PDF/ocr_dataset_1002.jsonl

{"citation_id": "19930085957", "source_url": "https://ntrs.nasa.gov/api/citations/19930085957/downloads/19930085957.pdf", "page_number": 27, "total_pages": 27, "image_filename": "19930085957_p27.jpg", "text": "26\nNACA RM L9E02\n\n1.6\n1.2\n.8\n$C_L$ .4\n0\n0 8 16 24\n$\\alpha$, deg\n\n0\n-.08 $C_m$\n-.16\n\n1.6\n1.2\n.8\n$C_L$ .4\n0\n0 8 16 24\n$\\alpha$, deg\n\n0\n-.08 $C_m$\n-.16\n\n[Legend]\n[ ] Gross flow\n[ ] Rough\n[ ] Intermittent stall\n[ ] Complete stall\n\n$C_L = 1.16$ $\\alpha = 13.3^\\circ$\n$C_L = 1.13$ $\\alpha = 13.2^\\circ$\n\n$C_L = 1.34$ $\\alpha = 17.4^\\circ$\n$C_L = 1.28$ $\\alpha = 17.3^\\circ$\n\n$C_L = 1.37$ $\\alpha = 18.5^\\circ$\n$C_L = 1.31$ $\\alpha = 19.4^\\circ$\n\n[NACA]\n\n$C_L = 1.40$ $\\alpha = 21.5^\\circ$\nLeading edge smooth\n$C_L = 1.28$ $\\alpha = 21.3^\\circ$\nLeading edge rough\n\n(b) Split flaps on; R = $3.0 \\times 10^6$.\nFigure 13.- Concluded.", "timestamp": "2026-07-22T04:12:47.203887+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 34, "total_pages": 59, "image_filename": "19930085936_p34.jpg", "text": "NACA RM No. E9B03\n33\n\nPressure coefficient, $C_p$\n\nAngle of yaw (deg)\n-12\n-6\n0\n6\n12\n\nDistance from tip, x/L\n\n(b) $\\theta = 45^\\circ$ longitudinal plane.\n\nFigure 6. - Continued. Pressure distributions along longitudinal planes at $0^\\circ$ angle of attack for range of yaw angles.", "timestamp": "2026-07-22T04:12:47.519605+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 15, "total_pages": 54, "image_filename": "19930086015_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:12:49.161551+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 20, "total_pages": 67, "image_filename": "19930085965_p20.jpg", "text": "NACA RM E9E06\n19\n\n$$P_1 = \\frac{2l_m(d+w)}{g}$$\n\nwhere w refers to width, as shown in figure 12. Also\n\n$$P_2 = 0.52 \\text{ w}$$\n\n$$P_3 = \\frac{2w}{\\pi} \\log_e \\left(\\frac{g+t}{g}\\right)$$\n\nThe plane ABC is so positioned that the average value of volume 5 remains unchanged. In this manner, volumes 4 and 5 can be combined in one calculation. On a scale drawing, the line BA was found to make an angle of $50^\\circ$ with the horizontal, the angle ABD therefore being $140^\\circ$. Then\n\n$$P_{4+5} = \\left(\\frac{180}{140}\\right) \\frac{w}{\\pi} \\log_e \\left(\\frac{g+h+t}{g+t}\\right) = \\frac{1.28 \\text{ w}}{\\pi} \\log_e \\left(\\frac{g+h+t}{g+t}\\right)$$\n\nand\n\n$$P_6 = 0.52 \\text{ d}$$\n\n$$P_7 = \\frac{2d}{\\pi} \\log_e \\left(\\frac{g+t+h}{g}\\right)$$\n\nVolume 8 is considered a partial quadrant of a spherical shell of thickness t+h. The fraction used as a factor in making the computation is the ratio between 180 and the arithmetic mean of 140 and 90, or\n\n$$P_8 = \\frac{180}{\\left(\\frac{140+90}{2}\\right)} \\frac{t+h}{4} = 0.39(t+h)$$\n\nThe total permeance for one end will be\n\n$$2\\left(\\frac{P_1}{2} + P_2 + P_3 + P_{4+5} + P_6 + P_7 + 2P_8\\right)$$", "timestamp": "2026-07-22T04:12:50.655110+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 17, "total_pages": 36, "image_filename": "19930086003_p17.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9J08\n\n[Figure: Photograph of a model with 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil mounted on the bump.]\n\nNACA\nL-61514\n\nFigure 4.— Photograph of a model with 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil mounted on the bump.\n\nCONFIDENTIAL\n\n15", "timestamp": "2026-07-22T04:12:50.843394+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 16, "total_pages": 32, "image_filename": "19930085982_p16.jpg", "text": "14 NACA RM E5E13\n\nshould therefore be laid out along a developed conical surface that approximates the streamlines rather than along a developed cylindrical surface, as was done in this design. In addition, a much greater allowance must be made for the effect of solidity and stagger on turning than was made in this design.\n\nIn figure 9, an arrow indicates the equivalent angle of attack for the weight flow giving peak efficiency. The high efficiency obtained at this condition indicates that all sections of the blade are operating in a relatively low drag region of the angle-of-attack range. A study of drag and pressure-distribution characteristics of cascades of blades has shown that for a range of angles of attack on either side of that for optimum pressure distribution, efficient blade performance can be obtained. Beyond this range of angle of attack, the blade drag increases rapidly. For a given blade section, the angle of attack and resulting turning angle for optimum performance decreases with decreasing solidity and increasing stagger. Thus, the wide range of turning angles from hub to tip on this compressor can be efficiently obtained with a constant-camber blade over a limited range of weight flow. At the blade-tip section, the slope of the turning-angle curve increases very rapidly as the weight flow is increased beyond 20 pounds per second, which corresponds to an angle of attack of approximately $3^\\circ$. A rapid drop in enthalpy addition accompanied by an increase in losses results as the weight flow increases. At the hub the blade appears to have a tendency to stall at an angle of attack of approximately $17^\\circ$, and although the turning angle increases with increasing angle of attack beyond this value, this increase is probably accompanied by drag increases.\n\nFrom these performance characteristics of the tip and hub sections, it appears that in the range of high-efficiency operation of the compressor the tip section is operating near the low region of angle of attack of the efficient blade operating range, whereas the hub is operating near the high region. Thus, a small variation in weight flow in either direction tends to cause a relatively large increase in losses. By the use of a lower cambered blade section at the tip and a higher cambered blade section at the hub, it would therefore appear possible to flatten the efficiency curve somewhat and to minimize the poor radial distribution of energy addition at the extremes of the flow range.\n\nSUMMARY OF RESULTS\n\nThe following results were obtained from experimental operation of a typical inlet stage of an axial-flow compressor designed on the", "timestamp": "2026-07-22T04:12:57.204282+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 38, "total_pages": 47, "image_filename": "19930085918_p38.jpg", "text": "NACA RM A9D29\n37\n\nPressure coefficient, P\n-44\n-40\n-36\n-32\n-28\n-24\n-20\n-16\n-12\n-8\n-4\n0\n4\n8\n\n-90\n-76\n-72\n-68\n-64\n-60\n-56\n-52\n-48\n-44\n-40\n-36\n-32\n-28\n-24\n-20\n-16\n-12\n-8\n-4\n0\n4\n8\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nSpanwise\nstation, 2y/b\n28.1%\n\n57.4%\n\n85.0%\n\nChordwise station, x/c\n2\n4\n6\n10\n\n(g) $\\alpha=28.8^\\circ$\n\nNACA\n\nFigure 9.—Concluded.", "timestamp": "2026-07-22T04:13:01.915762+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 45, "total_pages": 104, "image_filename": "19930085842_p45.jpg", "text": "NACA RM L9C29\n41\n\n[Figure: NACA LMAL 45070]\n$\\alpha = 25.4^\\circ$; $C_L = 0.91$\n\n[Figure: NACA LMAL 45071]\n$\\alpha = 29.2^\\circ$; $C_L = 0.93$\n\n[Figure: NACA LMAL 45072]\n$\\alpha = 32.2^\\circ$; $C_L = 0.97$\n\n[Figure: NACA LMAL 45073]\n$\\alpha = 35.2^\\circ$; $C_L = 0.95$\n\n(a) Configuration 1.\n\nFigure 15.— Tuft observations on the $\\frac{1}{3}$-scale model of the airplane.\nPropellers removed.\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS.", "timestamp": "2026-07-22T04:13:02.514238+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 31, "total_pages": 50, "image_filename": "19930085952_p31.jpg", "text": "```markdown\n30\nNACA RM L9C24\n\n<!-- Image (79, 126, 846, 760) -->\n\nFigure 10.- Increments of pitching-moment and lift coefficients due to propeller operation for model with all-movable horizontal tail removed. Full-power operation; $\\delta_F = 0^\\circ$.\n```", "timestamp": "2026-07-22T04:13:03.908735+00:00"}
{"citation_id": "19930085843", "source_url": "https://ntrs.nasa.gov/api/citations/19930085843/downloads/19930085843.pdf", "page_number": 63, "total_pages": 65, "image_filename": "19930085843_p63.jpg", "text": "NACA RM L9C31\n61\n\n<!-- Image (258, 123, 847, 823) -->\n\nFigure 21.- A comparison of effectiveness for several Mach numbers as obtained by three different test methods on models of a tailless airplane. Vertical fins on; $C_L = 0$.", "timestamp": "2026-07-22T04:13:08.099235+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 22, "total_pages": 43, "image_filename": "19930085958_p22.jpg", "text": "NACA RM No. L9B11\n21\n\n[Figure: A photograph of a model aircraft mounted in a wind tunnel. The model is viewed from the front. A label in the bottom right corner reads \"NACA L-56820\".]\n\n(a) $0.60\\frac{b}{2}$ drooped-nose flaps deflected $40^\\circ$.\n\n[Figure: A photograph of a model aircraft mounted in a wind tunnel. The model is viewed from the front. A label in the bottom right corner reads \"NACA L-56821\".]\n\n(b) $0.55\\frac{b}{2}$ extensible leading-edge flaps. Horizontal tail in highest position.\n\nFigure 3.- The $40^\\circ$ sweptback wing-fuselage combination mounted for testing in the Langley 19-foot pressure tunnel. Split flaps and upper surface fences on; low-wing position.", "timestamp": "2026-07-22T04:13:08.262021+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 61, "total_pages": 82, "image_filename": "19930085551_p61.jpg", "text": "60\nNACA RM No. L8K30\n\n<!-- Image (246, 122, 650, 338) -->\n\n<!-- Image (288, 401, 650, 523) -->\n\n<!-- Image (292, 566, 653, 824) -->\n\n(a) 120 miles per hour.\nFigure 15.- Sideslip characteristics with asymmetric load.\nC-54D airplane; clean condition; normal rated power; right\nwing-tip gas tank empty.", "timestamp": "2026-07-22T04:13:11.439116+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 5, "total_pages": 20, "image_filename": "19930086060_p5.jpg", "text": "NACA RM L9F02 CONFIDENTIAL 3\n\ndiscussion of the method and of the accuracy of results obtained by use\nof this technique is found in reference 1. Two models were flown for\neach configuration investigated; however, one model (10b) failed to\ngive satisfactory results and the data were omitted from the present\npaper.\n\nRESULTS AND DISCUSSION\n\nIn figure 4 are shown the curves of drag coefficient, based on\nfrontal area, against Mach number for the three tested configurations.\nThe small degree of scatter in the data for identical models of a\nconfiguration is an indication of the reliability that may be placed\nupon the results. The variation of drag coefficient with Mach number\nmay be considered through three ranges of velocity corresponding to\nthe flow field around the body: the subsonic range which is terminated\nat the beginning of the rise of drag coefficient, the transonic range\nover which the drag coefficient rise is marked, and the supersonic\nrange which begins at the end of the rise of drag coefficient. The\nMach numbers which mark the transition between the speed ranges for\nthe experimental curves are only approximate inasmuch as there are not\nsharp demarcations between types of flows.\n\nThe following table lists the approximate experimental transition\nMach numbers for each of the body shapes investigated herein and also\ncompares the transonic-supersonic transition with the shock-wave\nattachment Mach numbers for cones as given by some experimental results\n(reference 2) and by theoretical results from reference 3.\n\n| Position of maximum diameter K (percent) | Transition Mach number | | | |\n|---|---|---|---|---|\n| | Subsonic-transonic | Transonic-supersonic | | |\n| | Experiment (fig. 4) | Experiment (fig. 4) | Experiment (reference 2) | Theory (reference 3) |\n| 20 | 0.8 | 1.5 | 1.68 | 1.96 |\n| 40 | .92 | 1.2 | 1.20 | 1.28 |\n| 60 | .94 | 1.1 | 1.09 | 1.15 |\n\nFrom figure 4 the 60-percent station appears to be the best as\nregards zero-lift drag through the Mach number range of these tests.\nBelow M = 0.8 the position of maximum diameter had no effect.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:13:13.462653+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 95, "total_pages": 122, "image_filename": "19930082090_p95.jpg", "text": "NACA TN No. 1455\n93\n\nVentilating air\n5.0\"\nI.D.\n30\"\n26\"\n7.83\"\nI.D.\nExhaust\ngas\n15\"\n3.25\"\n2.25\"\n24.5\"\n7.83\"\nI.D.\nExhaust\ngas\n58.5\"\n54\"\n5.0\"\nI.D.\no Static-pressure tap\nx Temperature traverse\nNACA\nVentilating air\n\nFigure 46.- Schematic diagram of test setup of heat exchanger L and air\nshroud, showing location of static-pressure and temperature measuring\nstations.", "timestamp": "2026-07-22T04:13:15.562167+00:00"}
{"citation_id": "19930085588", "source_url": "https://ntrs.nasa.gov/api/citations/19930085588/downloads/19930085588.pdf", "page_number": 46, "total_pages": 51, "image_filename": "19930085588_p46.jpg", "text": "```markdown\nNACA RM No. L5I08\n\n2.8\n2.4\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-1.6\n-2.0\n-16\n-8\n0\n8\n16\n24\nSection angle of attack, $\\alpha_0$, deg\n\nSection lift coefficient, $c_l$\n\nR\n$\\nabla$ 13.9 x $10^6$\n$\\diamond$ 9.1\n$\\square$ 6.1\n$\\Delta$ Standard roughness\n6.1 x $10^6$\n\n.032\n.028\n.024\n.020\n.016\n.012\n.008\n.004\n0\n-.4\n-.8\n-1.2\n-1.6\n0\n.2\n.4\n.6\n.8\n1.0\nx/c\n\nSection drag coefficient, $c_d$\n\nR\n$\\nabla$ 13.9 x $10^6$\n$\\diamond$ 9.1\n$\\square$ 6.1\n$\\Delta$ Standard roughness\n6.1 x $10^6$\n\n-1.6\n-1.2\n-.8\n-.4\n0\n.4\n.8\n1.2\n1.6\nSection lift coefficient, $c_l$\n\n[Figure: Airfoil section profile]\n\nNACA\n\nFigure 29.- Aerodynamic characteristics of airfoil section 0, 24-inch chord.\n\n45\n```", "timestamp": "2026-07-22T04:13:18.244208+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 46, "total_pages": 92, "image_filename": "19930085930_p46.jpg", "text": "~~UNCLASSIFIED~~\nCONFIDENTIAL\n\nNACA RM L9D07\n\n0 2 4 6 8 10 12\nINCHES\n\nNozzle\nAdjustment\nmechanism\n\nTurning passage\n\nNACA\nL-58132\n\nFigure 14.- Model mounted in test setup with one side wall removed.\n~~UNCLASSIFIED~~\nCONFIDENTIAL\n\n45", "timestamp": "2026-07-22T04:13:18.431797+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 62, "total_pages": 118, "image_filename": "19930085838_p62.jpg", "text": "60\nNACA RM No. 19B23\n\n<!-- Image (108, 110, 839, 934) -->\n\n(a) $\\delta_f = 25^\\circ$\nFigure 8.- Continued.", "timestamp": "2026-07-22T04:13:18.606826+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 2, "total_pages": 42, "image_filename": "19930086078_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:13:21.525592+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 35, "total_pages": 59, "image_filename": "19930085936_p35.jpg", "text": "34\nNACA RM No. E9803\n\n<!-- Image (169, 109, 884, 848) -->\n\n(c) $\\theta = 180^\\circ$ longitudinal plane.\nFigure 6. - Continued. Pressure distributions along longitudinal planes at $0^\\circ$ angle of attack for range of yaw angles.", "timestamp": "2026-07-22T04:13:27.102780+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 12, "total_pages": 34, "image_filename": "19930086022_p12.jpg", "text": "10\nNACA RM L9E24\n\nA possible high-speed flight arrangement may incorporate an\naileron with a geared tab and no aerodynamic balance. Although this\nparticular type of design may prove satisfactory in some instances, it\nshould be pointed out that high positive values of $C_{h_a}/C_{h_\\delta}$ are\nattained in practice through the use of this type of lateral-control\narrangement. It is evident, therefore, that a geared-tab arrangement\nwithout aerodynamic balance would not prove satisfactory for this\nparticular case, inasmuch as it would tend to aggravate the already\nlarge positive values of $C_{h_a}/C_{h_\\delta}$.\n\nAnother approach to the high-speed flight problem is to incorpo-\nrate in the aileron design some degree of aerodynamic balance. In an\neffort to show the effect of an internally sealed aerodynamic balance\non $C_{h_\\delta}'$ (the rate of change of aileron hinge moment with deflection\nin a steady roll), some calculations by means of equations presented\nin reference 2 were made for various degrees of aerodynamic balance for\nthe configurations investigated and are presented in figure 12. It can\nbe seen that in the case of the plain wing, the degree of aerodynamic\nbalance required for $C_{h_\\delta}' = 0$ varies from approximately 45 percent\nof the aileron chord at $\\alpha = 0^\\circ$ to 60 percent at $\\alpha = 16^\\circ$. With flaps\ndeflected, the required aerodynamic balance varied from approximately 50\nto 30 percent of the aileron chord for angles of attack of $0^\\circ$ and $16^\\circ$,\nrespectively. With upper-surface fences installed, the aerodynamic\nbalance required was approximately 50 percent of the aileron chord\nthroughout the angle-of-attack range.\n\nAileron load coefficients.— The aileron load coefficients are\npresented in figures 4 to 6 for the purpose of supplying design infor-\nmation on the aerodynamic forces that would be anticipated on a\ngeometrical similar aileron. Due to the limitations of the strain-gage\narrangement, it was not possible to measure the aileron drag forces\nparallel to the wing chord line and, therefore, no attempt was made to\npresent true normal-force coefficients.\n\nCONCLUSIONS\n\nFrom an investigation of the lateral control characteristics at a\nReynolds number of 6,800,000 of a wing with the leading edge swept\nback $42^\\circ$ with and without high-lift devices, the following conclusions\ncan be made:\n\n1. At low total aileron deflections the rolling moments produced\nwere not influenced by modal configurations. At large total aileron\ndeflections ($50^\\circ$), lower values of rolling moments throughout the", "timestamp": "2026-07-22T04:13:27.348966+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 16, "total_pages": 54, "image_filename": "19930086015_p16.jpg", "text": "NACA RM A9E24 CONFIDENTIAL 15\n\n<!-- Image (126, 195, 905, 343) -->\n\n<!-- Image (126, 403, 838, 719) -->\n\nAll dimensions in inches\nSketch not to scale.\nFigure 2.—Static-pressure survey apparatus.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:13:29.101306+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 1, "total_pages": 44, "image_filename": "19930086081_p1.jpg", "text": "NACA RM L9H05\n290\nCopy\nRM L9H05\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nCONTROL EFFECTIVENESS LOAD AND HINGE-MOMENT\nCHARACTERISTICS OF A TIP CONTROL SURFACE\nON A DELTA WING AT A MACH NUMBER OF 1.9\n\nBy D. William Conner and Ellery B. May, Jr.\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nU.S.C. 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: NACA RESEARCH ABSTRACT NO. 145\nDATE: AUGUST 28, 1956\nWHL\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nOctober 7, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:13:33.057975+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 18, "total_pages": 36, "image_filename": "19930086003_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:13:35.013341+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 21, "total_pages": 67, "image_filename": "19930085965_p21.jpg", "text": "20\nNACA RM E9E06\n11125\n\nThe total air-gap permeance will equal one-half of this value, or\n$$P_T = \\frac{P_1}{2} + P_2 + P_3 + P_{4+5} + P_6 + P_7 + 2P_8 \\quad (21)$$\n\nThe following equation can be written from the total-air-gap magnetomotive-force equation for the real component\n$$N_a I_{\\text{max } a} = \\frac{A}{0.4 \\pi l_o} \\frac{\\Phi_{\\text{max } 2} \\cos (45^\\circ + \\beta_2 - \\delta_2)}{P_T}$$\n\nSubstitution of equation (10b) gives the equation for air-gap ampere turns\n$$N_a I_{\\text{max } a} = \\frac{A}{0.4 \\pi \\sqrt{2} cl_o} \\frac{\\cos (45^\\circ + \\beta_2 - \\delta_2) B_{\\text{max } 1}}{P_T} \\sqrt{1 + 2 \\cos (45^\\circ + \\alpha_2) \\left( \\frac{H_{\\text{max } 2}}{H_{\\text{max } 1}} - 1 \\right)} \\quad (22)$$\n\nThe final determination needed is that for the magnetomotive-force drop along the blade. Previous discussion showed that the flux density at the outer surface of the blade is the result of the impressed magnetomotive force alone; the following equation is therefore true:\n$$H_{\\text{max } 2} = \\frac{0.4 \\pi N_b I_{\\text{max } b}}{L}$$\n\nThus\n$$N_b I_{\\text{max } b} = \\frac{L H_{\\text{max } 2}}{0.4 \\pi} \\quad (23)$$\n\nThe total effective ampere turns required are\n$$NI = \\frac{N_a I_{\\text{max } a} + N_b I_{\\text{max } b}}{\\sqrt{2}} \\quad (24)$$\n\nWhen the previously derived equations are used, the eddy-current power per square inch of blade surface appearing as heat can be calculated from equation (20b), and the voltage can be", "timestamp": "2026-07-22T04:13:42.386945+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 46, "total_pages": 104, "image_filename": "19930085842_p46.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:13:42.870738+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 32, "total_pages": 50, "image_filename": "19930085952_p32.jpg", "text": "NACA RM L9C24\n31\n\n<!-- Image (124, 168, 909, 410) -->\n\n(a) Articulated propellers.\n\nTotal increment of $\\Delta C_{mp}$\nIncrement due to propellers and wing\nIncrement due to power effects on all-movable horizontal tail\n\n<!-- Image (124, 518, 909, 858) -->\n\n(b) Rigid propellers.\n\nFigure 11.- Increments of pitching-moment coefficients due to propeller operation. Full-power operation; $\\delta_F = 0^\\circ$.", "timestamp": "2026-07-22T04:13:45.177264+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 17, "total_pages": 32, "image_filename": "19930085982_p17.jpg", "text": "NACA RM E5E13\n15\n\nbasis of constant total enthalpy and symmetrical velocity diagram at all radii:\n\n1. At the design speed of 18,060 rpm, a mean blade-element efficiency of 0.97 was obtained at an equivalent weight flow of 24.5 pounds per second and an over-all total-pressure ratio of 1.18. The maximum pressure ratio obtained was 1.28 at a weight flow of 21.0 pounds per second. These combinations of weight flow and pressure ratio were good for a compressor stage with a hub-to-tip-radius ratio of 0.50. In addition, the high rotative speed of this inlet-stage design will permit higher pressure ratios over the later stages of a multistage compressor.\n\n2. The radial distribution of axial velocity at the measuring station immediately downstream of the rotor was somewhere between the design distribution, which assumed no change in axial velocity through the rotor row, and that distribution which would have been obtained had the flow been in simple-radial equilibrium.\n\n3. The enthalpy addition increased more rapidly at the blade tip than at the hub with decreasing weight flow. In the range of peak efficiency, the design conditions of constant enthalpy at all radii were very nearly achieved.\n\n4. The equation suggested by Kantrowitz and Daum (with constant $K = 0.9$) used in setting the blades was invalid in the range of solidities and stagger angles encountered near the tip section of this design. In general, the turning obtained was $4^\\circ$ to $9^\\circ$ below that predicted by assuming straight-through flow. As a result, the design pressure ratio and peak efficiency were obtained at a flow considerably below that for which the compressor was designed.\n\n5. The data indicate that a variable-camber blade would minimize the radial variations in energy addition and would tend to increase the range of high-efficiency operation over that obtained with this constant-camber blade.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.\n\nREFERENCES\n\n1. Sinnette, John T., Jr.: Analysis of Effect of Basic Design Variables on Subsonic Axial-Flow-Compressor Performance. NACA Rep. 901, 1948.", "timestamp": "2026-07-22T04:13:45.772344+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 23, "total_pages": 43, "image_filename": "19930085958_p23.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:13:47.058535+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 39, "total_pages": 47, "image_filename": "19930085918_p39.jpg", "text": "```markdown\n38\nNACA RM A9D29\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\n<!-- Image (155, 233, 786, 686) -->\n\n(a) $\\alpha = 0.1^\\circ$.\n\nFigure 10: Chordwise pressure distributions for $45^\\circ$ swept-forward wing with a cambered nose.\n```", "timestamp": "2026-07-22T04:13:47.285250+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 62, "total_pages": 82, "image_filename": "19930085551_p62.jpg", "text": "NACA RM No. L8K30\n61\n\n[Figure: Graphs showing Rudder pedal force, Control wheel force, Angle of bank, and Control position versus Sideslip angle]\n\n(b) 150 miles per hour.\nFigure 15.— Continued.", "timestamp": "2026-07-22T04:13:49.494935+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 96, "total_pages": 122, "image_filename": "19930082090_p96.jpg", "text": "94\nNACA TN No. 1455\n\n<!-- Image (123, 119, 759, 912) -->\n\nFigure 47.- Thermal output and isothermal frictional pressure drops of pin heat exchanger L.", "timestamp": "2026-07-22T04:13:57.161365+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 47, "total_pages": 92, "image_filename": "19930085930_p47.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:14:02.006389+00:00"}
{"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 75, "total_pages": 85, "image_filename": "19930085529_p75.jpg", "text": "74\nNACA RM No. L8A30a\n\nTABLE 68\n$$\n\\left[ \\Lambda = -30^\\circ, \\delta_{te} = 10.0^\\circ, \\alpha = -6^\\circ \\right]\n$$\nCONFIDENTIAL\n\n| | | UPPER SURFACE | | | | | | | LOWER SURFACE | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Tube** | **Per-cent chord** | **Mach Number** | | | | **Tube** | **Per-cent chord** | **Mach Number** | | | |\n| | | **0.60** | **0.80** | **0.85** | **0.89** | | | **0.60** | **0.80** | **0.85** | **0.89** |\n| A 1 | 2.0 | -- | -- | -- | -- | 86 | 3.0 | -- | -- | -- | -- |\n| 2 | 6.0 | -- | -- | -- | -- | 87 | 10.0 | -- | -- | -- | -- |\n| 3 | 15.0 | -- | -- | -- | -- | 88 | 25.0 | -- | -- | -- | -- |\n| 4 | 27.5 | -- | -- | -- | -- | 89 | 41.0 | -- | -- | -- | -- |\n| 5 | 40.0 | -- | -- | -- | -- | 90 | 52.5 | +0.108 | +0.101 | +0.102 | +0.061 |\n| 6 | 50.0 | -0.198 | -0.223 | -0.223 | -0.183 | 91 | 58.5 | -.068 | -.094 | -.091 | -.093 |\n| 7 | 59.0 | -.167 | -.181 | -.170 | -.159 | 92 | 72.5 | -.028 | -.009 | .001 | .020 |\n| 8 | 67.5 | -.148 | -.096 | -.086 | -.069 | 93 | 84.0 | -- | -- | -- | -- |\n| 9 | 77.5 | -- | -- | -- | -- | 94 | 94.0 | -- | -- | -- | -- |\n| 10 | 88.0 | -- | -- | -- | -- | | | | | | |\n| 11 | 98.0 | -- | -- | -- | -- | | | | | | |\n| | | | | | | | | | | | |\n| B12 | 2.0 | .192 | .191 | .100 | .101 | 95 | 3.0 | -.752 | -- | -.961 | -.832 |\n| 13 | 6.0 | .047 | .065 | .070 | .075 | 96 | 10.0 | -.367 | -.641 | -.221 | -.221 |\n| 14 | 15.0 | -.141 | -.185 | -.203 | -.209 | 97 | 25.0 | -.282 | -.336 | -.462 | -.811 |\n| 15 | 27.5 | -.299 | -.390 | -.378 | -.416 | 98 | 41.0 | -.256 | -.281 | -.318 | -.414 |\n| 16 | 40.0 | -.269 | -.343 | -.407 | -.444 | 99 | 52.5 | -.185 | -.195 | -.190 | -.094 |\n| 17 | 50.0 | -.261 | -.318 | -.337 | -.404 | 100 | 58.5 | -.131 | -.118 | -.129 | -.016 |\n| 18 | 59.0 | -.231 | -.261 | -.323 | -.454 | 101 | 72.5 | -.039 | -.025 | -.014 | -.019 |\n| 19 | 67.5 | -.170 | -.227 | -.145 | -.064 | 102 | 84.0 | .091 | .047 | .046 | .077 |\n| 20 | 77.5 | -.095 | -.093 | -.022 | -.058 | 103 | 94.5 | .114 | .133 | .142 | .162 |\n| 21 | 88.0 | .084 | .097 | .061 | .074 | | | | | | |\n| 22 | 95.5 | -- | -- | -- | -- | | | | | | |\n| | | | | | | | | | | | |\n| C23 | 2.0 | .137 | .174 | .434 | .422 | 104 | 3.0 | -.504 | -.641 | -.693 | -.619 |\n| 24 | 6.0 | .068 | .111 | .131 | .143 | 105 | 10.0 | -.341 | -.482 | -.528 | -.564 |\n| 25 | 15.0 | -.113 | -.111 | -.099 | -.091 | 106 | 25.0 | -.301 | -.408 | -.498 | -.523 |\n| 26 | 27.5 | -.215 | -.268 | -.278 | -.286 | 107 | 41.0 | -.239 | -.300 | -.419 | -.506 |\n| 27 | 40.0 | -.272 | -.393 | -.413 | -.473 | 108 | 52.5 | -.223 | -.244 | -.241 | -.303 |\n| 28 | 50.0 | -.246 | -.334 | -.343 | -.413 | 109 | 58.5 | -.143 | -.158 | -.146 | -.200 |\n| 29 | 59.0 | -.218 | -.287 | -.337 | -.481 | 110 | 72.5 | -.044 | -.042 | -.032 | -.006 |\n| 30 | 67.5 | -.180 | -.235 | -.265 | -.287 | 111 | 85.0 | .034 | .047 | .046 | .077 |\n| 31 | 77.5 | -.083 | -.081 | -.083 | -.085 | 112 | 94.6 | .096 | .110 | .113 | .125 |\n| 32 | 88.0 | .030 | .037 | .041 | .049 | | | | | | |\n| 33 | 95.5 | -- | -- | -- | -- | | | | | | |\n| | | | | | | | | | | | |\n| D34 | 2.0 | .068 | .135 | .362 | .379 | 113 | 3.0 | -.439 | -.560 | -.626 | -.583 |\n| 35 | 15.0 | -.113 | -.160 | -.097 | -.089 | 114 | 10.0 | -.308 | -.386 | -.482 | -.493 |\n| 36 | 27.5 | -.209 | -.241 | -.251 | -.260 | 115 | 25.0 | -.258 | -.325 | -.416 | -.447 |\n| 37 | 40.0 | -.275 | -.347 | -.384 | -.435 | 116 | 41.0 | -.261 | -.307 | -.389 | -.445 |\n| 38 | 50.0 | -.279 | -.345 | -.397 | -.504 | 117 | 52.5 | -.201 | -.224 | -.296 | -.347 |\n| 39 | 59.0 | -.249 | -.311 | -.371 | -.547 | 118 | 58.5 | -.114 | -.158 | -.180 | -.244 |\n| 40 | 67.5 | -- | -- | -- | -- | 119 | 72.5 | -.034 | -.058 | -.071 | -.104 |\n| 41 | 77.5 | -.106 | -.091 | -.071 | -.110 | 120 | 87.4 | .051 | .050 | .048 | .077 |\n| 42 | 87.5 | -.001 | -.001 | -.005 | -.030 | 121 | 94.0 | .074 | .080 | .074 | .082 |\n| 43 | 94.2 | .064 | .071 | .087 | .095 | | | | | | |\n| | | | | | | | | | | | |\n| E44 | 2.0 | .022 | .065 | .233 | .231 | 122 | 3.0 | -.261 | -.310 | -.340 | -.372 |\n| 45 | 6.0 | -.009 | .016 | .039 | .074 | 123 | 10.0 | -.215 | -.262 | -.287 | -.303 |\n| 46 | 15.0 | -.157 | -.176 | -.160 | -.145 | 124 | 25.0 | -.201 | -.244 | -.277 | -.315 |\n| 47 | 27.5 | -.232 | -.305 | -.320 | -.306 | 125 | 41.0 | -.201 | -.246 | -.271 | -.294 |\n| 48 | 40.0 | -.328 | -.428 | -.462 | -.473 | 126 | 52.5 | -.146 | -.187 | -.191 | -.223 |\n| 49 | 50.0 | -.306 | -.402 | -.470 | -.582 | 127 | 58.5 | -.106 | -.125 | -.134 | -.136 |\n| 50 | 59.0 | -.279 | -.349 | -.438 | -.618 | 128 | 72.5 | -.040 | -.052 | -.052 | -.066 |\n| 51 | 67.5 | -.217 | -.301 | -.366 | -.480 | 129 | 78.0 | .001 | -.012 | -.022 | -.037 |\n| 52 | 77.5 | -.148 | -.211 | -.246 | -.279 | 130 | 85.0 | .050 | .044 | .046 | .066 |\n| 53 | 88.5 | .103 | .101 | .100 | .094 | 131 | 94.1 | .086 | .088 | .084 | .077 |\n| 54 | 95.5 | .070 | .089 | .070 | .071 | | | | | | |\n| | | | | | | | | | | | |\n| F55 | 2.0 | .009 | .066 | .201 | .238 | 132 | 3.0 | -.261 | -.327 | -.360 | -.394 |\n| 56 | 6.0 | -.017 | .015 | .043 | .075 | 133 | 10.0 | -.198 | -.248 | -.270 | -.297 |\n| 57 | 15.0 | -.149 | -.159 | -.141 | -.132 | 134 | 25.0 | -.197 | -.231 | -.258 | -.281 |\n| 58 | 27.5 | -.227 | -.293 | -.310 | -.294 | 135 | 41.0 | -.217 | -.213 | -.232 | -.268 |\n| 59 | 40.0 | -.328 | -.420 | -.460 | -.462 | 136 | 52.5 | -.118 | -.159 | -.170 | -.211 |\n| 60 | 50.0 | -.314 | -.428 | -.491 | -.570 | 137 | 58.5 | -.040 | -.087 | -.094 | -.092 |\n| 61 | 59.0 | -.311 | -.381 | -.451 | -.523 | 138 | 72.5 | .106 | .103 | .101 | .101 |\n| 62 | 67.5 | -.284 | -.331 | -.395 | -.474 | 139 | 83.4 | .129 | .129 | .129 | .129 |\n| 63 | 86.4 | .114 | .123 | .124 | .124 | 140 | 94.0 | .102 | .094 | .089 | .078 |\n| 64 | 94.6 | .043 | .044 | .044 | .033 | | | | | | |\n| | | | | | | | | | | | |\n| G65 | 2.0 | .018 | .092 | .328 | .353 | 141 | 3.0 | -.220 | -.282 | -.314 | -.337 |\n| 66 | 6.0 | .002 | .037 | .069 | .091 | 142 | 10.0 | -.147 | -.204 | -.229 | -.251 |\n| 67 | 15.0 | -.151 | -.148 | -.125 | -.098 | 143 | 25.0 | -.177 | -.237 | -.273 | -.312 |\n| 68 | 27.5 | -.224 | -.285", "timestamp": "2026-07-22T04:14:08.382233+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 63, "total_pages": 118, "image_filename": "19930085838_p63.jpg", "text": "```markdown\nNACA RM No. L9B23\n\nAileron section hinge-moment coefficient, $c_{h_a}$\n\n$\\delta_i = 0^\\circ$\n$\\delta_a$ (deg)\n-5\n-10\n-15\n\n$\\delta_i = 0^\\circ$\n$\\delta_a$ (deg)\n0\n5\n10\n15\n\nSection angle of attack, $\\alpha_o$, deg\n\n(e) $\\delta_r = 40^\\circ$.\nFigure 8.- Continued.\n\nNACA\n\n61\n```", "timestamp": "2026-07-22T04:14:11.876254+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 17, "total_pages": 54, "image_filename": "19930086015_p17.jpg", "text": "16\nCONFIDENTIAL\nNACA RM A9E24\n\n<!-- Image (349, 124, 866, 277) -->\n\nCone detail\n\n<!-- Image (114, 312, 812, 823) -->\n\nFigure 3 - Stream-angle survey apparatus.\n\nAll dimensions in inches\nSketch not to scale\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:14:12.119598+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 2, "total_pages": 44, "image_filename": "19930086081_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:14:15.461187+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 3, "total_pages": 42, "image_filename": "19930086078_p3.jpg", "text": "NACA RM L9H04\nCONFIDENTIAL\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nRESEARCH MEMORANDUM\nINVESTIGATION OF EXTENSIBLE WING-TIP AILERONS ON AN\nUNTAPERED SEMISPAN WING AT 0° AND 45° SWEEPBACK\nBy John R. Hagerman and William M. O'Hare\nSUMMARY\nA low-speed wind-tunnel investigation was made to determine the\nlateral control characteristics of extensible wing-tip ailerons on an\nuntapered semispan wing having two configurations; one configuration\nwas unswept and had an aspect ratio of 3.13 and the other configuration\nwas swept back 45° and had an aspect ratio of 1.59. Three plan forms\nof extensible ailerons were investigated on each wing configuration at\nvarious amounts of extension and deflection relative to the wing-chord\nplane. Also, wing aerodynamic characteristics were determined for the\ntwo plain-wing configurations.\nThe results indicate that sufficient aileron effectiveness was\ngenerally obtained at moderate and high lift coefficients with the\nextensible ailerons investigated. However, the control effectiveness\nat low lift coefficients appears to be inadequate for satisfactory\napplication to an airplane. It is thought that the extensible ailerons\nmay be sufficiently effective for some types of missiles.\nYawing moments produced by the extensible ailerons investigated\nwere comparable to those produced by conventional flap-type ailerons.\nINTRODUCTION\nThe National Advisory Committee for Aeronautics is currently\ninvestigating the lateral-control problem associated with transonic\nand supersonic wing configurations. Because conventional flap-type\nailerons do not always provide adequate lateral control throughout\nthe speed range, particularly above the wing critical speed, other\nlateral-control devices are being investigated. Among the lateral-\ncontrol devices being investigated are extensible wing-tip ailerons.\nCONFIDENTIAL", "timestamp": "2026-07-22T04:14:16.386541+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 19, "total_pages": 36, "image_filename": "19930086003_p19.jpg", "text": "NACA RM L9I08\n17\n\nCONFIDENTIAL\n\n[Figure: A cutaway view of a model wing with a sponge-wiper seal installation. The model has a 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section. Three vertical fins are visible in the background.]\n\nFigure 5.- A cutaway view showing the sponge-wiper seal installation on the model with 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section.\n\nNACA\nL-61938\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:14:24.101749+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 47, "total_pages": 104, "image_filename": "19930085842_p47.jpg", "text": "NACA RM L9C29\n43\n\n[Figure: Four photographs of a model in a wind tunnel, arranged in a 2x2 grid.]\n\n$\\alpha = 26.3^\\circ; C_L = 1.01$\n\n$\\alpha = 29.3^\\circ; C_L = 1.12$\n\n$\\alpha = 35.3^\\circ; C_L = 1.31$\n\n$\\alpha = 44.3^\\circ; C_L = 0.81$\n\n(b) Configuration 10.\nFigure 15.— Concluded.\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS.", "timestamp": "2026-07-22T04:14:28.642370+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 18, "total_pages": 32, "image_filename": "19930085982_p18.jpg", "text": "16\nNACA RM E9E13\n\n2. Kantrowitz, Arthur, and Daum, Fred L.: Preliminary Experimental\nInvestigation of Airfoils in Cascade. NACA CB, July 1942.\n\n3. Mankuta, Harry, and Guentert, Donald C.: Investigation of Per-\nformance of Single-Stage Axial-Flow Compressor Using NACA\n5509-34 Blade Section. NACA RM E8F30, 1948.\n\n4. Bogdonoff, Seymour M., and Bogdonoff, Harriet E.: Blade Design\nData for Axial-Flow Fans and Compressors. NACA ACR L5F07a,\n1945.", "timestamp": "2026-07-22T04:14:29.502379+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 24, "total_pages": 43, "image_filename": "19930085958_p24.jpg", "text": "NACA RM No. L9B11\n23\n\n<!-- Image (149, 106, 899, 813) -->\n\nFigure 4.- Characteristics of a 42° sweptback wing with and without split flaps and upper-surface stall-control fences.", "timestamp": "2026-07-22T04:14:30.316843+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 63, "total_pages": 82, "image_filename": "19930085551_p63.jpg", "text": "62\nNACA RM No. L8K30\n\n<!-- Image (256, 116, 739, 876) -->\n\n(c) 200 miles per hour.\nFigure 15.- Continued.", "timestamp": "2026-07-22T04:14:37.777275+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 40, "total_pages": 47, "image_filename": "19930085918_p40.jpg", "text": "NACA RM A9D29\n39\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nPressure coefficient, P\n-16\n-12\n-.8\n-.4\n0\n.4\n.8\n\n-12\n-.8\n-.4\n0\n.4\n.8\n\n-12\n-.8\n-.4\n0\n.4\n.8\n\nSpanwise\nstation, 2y/b\n28.1%\n57.4%\n85.0%\n\nChordwise station, x/c\n2\n4\n6\n8\n10\n\n(b) $\\alpha=6.3^\\circ$\n\nNACA\n\nFigure 10.—Continued.", "timestamp": "2026-07-22T04:14:38.209499+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 36, "total_pages": 59, "image_filename": "19930085936_p36.jpg", "text": "```markdown\nNACA RM No. E9B03\n35\n\n<!-- Image (151, 133, 855, 848) -->\n\n(d) $\\theta = 225^\\circ$ longitudinal plane.\nFigure 6. - Continued. Pressure distributions along longitudinal\nplanes at $0^\\circ$ angle of attack for range of yaw angles.\n```", "timestamp": "2026-07-22T04:14:42.440335+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 13, "total_pages": 34, "image_filename": "19930086022_p13.jpg", "text": "NACA RM LGE24\n11\n\nangle-of-attack range were obtained for the flap-deflected conditions\nthan for the plain wing because of the loss in effectiveness of the\ndowngoing aileron.\n\n2. For a total aileron deflection of $50^\\circ$, longitudinal trim changes\nwere obtained for the flap-deflected conditions which would require a\nsmall degree of elevator travel to balance.\n\n3. The hinge moment of the unbalanced aileron was approximately\nuniform in the low angle-of-attack range and decreased appreciably in\nthe higher angle-of-attack range for the plain wing and for the wing\nwith leading- and trailing-edge flaps deflected. The addition of\nfences to the flapped configuration resulted in a more uniform hinge-\nmoment variation. Accordingly, only for the latter case was the\naerodynamic-balance requirement essentially constant throughout the\nangle-of-attack range.\n\n4. The calculated value of the rolling-effectiveness parameter for\nthe plain wing at zero angle of attack agreed within 2 percent with the\nexperimental value.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.", "timestamp": "2026-07-22T04:14:44.292370+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 97, "total_pages": 122, "image_filename": "19930082090_p97.jpg", "text": "NACA TN No. 1455\n95\n\n<!-- Image (157, 166, 912, 751) -->\n\nFigure 48.- Schematic diagram of cast-aluminum heat exchanger M and air shroud. Weight of heat exchanger, 31.0 pounds.\n\n| | Air side | Gas side |\n| :--- | :--- | :--- |\n| Cross-sectional area, sq ft | 0.170 | 0.111 |", "timestamp": "2026-07-22T04:14:45.915470+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 48, "total_pages": 92, "image_filename": "19930085930_p48.jpg", "text": "CONFIDENTIAL\n\nNACA RM L59D7\n\n0 2 4 6 8 10 12\nINCHES\n\nNozzle walls\nBleed off\nConcave surface\nConvex surface\nSurvey rake\n\nNACA\nL-58133\n\nFigure 15.- Model mounted in test setup with one side wall removed.\n\nCONFIDENTIAL\n\n47", "timestamp": "2026-07-22T04:14:48.904897+00:00"}
{"citation_id": "19930085588", "source_url": "https://ntrs.nasa.gov/api/citations/19930085588/downloads/19930085588.pdf", "page_number": 47, "total_pages": 51, "image_filename": "19930085588_p47.jpg", "text": "46\n\n2.8\n2.6\n2.4\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-1.6\n-2.0\n-2.4\n-2.8\nSection lift coefficient, $c_l$\n-24 -16 -8 0 8 16 24\nSection angle of attack, $\\alpha_o$, deg\n\nR\n$\\nabla$ 14.2 $\\times$ 10$^6$\n$\\circ$ 9.0\n$\\circ$ 6.1\n$\\circ$ 3.1\n$\\Delta$ Standard roughness\n6.1 $\\times$ 10$^6$\n\n.032\n.028\n.024\n.020\n.016\n.012\n.008\n.004\n0\n-.4 -.8 -1.2 -1.6\nSection drag coefficient, $c_d$\n-1.6 -1.2 -.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\n0 .2 .4 .6 .8 1.0\nx/c\n\nR\n$\\circ$ 9.0 $\\times$ 10$^6$\n$\\circ$ 6.1\n$\\circ$ 3.1\n$\\Delta$ Standard roughness\n6.1 $\\times$ 10$^6$\n\nNACA\n\nFigure 30.—Aerodynamic characteristics of airfoil section P, 24-inch chord.\n\nNACA RM No. 18108", "timestamp": "2026-07-22T04:14:51.886686+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 6, "total_pages": 20, "image_filename": "19930086060_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM L9F02\n\nFor most of the transonic range ($0.82 < M < 1.02$), the 60-percent and 40-percent stations had significantly less drag than resulted from the 20-percent station of maximum diameter. Above $M = 1.10$ the 60-percent station resulted in approximately 15 percent less drag than did the 40-percent station and approximately 50 percent less drag than did the 20-percent station of maximum diameter. The foregoing discussion holds roughly true regardless of whether drag coefficient $C_D$ is based on frontal area, wetted skin area, or $(vol)^{2/3}$.\n\nBy means of the method of Von Kármán and Moore (reference 4), the pressure distribution was calculated at $M = 1.40$ for the configuration tested and, in addition, for an 80-percent position of maximum diameter $D_{max}$. These distributions are shown in figure 5. The calculations were made for only one supersonic Mach number in consideration of Laitone's work (reference 5) in which the method was concluded to be most accurate near $M = \\sqrt{2}$. Although the method cannot be rigidly applied to the 20-percent maximum-diameter position (due to the questionable nature of the flow at $M = 1.4$), pressure distribution of the flow is included for the purpose of indicating the type of variation that might be expected for extreme forward positions of maximum diameter. As will be seen subsequently, inclusion of pressure distribution is further justified in view of its favorable agreement with experiment. The theoretical distributions are shown to be basically of two distinct types of variations depending upon the position of $D_{max}$ relative to the position of symmetry. The position of symmetry is that station of $D_{max}$ for which the nose and stern of the body are of equal curvature. For the test bodies of this paper, the 57.1-percent station is the relative position of symmetry. From figure 5, where the maximum diameter is well forward of the symmetry position, the characteristic variation is one of a strong compression at the nose followed by a rapid expansion to peak suction at the maximum diameter and then a gradual recompression to the stern. Where the maximum diameter is well behind the symmetry position, the characteristic variation is one of a relatively weak compression at the nose followed by a gradual expansion to the maximum diameter, and then by an extremely rapid expansion to a very large peak suction on the boattail. Since the method of pressure distributions is, however, based upon the assumption of small disturbances, it is doubtful whether peak suctions of the order indicated in figure 5(d) would be correct to within the limits of the theory.\n\nThe variation of total drag with position of $D_{max}$ has been calculated at $M = 1.40$ and is compared with the experimental variation in figure 6. The experimental variation was based on three test points and was, in part, guided by the calculated variation. The general agreement between calculated variation and the test points\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:14:56.568500+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 64, "total_pages": 118, "image_filename": "19930085838_p64.jpg", "text": "62\nNACA RM No. L9B23\n\n<!-- Image (107, 110, 838, 934) -->\n\n(r) $\\delta_f = 40^\\circ$.\nFigure 8.- Continued.", "timestamp": "2026-07-22T04:14:56.775402+00:00"}