefficient\n\n$$\nC_D = \\frac{D}{1/2 \\rho S V^2}\n$$\n\nwhich could, of course, also be plotted as a function of angle of attack to establish a suitably faired curve. With the data in this form, the relation of excess thrust\n\n[Figure: Two graphs labeled (a) and (b). Graph (a) shows Lift coefficient, $C_L$ and Drag coefficient, $C_D$ vs Angle of attack, deg. Graph (b) shows Lift coefficient, $C_L$ vs Drag coefficient, $C_D$. Both graphs have data points separated by height > 10 ft and height < 10 ft.]\n\n(a) Variation with angle of attack.\n(b) Polar diagrams.\n\nFIGURE 2.—Lift and drag characteristics of the Verville AT airplane as determined from take-off tests.\n\nby the phototheodolite records. With the foregoing information it was possible to determine values for the lift $L$ and the excess thrust $T_{ex}$ by resolving the forces $F_x$ and $F_z$ along the flight-path axes or\n\n$$\nL = F_z \\cos \\alpha + F_x \\sin \\alpha = \\frac{W}{g} a_z (\\cos \\alpha + \\tan \\theta \\sin \\alpha)\n$$\n\n$$\nT_{ex} = F_x \\cos \\alpha - F_z \\sin \\alpha = \\frac{W}{g} a_x (\\tan \\theta \\cos \\alpha - \\sin \\alpha)\n$$\n\nThe values of lift were converted to the coefficient form $C_L$ with the relation\n\n$$\nC_L = \\frac{L}{1/2 \\rho S V^2}\n$$\n\nThus the data could be readily plotted and faired as a function of angle of attack. (See fig. 2 (a).)\n\nThe full-throttle excess thrust is, in effect, a function\n\nto air speed and lift coefficient was determined by using the faired results in a reversal of the procedure.\n\nIn order to take into account the effect of ground proximity on the lift and drag characteristics, hence on the excess thrust, the data were divided into two groups and were plotted separately, according to whether they were obtained when the wheels of the airplane were above or below a height of 10 feet from the ground. This height was arbitrarily chosen as the line of demarcation between the region of strongest ground effect and the region in which, for the purposes of the present investigation, the ground effect could be considered as negligible. The data available were insufficient to warrant further division.\n\nThe lift and drag coefficients evaluated by the foregoing methods are plotted against angle of attack in figure 2 (a) and as polars in figure 2 (b). In figure 3", "timestamp": "2026-07-19T18:35:45.899702+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 23, "total_pages": 47, "image_filename": "19930093641_p23.jpg", "text": "```markdown\n19\n\nFIGURE LEGENDS\n\nFigure 1.- Installation of the 4-engine model in the full-scale wind tunnel: Bare-wing case.\n\nFigure 2.- Installation of the 4-engine model in the full-scale wind tunnel; Conventional nacelles and external radiators for liquid-cooled engines.\n\nFigure 2(a).- Bottom view - Installation of the 4-engine model in the full-scale wind tunnel: Conventional nacelles and external radiators for liquid-cooled engines.\n\nFigure 3.- Installation of the 4-engine model in full-scale wind tunnel. Four-inch diameter extension shaft housings and pusher-propeller arrangement.\n\nFigure 4.- Installation of the 4-engine model in the full-scale wind tunnel: Four-inch-diameter extension-shaft housings and tractor propellers 0.39c ahead of wing.\n\nFigure 5.- Installation of the 4-engine model in full-scale wind tunnel: Eight-inch-diameter cowls and tractor propellers at 0.26c ahead of wing.\n\nFigure 6.- Installation of the 4-engine model in the full-scale wind tunnel: Four-inch diameter extension shaft housings and tractor propellers at 0.13c ahead of wing.\n\nFigure 7.- Diagram of model.\n\nFigure 8.- Blade dimensions for 3-blade model propellers.\n\nFigure 9.- Aerodynamic characteristics of model. Bare wing, without nacelles or radiators; $\\delta_e$, $0^\\circ$; $\\delta_f$, $0^\\circ$; approximate test air speed, 59 m.p.h.\n\nFigure 10.- Aerodynamic characteristics of model. Wing nacelles and radiators for liquid-cooled engines; $\\delta_e$; $0^\\circ$; approximate test air speed, 59 m.p.h.\n\nFigure 11.- Aerodynamic characteristics of model. Pusher model; housing diameter, 4 inches; spinners on; approximate test air speed, 59 m.p.h.\n```", "timestamp": "2026-07-19T18:35:48.504856+00:00"}
{"citation_id": "19930093281", "source_url": "https://ntrs.nasa.gov/api/citations/19930093281/downloads/19930093281.pdf", "page_number": 1, "total_pages": 22, "image_filename": "19930093281_p1.jpg", "text": "FILE COPY\nNO. 3-W\n\nN 62 65279\n\n→ ACR Dec. 1939\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nWARTIME REPORT\n\nORIGINALLY ISSUED\nDecember 1939 as\nAdvance Confidential Report\n\nTHE EFFECT OF STREAMLINING THE AFTERBODY OF\nAN N.A.C.A. COWLING\n\nBy George W. Stickle, John L. Crigler, and Irven Naiman\n\nLangley Memorial Aeronautical Laboratory\nLangley Field, Va.\n\nFILE COPY\nTo be returned to\nthe files of the National\nAdvisory Committee\nfor Aeronautics\nWashington D. C.\n\n40\n\nNACA\n\nWASHINGTON\n\nNACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of\nadvance research results to an authorized group requiring them for the war effort. They were pre-\nviously held under a security status but are now unclassified. Some of these reports were not tech-\nnically edited. All have been reproduced without change in order to expedite general distribution.\n\nL - 279", "timestamp": "2026-07-19T18:35:58.226244+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 82, "total_pages": 102, "image_filename": "19930094542_p82.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:36:08.242447+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 41, "total_pages": 43, "image_filename": "19930094538_p41.jpg", "text": "W. A. C. A. Mechanical Memorandum No. 878\n\nCylinder II\n0.006\n0.004\n0.002\n0\n10 20 30 10⁴\nT in kg cm\nCylinder I\nCylinder II\nCylinder III\n1st Test\n2nd\nPoint of failure\n\nCylinder III\n0.006\n0.004\n0.002\n0\n10 20 30 10⁴\nT in kg cm\nL = 1050 kg cm\n\nCylinder IV\n0.008\n0.006\n0.004\n0.002\n0\n10 20 30 10⁴\nT in kg cm\nL = 1050 kg cm\nExperimental value\nComputed for complete tension field\nComputed by Ebner & Heck\nComputed for incomplete tension field\n\nFigure 20.- Experimental and theoretical angle of twist $\\psi$.\n\nAngle of principal axis\n80°\n60°\n40°\n20°\n0°\n0 50 100 150 200 250 300 350\nT kg cm\n5. Cylinder I - 1050 kg cm\n6. Cylinder II - 1050 kg cm\nFor incomplete tension field\nFor complete tension field\nAngle of wrinkles\nFor incomplete tension field\nFor complete tension field\nAngle of wrinkles\nCylinder II\nCylinder III\n\nFigure 21.- Angle $\\alpha$ of the principal axes.\n\nFigs. 20, 21", "timestamp": "2026-07-19T18:36:15.105511+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 13, "total_pages": 16, "image_filename": "19930094559_p13.jpg", "text": "N.A.C.A. Technical Memorandum No. 857\nFigs.3,4,5,6\n\n[Figure: A series of sequential images showing fuel spray and combustion patterns in a chamber, labeled \"Figure 3.\"]\nFigure 3.\n\n[Figure: A series of sequential images showing fuel spray and combustion patterns in a chamber, labeled \"Figure 4.\"]\nFigure 4.\n\n[Figure: A series of sequential images showing fuel spray and combustion patterns in a chamber, labeled \"Figure 5.\"]\nFigure 5.\n\n[Figure: A series of sequential images showing fuel spray and combustion patterns in a chamber, labeled \"Figure 6.\"]\nFigure 6.\n\nFigures 3,4,5,6.- Effect of air velocity on fuel distribution\nand combustion. Turbulence chamber model.\nFuel: Derop gasoil fuel quantity B= 50 mg, Bosch pump PE 1 B\n100/100, n=400 r.p.m., Bosch nozzle DN 12 SD 12.\nOpening pressure =85 atm. Seconds per picture =.00245 s.", "timestamp": "2026-07-19T18:36:16.504915+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 17, "total_pages": 18, "image_filename": "19930094551_p17.jpg", "text": "N.A.C.A. Technical Memorandum No. 865\nFigs. 9, 9a\n\n[Figure: A large nomograph with multiple scales and lines. The vertical axis on the left is labeled 'a c/gc_s' with values from 0 to 35. The horizontal axis at the bottom is labeled '2A + G_s / C_s' with values from 1 to 20. There are numerous diagonal lines labeled with angles from 10° to 90°. A formula is shown: 'a = (2A + G_s) / C_s -> sin α_s, cos²α_s + (G_s/C_s) sin²α_s'. To the right is a smaller graph with a vertical axis labeled 'sin α_s cos²α_s + (G_s/C_s) sin²α_s' and a horizontal axis labeled 'sin α_s cos²α_s + (G_s/C_s) sin²α_s'. This smaller graph has three lines labeled 'G_s = 2000', '1000', and '500'. The caption for this entire section is 'Figure 9a.- Supplement to nomograph.']\n\n[Figure: An explanatory sketch showing a geometric construction. It includes axes labeled 'x' and 'y', points 'x_1' and 'y_1', and angles. Formulas shown include 'a = (2A + G_s) / C_s' and 'β = G_s / C_s'. Below the sketch is the text: 'Explanatory sketch to the nomographic method: (see Fig. 9): after finding points x and y the relevant intersecting straight is shifted parallel until the ordinates, for example, at x_1 and Y_1 disclose equal angles.']\n\n[Figure: A nomograph with a vertical axis labeled '2A/C_s' with values from 0 to 25, and a horizontal axis labeled 'β' with values from 0 to 8. There are three diagonal lines labeled 'G_s = 174', 'G_s = 1335', and 'G_s = 8604'. The caption is 'Figure 9.- Nomograph for obtaining the suspension tube curve.']", "timestamp": "2026-07-19T18:36:21.130253+00:00"}
{"citation_id": "19930091719", "source_url": "https://ntrs.nasa.gov/api/citations/19930091719/downloads/19930091719.pdf", "page_number": 2, "total_pages": 33, "image_filename": "19930091719_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... | meter | m | foot (or mile) | ft. (or mi) |\n| Time... | second | s | second (or hour) | sec. (or hr.) |\n| Force... | weight of 1 kilogram | kg | weight of 1 pound | lb. |\n| Power... | horsepower (metric) | k.p.h. | horsepower | hp. |\n| Speed... | meters per second | m.p.s. | miles per hour | m.p.h. |\n| | | | feet per second | f.p.s. |\n\n## 2. GENERAL SYMBOLS\n\n$W$, Weight=$mg$\n$g$, Standard acceleration of gravity=9.80665 m/s$^2$ or 32.1740 ft./sec.$^2$\n$m$, Mass=$\\frac{W}{g}$\n$I$, Moment of inertia=$mk^2$. (Indicate axis of radius of gyration $k$ by proper subscript.)\n$\\mu$, Coefficient of viscosity\n$\\nu$, Kinematic viscosity\n$\\rho$, Density (mass per unit volume)\nStandard density of dry air, 0.12497 kg-m$^{-4}$-s$^2$ at 15$^\\circ$ C. and 760 mm; or 0.002378 lb.-ft.$^{-4}$ sec.$^2$\nSpecific weight of \"standard\" air, 1.2255 kg/m$^3$ or 0.07651 lb./cu. ft.\n\n## 3. AERODYNAMIC SYMBOLS\n\n$S_t$, Area\n$S_w$, Area of wing\n$G$, Gap\n$b$, Span\n$c$, Chord\n$b^2$, Aspect ratio\n$S$,\n$V$, True air speed\n$q$, Dynamic pressure=$\\frac{1}{2}\\rho V^2$\n$L$, Lift, absolute coefficient $C_L=\\frac{L}{qS}$\n$D$, Drag, absolute coefficient $C_D=\\frac{D}{qS}$\n$D_0$, Profile drag, absolute coefficient $C_{D_0}=\\frac{D_0}{qS}$\n$D_i$, Induced drag, absolute coefficient $C_{D_i}=\\frac{D_i}{qS}$\n$D_p$, Parasite drag, absolute coefficient $C_{D_p}=\\frac{D_p}{qS}$\n$C$, Cross-wind force, absolute coefficient $C_C=\\frac{C}{qS}$\n$R$, Resultant force\n$i_w$, Angle of setting of wings (relative to thrust line)\n$i_t$, Angle of stabilizer setting (relative to thrust line)\n$Q$, Resultant moment\n$\\Omega$, Resultant angular velocity\n$\\frac{Vl}{\\mu}$, Reynolds Number, where $l$ is a linear dimension (e.g., for a model airfoil 3 in. chord, 100 m.p.h. normal pressure at 15$^\\circ$ 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$C_p$, Center-of-pressure coefficient (ratio of distance of c.p. from leading edge to chord length)\n$\\alpha$, Angle of attack\n$\\epsilon$, Angle of downwash\n$\\alpha_0$, Angle of attack, infinite aspect ratio\n$\\alpha_i$, Angle of attack, induced\n$\\alpha_a$, Angle of attack, absolute (measured from zero-lift position)\n$\\gamma$, Flight-path angle", "timestamp": "2026-07-19T18:36:22.879030+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 15, "total_pages": 16, "image_filename": "19930094564_p15.jpg", "text": "N.A.C.A. Technical Memorandum No. 852\nFigs. 2,3,4,5\n\n1.6\n1.4\n1.2\n1.0\n.8\n0 .04 .08 .12 .16 .20 .24\n$\\frac{d}{t}$ [wing thickness / wing chord]\nTips blunt\nTips rounded\nFigure 2.- Maximum lift of the N.A.C.A. airfoil series 2409 to 2421 with blunt and round wing tips. $R_{eff} = 3 \\times 10^6$\n\n1.6\n1.4\n1.2\n1.0\n.8\n0 1 2 3 4\nWing chord for 100 km/h landing speed, m\n2409\n2412\n2415\n2416\n2421\n$R_{eff}$\n1 2 3 4 5 6 7 8x10$^5$\nTunnel Source\n• DVL 5x7m\n○ GALCIT 3m (7)\n× VDT 1.5m (2)\nFigure 3.- $c_{a_{max}}$ of N.A.C.A. airfoil series 2409 to 2421\n\n1.4\n1.2\n1.0\n.8\n.6\n0 1 2 3 4\nWing chord for 100 km/h landing speed, m\n0012\n0015\n0018\n0021\n0009\n$R_{eff}$\n1 2 3 4 5 6 7 8x10$^5$\nTunnel Source\n• DVL 5 x 7m\n× VDT 1.5m (2)\nFigure 4.- $c_{a_{max}}$ of N.A.C.A. airfoil series 0009 to 0021\n\n1.6\n1.4\n1.2\n1.0\n0 1 2 3 4\nWing chord for 100 km/h landing speed, m\n23009\n23012\n23018\n$R_{eff}$\n1 2 3 4 5 6 7 8x10$^5$\nTunnel Source\n• DVL 5 x 7m\n× VDT 1.5m (3)\nFigure 5.- $c_{a_{max}}$ of N.A.C.A. airfoil series 23009 to 23018", "timestamp": "2026-07-19T18:36:29.597806+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 30, "total_pages": 43, "image_filename": "19930094544_p30.jpg", "text": "28 N.A.C.A. Technical Memorandum No. 872\n\n31. Von Kármán, Th.: Berechnung der Druckverteilung an Luftschiffkörpern (Determination of the Pressure Distribution on Airship Hulls). Transactions of the Aerodynamische Institut der Technischen Hochschule Aachen, No. 6, pp. 2-17.\n\n32. Munk, Max M.: The Aerodynamic Forces on Airship Hulls. T.R. No. 184, N.A.C.A., 1923.\n\n33. Fulton, G.: Some Features of a Modern Airship (U.S.S. \"Akron\"). Trans., Society of Naval Architects and Marine Engineers, vol. 39, 1931.\n\n34. Cox, H. R.: The External Forces on an Airship Structure with Special Reference to the Requirements of Rigid Airship Design. Jour., Royal Aeronautical Society, vol. 33, no. 225, 1929, pp. 725-811.\n\n35. Aeronautical Research Committee: Report of the Airworthiness of Airships Panel. R. & M. No. 970, British A.R.C., 1924.\n\n36. Seydel, E.: Elastizitätstheorie des starren Luftschiffs. (Elastic Theory of the Rigid Airship). (From the writings by H. Müller-Breslau) Z.F.M., vol. 23, no. 2, 1932, pp. 46-51. Luftfahrtforschung, vol. 9, no. 2, 1931, pp. 57-64.\n\n37. Southwell, R. V.: On the Calculation of Stresses in the Hulls of Rigid Airships. R. & M. No. 1057, British A.R.C., 1927.\n\n38. Burgess, C. P.: Airship Design. The Ronald Press, New York, 1927.\n\n39. Lewitt, E. H.: The Rigid Airship. Pitman & Sons Ltd., London, 1925.", "timestamp": "2026-07-19T18:36:31.764415+00:00"}
{"citation_id": "19930091701", "source_url": "https://ntrs.nasa.gov/api/citations/19930091701/downloads/19930091701.pdf", "page_number": 8, "total_pages": 18, "image_filename": "19930091701_p8.jpg", "text": "4\nREPORT NO. 626—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nthe excess thrust within and outside of the region of principal ground effect is shown as a function of lift coefficient and air speed.\n\nStep-by-step integrations.—At quarter-second intervals throughout the transition phase of the take-off, the vertical acceleration $a_z$ and the horizontal acceleration $a_x$ of the airplane were calculated by successive approximations according to the relations\n\n$$a_z = \\frac{g(L \\cos \\gamma + T_{ex} \\sin \\gamma - W)}{W}$$\n\nand\n\n$$a_x = \\frac{g(T_{ex} \\cos \\gamma - L \\sin \\gamma)}{W}$$\n\nCorresponding velocities were determined from\n\n$$V_x = V_{x_0} + \\frac{0.25(a_{x_0} + a_{x_1})}{2} + \\frac{0.25(a_{x_1} + a_{x_2})}{2} + \\dots + \\frac{0.25(a_{x_{n-1}} + a_{x_n})}{2}$$\n\nand\n\n$$V_z = V_{z_0} + \\frac{0.25(a_{z_0} + a_{z_1})}{2} + \\frac{0.25(a_{z_1} + a_{z_2})}{2} + \\dots + \\frac{0.25(a_{z_{n-1}} + a_{z_n})}{2}$$\n\nVertical and horizontal displacements were similarly determined; the flight-path angle was obtained from\n\n$$\\gamma = \\tan^{-1} \\frac{V_z}{V_x}$$\n\nThe initial values of $a_z$ and $V_x$, i. e., at the instant of leaving the ground, were, of course,\n\n$$a_{z_0} = 0$$\n\nand\n\n$$V_{x_0} = 0$$\n\nThe horizontal speed $V_{x_0}$ at the same instant was the assumed take-off speed and, since at this instant $L=W$, the value of the excess thrust $T_{ex}$ and thence the value of $a_{x_0}$ could be determined. For subsequent intervals the quantities involved in the calculations were determined by the usual methods of successive approximation.\n\nThe course of the lift coefficient in the early part of the transition was prescribed by the assumption that the transition should be of as short duration as possible. This limitation, of course, required that the airplane be pulled up quickly to the angle of attack for maximum lift coefficient, as soon as the desired speed for taking off was attained, and held at this angle as long as possible. The lift coefficient was then reduced in time to prevent the flight-path velocity from decreasing, by reason of the increasing climb angle, below the value designated for the steady climb and to permit the adjustment of the lift coefficient necessary to provide a smooth approach to the steady-climb conditions without exceeding reasonable values for the corresponding rate of change of the angle of attack. Examples of the variation in lift coefficient followed in performing the calculations are shown in figure 4.\n\n<!-- Image (524, 155, 911, 582) -->\n\nFIGURE 3.—Excess-thrust characteristics of the Verville AT airplane.\n\n<!-- Image (524, 600, 911, 768) -->\n\nFIGURE 4.—Examples of assumed variation in lift coefficient during transition. The Verville AT airplane.\n\nThe excess thrust corresponding to the lift coefficient and speed occurring at a particular instant was taken from the curves of figure 3, according to whether the height at that instant was greater or less than 10 feet. In this way allowance was made for the ground effect.", "timestamp": "2026-07-19T18:36:47.757642+00:00"}
{"citation_id": "19930091724", "source_url": "https://ntrs.nasa.gov/api/citations/19930091724/downloads/19930091724.pdf", "page_number": 1, "total_pages": 20, "image_filename": "19930091724_p1.jpg", "text": "NATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nREPORT No. 649\n\nTHE “PACK” METHOD FOR COMPRESSIVE TESTS\nOF THIN SPECIMENS OF MATERIALS USED\nIN THIN-WALL STRUCTURES\n\nBy C. S. AITCHISON and L. B. TUCKERMAN\n\n[Figure: Seal of the United States]\n\n1939\n\nFor sale by the Superintendent of Documents, Washington, D. C. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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"timestamp": "2026-07-19T18:36:51.172874+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 24, "total_pages": 47, "image_filename": "19930093641_p24.jpg", "text": "20\n\nFigure 12.- Aerodynamic characteristics of model. Pusher model; housing diameter, 4 inches; spinners removed; $\\delta_{c}$, $0^{\\circ}$; approximate test air speed, 59 m.p.h.\n\nFigure 13.- Aerodynamic characteristics of model. Tractor position 1; housing diameter, 4 inches; spinners on; $\\delta_{c}$, $0^{\\circ}$; approximate test air speed, 59 m.p.h.\n\nFigure 14.- Aerodynamic characteristics of model. Tractor position 2; housing diameter, 4 inches; spinners on; $\\delta_{c}$, $0^{\\circ}$; approximate test air speed, 59 m.p.h.\n\nFigure 15.- Aerodynamic characteristics of model. Tractor position 2; cowling diameter, 8 inches; spinners on; $\\delta_{c}$, $0^{\\circ}$; $\\delta_{f}$, $0^{\\circ}$; approximate test air speed, 59 m.p.h.\n\nFigure 16.- Aerodynamic characteristics of model. Tractor position 3; housing diameter, 4 inches; spinners on; $\\delta_{f}$, $0^{\\circ}$; approximate test air speed, 59 m.p.h.\n\nFigure 17.- Aerodynamic characteristics of model. Tractor position 3; cowling diameter, 8 inches; spinners on; $\\delta_{c}$, $0^{\\circ}$; approximate test air speed, 59 m.p.h.\n\nFigure 18.- Aerodynamic characteristics of wing alone without fuselage or nacelles. $\\delta_{f}$, $0^{\\circ}$; approximate test air speed, 59 m.p.h. $\\delta/c$, 0.135; $t/c$, 0.083.\n\nFigure 19.- Scale effect on the drag coefficient for the models with wing nacelles and radiators and with the bare wing. $\\delta_{c}$, $0^{\\circ}$; $\\delta_{f}$, $0^{\\circ}$.\n\nFigure 20.- Scale effect on the increments of drag from nacelle and radiators for the model with wing nacelles and radiators.\n\nFigure 21.- Scale effect on the drag coefficient for the pusher model. $\\delta_{c}$, $0^{\\circ}$; $\\delta_{f}$, $0^{\\circ}$.\n\nFigure 22.- Scale effect on the drag coefficient for tractor positions 1, 2, and 3. Diameter of extension-shaft housing, 4 inches; spinners on; $\\delta_{c}$, $0^{\\circ}$; $\\delta_{f}$, $0^{\\circ}$.\n\nFigure 23.- Scale effect on the drag coefficient for tractor positions 2 and 3. Cowling diameter, 8 inches; $\\delta_{c}$, $0^{\\circ}$; $\\delta_{f}$, $0^{\\circ}$.", "timestamp": "2026-07-19T18:36:53.279076+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 18, "total_pages": 18, "image_filename": "19930094551_p18.jpg", "text": "N.A.C.A. Technical Memorandum No. 365\n\nFigs. 10,11,12\n\n[Figure: Graph with axes labeled $G \\sin^2 \\alpha_2$, $C_2$, and angles from $0^\\circ$ to $90^\\circ$]\n\nFigure 10.- Graph for defining the drag of a suspension tube element at different $C_2$ and $\\alpha_s$\n\n[Figure: Graph with axes labeled $G \\sin^2 \\alpha_2 \\cos \\alpha_2$, $C_3$, and angles from $0^\\circ$ to $90^\\circ$]\n\nFigure 11.- Graph for obtaining the lift of a suspension tube element at different $C_3$ and $\\alpha_s$\n\n[Figure: Graph with grid, curves labeled I, II, III, IV, V, and annotations “From graph” and “In flight”, alongside a side-view sketch of an airplane with lines extending from it to the graph]\n\nFigure 12.- Comparison of bearing method and nomographic determination of the position of the air speed head with respect to the airplane at different speeds.", "timestamp": "2026-07-19T18:37:08.282004+00:00"}