PDF/ocr_dataset_1001.jsonl

{"citation_id": "19930085917", "source_url": "https://ntrs.nasa.gov/api/citations/19930085917/downloads/19930085917.pdf", "page_number": 37, "total_pages": 38, "image_filename": "19930085917_p37.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:10:36.354248+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 2, "total_pages": 20, "image_filename": "19930086060_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:10:38.312879+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 42, "total_pages": 92, "image_filename": "19930085930_p42.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n\nNACA RM L9G07\n\n[Figure: Diagram of a blade profile with an arrow indicating direction. The NACA logo is present.]\n\n(b) Blade corresponding to figure 10(b). Stagger angle, $59^\\circ$.\n\nFigure 11.- Concluded.\n\n41", "timestamp": "2026-07-22T04:10:39.233007+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 18, "total_pages": 67, "image_filename": "19930085965_p18.jpg", "text": "NACA RM E9E06\n\ntemperature near the mean temperature of the water. From runs with and without insulation, insulation on the water jacket to reduce heat transfer between the atmosphere and the jacket apparently was unnecessary in this investigation.\n\nThe pulsating unidirectional operation was synthesized by superimposing a direct current and an alternating current in the inductor of the experimental unit, as shown in figure 11. The direct current to the inductor was introduced by means of a filtering network. The 0.55-henry choke, which was especially wound with a large air gap to prevent saturation, had approximately 10 times the inductance of the inductor and thus shunted only about 10 percent of the alternating current from the inductor. The inductance value of the remaining choke was not critical, but its impedance was sufficient, relative to the 20-microfarad condenser, to force most of the alternating current of the 0.55-henry choke through the measuring meter in series with the 20-microfarad condenser. Because the ratio of inductive reactance to resistance of the 0.55-henry choke and the inductor are sufficiently alike to make the currents in phase for all practical purposes, the difference between the readings of the two alternating-current ammeters was therefore the alternating current in the inductor. No appreciable alternating current passed through the direct-current ammeter or the battery circuit and the direct-current meter read the true direct current applied to the inductor.\n\nThe first investigations were made on annealed SAE 1020 steel with a total air gap between the laminated iron core and a sample blade of approximately 0.04 inch. The air gaps used for these investigations were approximately those that would exist in practice between the blade and the adjacent part through which the flux would flow. In order to permit a correlation between experimental results and the analysis, the blades were made of rectangular cross section. The actual size of the sample, however, which was 0.125 by 1 by $4\\frac{5}{8}$ inches, is approximately the same as the inlet guide vanes of a typical axial-flow compressor. The second and third experiments were made with a blade of Armco Magnetic Ingot Iron of the same size. In the second experiment, the air gap was kept at approximately 0.04 inch and in the third experiment the air gap was increased to approximately 0.06 inch. A sample of Hipernik, 0.094 by 1 by $4\\frac{5}{8}$ inches, with a total air gap of 0.04 inch was used in the fourth experiment. This material has high resistivity and high maximum flux density. The investigations on all samples were made by varying the number of magnetizing ampere turns and measuring the heat generated.", "timestamp": "2026-07-22T04:10:39.472393+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 92, "total_pages": 122, "image_filename": "19930082090_p92.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:10:43.420839+00:00"}
{"citation_id": "19930085957", "source_url": "https://ntrs.nasa.gov/api/citations/19930085957/downloads/19930085957.pdf", "page_number": 25, "total_pages": 27, "image_filename": "19930085957_p25.jpg", "text": "```markdown\n24\nNACA RM L9E02\n\n<!-- Image (172, 98, 752, 874) -->\n\nFigure 12.- Effect of roughness on the stalling characteristics of the\n$42^\\circ$ sweptback wing with $0.725 \\frac{b}{2}$ leading-edge flaps and half-span\nsplit flaps. $\\delta_F = 60^\\circ$; $R = 3.0 \\times 10^6$.\n```", "timestamp": "2026-07-22T04:10:54.715984+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 14, "total_pages": 36, "image_filename": "19930086003_p14.jpg", "text": "CONFIDENTIAL\n\nBump surface\n\n4.53\n45° .10\n.76\n1.60\n.40\n\nCenterline of balance\n\n1.18 Maximum diameter\n2.50\nEnd plate used with floating tail in fuselage\n\nCONFIDENTIAL\n\nNACA\n\n0 1 2\nScale, inches\n\nFigure 2.— Details of free-floating tail mounted in fuselage of a model with 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section.\n\nNACA RM L9108\n12", "timestamp": "2026-07-22T04:10:56.199722+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 9, "total_pages": 34, "image_filename": "19930086022_p9.jpg", "text": "```markdown\nNACA RM L5E24\n7\n\njet-boundary correction has been applied to the pitching moment.\nCorrections for jet-boundary effects on rolling and yawing moments\nwere found to be small and therefore have not been applied.\n\nAs previously stated, the aileron seal was not continuous along\nthe span of the aileron due to the installation of the strain-gage\nbeams. This discontinuity resulted in some degree of leakage across\nthe seal. A calibration of the leakage through the aileron seal\nindicated a leakage factor E of 0.13. The resultant balance-chamber\npressures presented herein have been corrected to a 100-percent sealed\ncondition. No correction for the effects of leakage on the aileron\nhinge-moment and load characteristics have been made to the data since\nthe amount of leakage (based on effective gap area defined in\nreference 5) was found to be very small, and the effects are believed\nto be negligible.\n\nRESULTS AND DISCUSSION\n\nLateral characteristics.— The aileron characteristics for the\nplain wing and the wing equipped with high-lift and stall-control\ndevices are presented in figures 4 to 6. Attention is called to the\nfact that the rolling-moment coefficients for the plain wing are not\nfaired beyond angles of attack near maximum lift. Beyond that point\nthe lift curve levels off and, as a result of the varying degree of\nflow separation that existed on the wing at the high angles of attack,\nthe rolling-moment data became very erratic. Yawing moments for\nintermediate aileron deflections have been omitted from figures 4 to 6\ninasmuch as the variations were found to be small and can be assumed to\nbe approximately linear through the range investigated.\n\nComparison of the rolling moments produced by equal up and down\naileron deflections for various configurations (fig. 7) indicates that\nat low values of aileron deflections the rolling moment produced was\nnot influenced by model configuration. At larger total aileron\ndeflections ($50^\\circ$) the addition of high-lift and stall-control devices\nresulted in somewhat lower values of rolling moment throughout the\nangle-of-attack range and, as might be expected from the loss in\nrolling moment, also resulted in less adverse yawing moments than were\nobtained for the plain wing.\n\nThe variation of $C_l$ with aileron deflection (fig. 8) (for the\nplain-wing condition) was approximately the same for corresponding\npositive and negative aileron deflections for all of the angles of\nattack investigated. On the other hand, with leading- and trailing-\nedge flaps deflected, the values of $C_l$ for positive aileron\n```", "timestamp": "2026-07-22T04:11:00.852720+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 42, "total_pages": 104, "image_filename": "19930085842_p42.jpg", "text": "38\nNACA RM L9C29\n\n[Figure: A graph with three curves plotted on a grid. The x-axis is labeled \"Angle of attack, $\\alpha$, deg\" and ranges from -10 to 50. The left y-axis is labeled \"Lift coefficient, $C_L$\" and ranges from 0 to 10. The right y-axis is labeled \"Drag coefficient, $C_D$\" and ranges from 0 to 8. The top y-axis is labeled \"Pitching-moment coefficient, $C_m$\" and ranges from -2 to 1. The curves are labeled $C_m$, $C_L$, and $C_D$. A box in the bottom right corner of the graph reads \"NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\".]\n\nFigure 12.- Variation of $C_L$, $C_D$, and $C_m$ with $\\alpha$ of a $\\frac{1}{3}$-scale model of the airplane. Complete model configuration; propellers removed.", "timestamp": "2026-07-22T04:11:05.632932+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 35, "total_pages": 47, "image_filename": "19930085918_p35.jpg", "text": "34\nNACA RM A9D29\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nPressure coefficient, P\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\n-32\n-28\n-24\n-20\n-16\n-12\n-8\n-4\n0\n-20\n-16\n-12\n-8\n-4\n0\n4\n8\n\nSpanwise\nstation, 2y/b\n28.1%\n57.4%\n85.0%\n\nChordwise station, x/c\n4\n6\n8\n10\n\nNACA\n\n(d) $\\alpha=16.6^\\circ$.\n\nFigure 9. - Continued", "timestamp": "2026-07-22T04:11:07.545809+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 28, "total_pages": 50, "image_filename": "19930085952_p28.jpg", "text": "NACA RM L9C24\n27\n\n$\\delta_a, deg$\n$C_m$\n$C_L$\n(b) Rigid propellers.\n\n$\\delta_a, deg$\n$C_m$\n$C_L$\n(a) Articulated propellers.\n\nFigure 5.- Variation of $C_m$ with $C_L$ for several ailavator deflections for simulated full-power operation. $\\delta_F = 0^\\circ$.", "timestamp": "2026-07-22T04:11:07.729266+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 20, "total_pages": 43, "image_filename": "19930085958_p20.jpg", "text": "NACA RM No. L9B11\n19\n\n68.25\n27.30\n0.75 b/2\n17.06\n0.60 b/2\n3.40\n3.40\n7.71\n8.40\n4.82\nDrooped-nose flap\nC\nA\nA\nB\nB\nSplit flap\n34.13\nC\nSection A-A\n(enlarged)\n8°\nSection B-B\n(enlarged)\n60°\nFence\n1.99\nSection C-C\n(enlarged)\n29.00\n1.70\n18.76\n0.70 b/2\n0.55 b/2\n1.70\n0.13c\n0.12c\nE\nExtensible leading-edge flap\nD\nD\n1.70\n0.18c\nB\nB\nSplit flap\nFence\n1.99\nSection D-D\n(enlarged)\n3°\n3°\nSection E-E\n(enlarged)\nNACA\n\nFigure 2.- Details of high-lift and stall-control devices on a 42° sweptback wing.", "timestamp": "2026-07-22T04:11:09.440092+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 58, "total_pages": 82, "image_filename": "19930085551_p58.jpg", "text": "NACA RM No. L8K30\n57\n\nYawing velocity,\nLeft 10 deg/sec Right\nPitching velocity,\nDown 10 rad/sec Up\nYaw\nPitch\n\nPull\nRight\nControl force, lb\nPush\nLeft\nAileron\nElevator\nRudder\n\nControl position,\nLeft deg Right\nDown Up\nAileron\nElevator\nRudder\n\nAltitude, ft\n8900\n8700\n\nIndicated\nairspeed, mph\n160\n120\nNACA\n\n0 4 8 12 16 20 24 28\nTime, sec\n\n(h) Approach condition; flaps 20°; gear down; power for level flight;\naltitude controlled; left turn; 140 miles per hour.\n\nFigure 14.- Continued.", "timestamp": "2026-07-22T04:11:09.609644+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 59, "total_pages": 118, "image_filename": "19930085838_p59.jpg", "text": "NACA RM No. L9B23\n57\n\n<!-- Image (145, 126, 877, 805) -->\n\n(a) $\\delta_f = 0^\\circ$.\nFigure 8.- Hinge-moment characteristics of a straight-sided Frise aileron on the approximately 17.7-percent-chord thick NACA 7-series-type airfoil with double slotted flap and flip. Aileron balance, $0.408c_a$; $R = 6 \\times 10^6$ (approx.)", "timestamp": "2026-07-22T04:11:17.612975+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 14, "total_pages": 32, "image_filename": "19930085982_p14.jpg", "text": "```markdown\n12\nNACA RM E5E13\n\nradial flow was not, however, sufficient to satisfy completely the\nrequirement of simple-radial equilibrium. As was expected, the\nactual flow through the rotor lay somewhere between the assumption\nof no change in axial-velocity distribution over the rotor row and\nthe assumption of the existence of simple-radial equilibrium at the\nrotor outlet.\n\nRotor-Row Performance\n\nRadial distribution of energy addition. - A good method to\nminimize mixing losses downstream of a compressor rotor and to\nfacilitate staging is to keep the total enthalpy from hub to tip\nconstant. This compressor was therefore designed for a vortex\naddition by the rotor with constant enthalpy addition across the\npassage. Figure 8 is a plot of the rotor enthalpy addition\n$$ \\frac{U \\Delta V_\\theta}{\\theta} $$\nagainst radius ratio at the rotor outlet $r/r_t$ for the\nrange of weight flows covered at three-fourths design speed. At\nthe high weight flows, the enthalpy addition at the blade tip is\nextremely low and increases towards the hub. As the weight flow is\ndecreased, the enthalpy addition at the tip increases rapidly as\ncompared with that at the hub. At weight flows of approximately\n19 pounds per second, the energy addition is relatively uniform\nacross the annulus. The weight flows that have the most uniform\nenergy distribution are in the high-efficiency range. At the low\nweight flows, the enthalpy addition continuously decreases from\nthe tip to the hub. Several interacting factors determine the\nshape of the curves in figure 8. Briefly, the more predominate\nfactors are:\n\n1. The high stagger angles and low solidities at the blade\ntip result in a somewhat lower value of the slope of the curve of\nturning angle against angle of attack for the tip section than for\nthe hub section. This effect is offset by the fact that, in gen-\neral, a given change in weight flow causes a much larger change in\nangle of attack at the tip than at the hub.\n\n2. The high wheel speed at the rotor tip in conjunction with\nthe high stagger angles results in a larger change in energy addi-\ntion for a given change in turning angle at the tip than at the hub.\n\n3. The effects of radial equilibrium of static pressure on\nthe flow at the rotor outlet causes a decrease in axial-velocity\nratio at the tip as the flow is decreased, whereas near the hub the\n\n1131\n```", "timestamp": "2026-07-22T04:11:18.855408+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 32, "total_pages": 59, "image_filename": "19930085936_p32.jpg", "text": "NACA RM No. E9B03\n31\n\nAngle of attack\n(deg)\n$\\circ$ -15\n$\\square$ -6\n$\\diamond$ 0\n$\\triangle$ 12\n$\\nabla$ 24\n\nPressure coefficient, $C_p$\n.10\n0\n-.10\n-.20\n-.30\n\nDistance from tip, x/L\n0 .2 .4 .6 .8 1.0\n\nNACA\n\n(e) $\\theta = 270^\\circ$ longitudinal plane.\n\nFigure 5. - Concluded. Pressure distributions along longitudinal planes at $0^\\circ$ yaw angle for range of angles of attack.", "timestamp": "2026-07-22T04:11:22.059578+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 43, "total_pages": 92, "image_filename": "19930085930_p43.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n42\n1.78\"\n2.00\"\n1.24\"\nNACA\nFigure 12.- Schematic setup for first variable-span model.\nNACA RM 19G07", "timestamp": "2026-07-22T04:11:22.970408+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 3, "total_pages": 20, "image_filename": "19930086060_p3.jpg", "text": "NACA RM L9F02\nCONFIDENTIAL\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nRESEARCH MEMORANDUM\nFLIGHT INVESTIGATION AT HIGH-SUBSONIC, TRANSONIC,\nAND SUPERSONIC SPEEDS TO DETERMINE ZERO-LIFT DRAG OF\nBODIES OF REVOLUTION HAVING FINENESS RATIO OF 6.04\nAND VARYING POSITIONS OF MAXIMUM DIAMETER\nBy Ellis R. Katz\nSUMMARY\nFlight investigation of rocket-powered models was performed at\nhigh-subsonic, transonic, and supersonic speeds to determine the zero-\nlift drag of fin-stabilized bodies of revolution differing only in\nposition of maximum diameter. The parabolic bodies were of\n6.04 fineness ratio and had cut-off sterns with equal base area for all\nmodels. Pressure and drag data are reported at maximum diameter\nstations of 20, 40, and 60 percent of the body length.\nAt supersonic speeds the 60-percent station resulted in the least\ndrag, and theoretical estimations at $M = 1.4$ indicated that the\n60-percent position may be nearly optimum. At transonic speeds, equal\ndrag resulted at the 40-percent and 60-percent stations and at subsonic\nspeeds the position of maximum diameter had no effect.\nINTRODUCTION\nPractical flight at transonic and supersonic speeds has dictated\nthe tendency toward large wing loadings for aircraft configurations.\nIn addition, the use of thinner and stronger wings has resulted in\naccommodating greater fuel and fixed-equipment loads within the fuselage\nof the aircraft. The sum of the above two effects has been to increase\nconsiderably the size of the fuselage of the high-speed aircraft\nrelative to the size of the wing. Thus, it is clear that the fuselage\ndrag of supersonic aircraft, which is of the order of 30 percent of the\nCONFIDENTIAL", "timestamp": "2026-07-22T04:11:31.521932+00:00"}
{"citation_id": "19930085957", "source_url": "https://ntrs.nasa.gov/api/citations/19930085957/downloads/19930085957.pdf", "page_number": 26, "total_pages": 27, "image_filename": "19930085957_p26.jpg", "text": "NACA RM L9B02\n25\n\n<!-- Image (258, 78, 819, 800) -->\n\n(a) Split flaps off; R = $4.7 \\times 10^6$.\n\nFigure 13.- Effect of roughness on the stalling characteristics of the $42^\\circ$ sweptback wing with $0.575 \\frac{b}{2}$ leading-edge flaps with and without half-span split flaps. $\\delta_F = 60^\\circ$.", "timestamp": "2026-07-22T04:11:32.026510+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 15, "total_pages": 36, "image_filename": "19930086003_p15.jpg", "text": "CONFIDENTIAL\n\nWing chord plane\nextended at CC = 0°\n\nFloating-tail geometry\nArea (Twice semispan) 0.0178 sq ft\nAspect ratio 4.0\nTaper ratio 0.60\n\n0.25 c of model\n4.53\n0.25-Chord line\n45°\n0.60\nB\n0.80\n0.76\n1.60\nB\nBump surface\n1.0\nc_t\n1/8 Diameter\nPivot center\n\nSection B-B\n1/8\n\nNACA\n\n0 1 2\nScale, inches\n\nCONFIDENTIAL\n\nFigure 3.- Details of free-floating tails used in surveys behind model with 45° sweptback wing,\naspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section.\n\nNACA RM L9J08\n13", "timestamp": "2026-07-22T04:11:32.545134+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 13, "total_pages": 54, "image_filename": "19930086015_p13.jpg", "text": "```markdown\n12\nCONFIDENTIAL\nNACA RM A9E24\n\nTABLE I\nCoordinates of Nozzle Blocks\n\n[Figure: Diagram of Nozzle Blocks with dimensions 495.0, 462.0, 66.0, 546.0, 144.0, and labels L_u, D, L_1, H_u, H_1, Test Section, Slide Plane]\n\nDimensions in inches.\n\n| $L_1$ | $H_1$ | $L_u$ | $H_u$ | $L_u$ | $H_u$ |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| 0 | 144.000 | 0 | -6.000 | 269.840 | 6.281 |\n| 146.724 | 69.992 | *48.000 | -61.360 | 271.791 | 8.174 |\n| 149.724 | 64.254 | *136.000 | -29.976 | 277.379 | 9.991 |\n| 152.724 | 62.788 | 144.000 | -35.842 | 283.203 | 11.734 |\n| 155.724 | 61.243 | 150.000 | -39.488 | 289.068 | 13.404 |\n| 158.724 | 59.766 | 156.000 | -34.724 | 294.998 | 15.002 |\n| 161.724 | 58.227 | 162.000 | -13.726 | 300.887 | 16.528 |\n| 164.724 | 56.743 | 168.000 | -38.440 | 306.892 | 17.984 |\n| 167.724 | 55.275 | 174.000 | -30.876 | 312.849 | 19.370 |\n| 170.724 | 53.809 | 180.000 | -69.092 | 318.160 | 20.534 |\n| 173.304 | 52.620 | 186.000 | -66.991 | 324.035 | 21.762 |\n| 176.043 | 51.395 | 192.000 | -64.713 | 329.924 | 22.929 |\n| 178.806 | 50.223 | 198.000 | -62.243 | 335.829 | 24.036 |\n| 181.593 | 49.103 | 200.000 | -61.381 | 341.738 | 25.086 |\n| 184.403 | 48.034 | 202.000 | -60.501 | 347.665 | 26.080 |\n| 187.237 | 47.017 | 204.000 | -19.605 | 353.601 | 27.022 |\n| 190.094 | 46.051 | 206.000 | -18.693 | 359.548 | 27.912 |\n| 192.975 | 45.135 | 208.000 | -17.765 | 365.506 | 28.764 |\n| 195.877 | 44.267 | 210.000 | -16.821 | 371.472 | 29.567 |\n| 198.803 | 43.449 | 212.000 | -15.865 | 377.449 | 30.326 |\n| 201.750 | 42.677 | 213.000 | -15.385 | 383.436 | 31.044 |\n| 204.718 | 41.953 | 213.899 | -14.953 | 389.433 | 31.723 |\n| 207.708 | 41.276 | 214.800 | -14.522 | 395.439 | 32.363 |\n| 210.719 | 40.643 | 215.700 | -14.092 | 401.451 | 32.960 |\n| 213.750 | 40.056 | 216.605 | -13.663 | 407.478 | 33.512 |\n| 217.600 | 39.378 | 217.509 | -13.235 | 413.514 | 34.016 |\n| 220.570 | 38.956 | 218.414 | -12.808 | 419.560 | 34.471 |\n| 223.548 | 38.582 | 221.133 | -11.539 | 425.617 | 34.874 |\n| 226.528 | 38.231 | 226.596 | -9.056 | 431.683 | 35.221 |\n| 229.513 | 37.909 | 232.096 | -6.642 | 437.762 | 35.504 |\n| 232.501 | 37.610 | 237.630 | -4.303 | 443.852 | 35.716 |\n| 235.494 | 37.336 | 243.198 | -2.037 | 449.954 | 35.860 |\n| 238.489 | 37.086 | 248.804 | .156 | 456.070 | 35.948 |\n| 241.487 | 36.860 | 254.446 | 2.273 | 462.000 | 36.000 |\n| 244.490 | 36.656 | 260.124 | 4.314 | 528.000 | 36.439 |\n| 247.496 | 36.476 | | | | |\n| 250.504 | 36.311 | | | | |\n| 253.517 | 36.193 | | | | |\n| 256.534 | 36.094 | | | | |\n| 259.556 | 36.027 | | | | |\n| 262.583 | 35.998 | | | | |\n| 269.000 | 36.000 | | | | |\n| 474.000 | 37.119 | | | | |\n| 500.366 | 38.732 | | | | |\n| 546.000 | 44.000 | | | | |\n\nCircular arc of radius 284.382 inches between (*) stations\n\nNACA\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T04:11:34.828658+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 19, "total_pages": 67, "image_filename": "19930085965_p19.jpg", "text": "18\nNACA RM E9E06\n\nThe fifth experiment was made with the synthesized pulsating unidirectional current. As illustrated by the curve in figure 11, the two currents were apportioned to subject the blade to a magnetomotive force that varied from 0 to $2\\sqrt{2}$ times the root-mean-square value of alternating magnetomotive force at each experimental condition. The Armco iron sample was used in this investigation and the total air-gap spacing was 0.04 inch.\n\nA 10-turn search coil was placed around the center of the sample blade in all experiments for the purpose of measuring the voltage and viewing the voltage wave shape on a cathode-ray oscilloscope. Ten turns were used to obtain sufficient signal voltage to make the hum voltage and other spurious voltages negligible in comparison.\n\nCALCULATION OF AMPERE TURNS\n\nThe correlation between the analytical equations and the experimental results depends on several new calculations that must be made in addition to computations using the equations already derived.\n\nThe blade was positioned in the experimental rig (fig. 10) to make the air gaps at each end of the blade approximately equal. Figure 12 shows the geometry used in calculating the permeances at each blade end. No attempt was made to draw figure 12 to scale or to make the relative sizes of the various volumes correct. (Volumes are designated by numbers 1 to 8.) The purpose of the figure is to give the geometric shapes of the various volumes to illustrate the application of formulas from reference 6. The suggestion given in this reference was followed in limiting the fringing-flux calculations to the entire space immediately surrounding the pole faces and not at a great distance along the blade (h+t+g was limited to approximately 0.75 in., where h, t, and g are dimensions, as shown in fig. 12). The geometry of figure 12 is based on the two propositions that the flux lines are normal to the metal surfaces and that such surfaces are equipotential. These propositions are obviously assumptions, because the magnetomotive-force drop along 0.75 inch of blade is appreciable.\n\nThe formulas for the permeances of the various volumes designated by the numbers in figure 12 follow for convenience. The air gap directly under the blade was assumed to have no flux, except for a volume equivalent to the depth of penetration. This permeance $P_1$ was therefore taken to be equal to the product of the depth of penetration $l_T$ and the circumference of the blade cross section, divided by the distance g:", "timestamp": "2026-07-22T04:11:34.845263+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 43, "total_pages": 104, "image_filename": "19930085842_p43.jpg", "text": "NACA RM L9C29\n39\n\n[Figure: A graph plotting Lift coefficient, $C_L$ (y-axis) against Angle of attack, $\\alpha$, deg (x-axis). The y-axis ranges from 0 to 1.4. The x-axis ranges from 20 to 50. Five curves are plotted, labeled as follows:\n- configuration 7\n- configuration 9\n- configuration 5\n- configuration 4\n- configuration 8\nThe curves generally rise from an angle of attack of 20 degrees, peak between 30 and 40 degrees, and then decline. The National Advisory Committee for Aeronautics logo is visible in the bottom right corner of the graph.]\n\nFigure 13.- Comparison of maximum lift coefficients obtained with five model configurations. (See table II.) Propellers removed.", "timestamp": "2026-07-22T04:11:38.750023+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 36, "total_pages": 47, "image_filename": "19930085918_p36.jpg", "text": "NACA RM A9D29\n35\n\n-112\n-108\n-104\n-100\n-96\n-92\n-88\n-84\n-80\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\n-44\n-40\n-36\n-32\n-28\n-24\n-20\n-16\n-12\n-8\n-4\n0\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\nPressure coefficient, P\n\nChordwise station, x/c\n.2\n.6\n.8\n1.0\n\n(b) $\\alpha=20.7^\\circ$\n\nNACA\n\nFigure 9. -Continued.", "timestamp": "2026-07-22T04:11:40.401241+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 29, "total_pages": 50, "image_filename": "19930085952_p29.jpg", "text": "28\nNACA RM L9C24\n\nNeutral point location - percent $\\bar{c}$\n$C_L$\nArticulated propellers\nRigid propellers\nNormal c.g. location\nNACA\n\nFigure 6.- Effect of propeller articulation on the stick-fixed longitudinal stability of the model for simulated full-power operation. $\\delta_F = 0^\\circ$.\n\n$C_m$\n$(C_{h_a}=0)$\n$C_L$\nArticulated propellers\nRigid propellers\nNACA\n\nFigure 7.- Effect of propeller articulation on the variation of $C_m$ with $C_L$ stick free for simulated full-power operation. $\\delta_F = 0^\\circ$.", "timestamp": "2026-07-22T04:11:42.152711+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 21, "total_pages": 43, "image_filename": "19930085958_p21.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:11:42.368517+00:00"}
{"citation_id": "19930085843", "source_url": "https://ntrs.nasa.gov/api/citations/19930085843/downloads/19930085843.pdf", "page_number": 61, "total_pages": 65, "image_filename": "19930085843_p61.jpg", "text": "NACA RM L59C31\n\nPitching-moment coefficient, $C_{m_{c/4}}=0$\n\nSemispan model, wing-fuselage\nSting model, wing-fuselage\nWing-flow model, wing-fuselage\n\n.04\n0\n-.04\n\n.4 .5 .6 .7 .8 .9 1.0 1.1\n\nMach number, M\n\n(b) Vertical fins off.\nFigure 19.- Concluded.\n\nNACA\n\n59", "timestamp": "2026-07-22T04:11:46.790366+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 59, "total_pages": 82, "image_filename": "19930085551_p59.jpg", "text": "58\nNACA RM No. L8K30\n\n<!-- Image (151, 120, 785, 842) -->\n\n(i) Final approach condition; flaps full down; gear down; power for level flight; no attempt to control altitude; right turn; 120 miles per hour.\n\nFigure 14.- Continued.", "timestamp": "2026-07-22T04:11:50.550146+00:00"}
{"citation_id": "19930085917", "source_url": "https://ntrs.nasa.gov/api/citations/19930085917/downloads/19930085917.pdf", "page_number": 38, "total_pages": 38, "image_filename": "19930085917_p38.jpg", "text": "NACA RM No. L9C18 CONFIDENTIAL 37\n\nM = 0.65 M = 0.65\n\nM = 0.78 M = 0.79\n\nM = 0.82 M = 0.82\n\nNACA 64₁-012 airfoil NACA 64₁A012 airfoil\n\n(c) α = 8°.\n\nFigure 6.- Concluded.\nCONFIDENTIAL\n\n[Figure: NACA logo]\nL-59837", "timestamp": "2026-07-22T04:11:53.810245+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 93, "total_pages": 122, "image_filename": "19930082090_p93.jpg", "text": "NACA TN No. 1455\n91\n\n[Figure: A black and white photograph of a cylindrical metal object with numerous pins on its surface, resting on a wooden table.]\n\nFigure 45.- Pin heat exchanger L.", "timestamp": "2026-07-22T04:11:55.018514+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 10, "total_pages": 34, "image_filename": "19930086022_p10.jpg", "text": "8\nNACA RM L9E24\n\ndeflections were lower than those obtained for the corresponding negative deflections. This effect increased with angle of attack. The addition of upper-surface fences appeared to have little effect on improving the flap-deflected aileron effectiveness. The reduced aileron effectiveness at large positive aileron deflections for the wing with high-lift and stall controls deflected was noted in reference 2 where aileron tests were made on a wing of approximately the same plan form but incorporating biconvex airfoil sections. Comprehensive information from which an exact reason could be obtained for the loss in aileron effectiveness was not available; however, it may be due in part to the existence of some degree of flow separation in the region of the aileron as is indicated by the lift and moment characteristics obtained.\n\nA comparison of the variation of $C_{l_\\delta}$ with angle of attack for various model configurations (fig. 9) indicates that the values of $C_{l_\\delta}$ (at $\\delta_a = 0^\\circ$) were approximately the same for all model configurations investigated up to an angle of attack of approximately $10^\\circ$ where $C_{l_\\delta}$ for the plain wing started to decrease. The value of $C_{l_\\delta}$ for the wing with leading- and trailing-edge flaps deflected was slightly greater throughout the higher angle-of-attack range. For the condition with leading- and trailing-edge high-lift devices and upper-surface fences, $C_{l_\\delta}$ started to decrease at an angle of attack of $12^\\circ$ and at an angle of attack of $16^\\circ$ was 50 percent lower than that obtained for the plain wing. The experimental value of $C_{l_\\delta}$ obtained for the plain wing at zero angle of attack was 0.00105. A value of $C_{l_\\delta}$ of 0.00103 was calculated for the same test condition by the method of reference 6 and is within 2 percent of the experimental value.\n\nAt zero angle of attack, favorable yawing moments were obtained for all model configurations (fig. 7); however, as the angle of attack was increased, the yawing moments for all configurations became adverse.\n\nPitching-moment characteristics.- The increments of pitching moment due to positive and negative aileron deflections were approximately equal for the plain wing but were unequal in the case of the wing with high-lift and stall-control devices deflected (figs. 4 to 6). As in the case of the rolling-moment characteristics, the large positive aileron deflections produced a smaller increment in pitching moment than corresponding negative deflections. Based upon the results of horizontal-tail tests of reference 7 and assuming that the elevator effectiveness would be 50 percent of the stabilizer effectiveness, it", "timestamp": "2026-07-22T04:11:55.289277+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 60, "total_pages": 118, "image_filename": "19930085838_p60.jpg", "text": "58\nNACA RM No. L9B23\n\n[Figure: A graph plotting Aileron section hinge-moment coefficient, $C_{h_a}$, against Section angle of attack, $\\alpha_o$, deg. The graph contains multiple data series with different symbols and legends. A NACA logo is present in the bottom right corner of the plot area.]\n\nAileron section hinge-moment coefficient, $C_{h_a}$\n\n$\\delta_a$ (deg) | $\\delta_t$ (deg)\n---|---\n0 | 0\n0 | -2\n\n$\\delta_a$ (deg) | $\\delta_t$ (deg)\n---|---\n0 | -5\n0 | -10\n\n$\\delta_a$ (deg) | $\\delta_t$ (deg)\n---|---\n0 | -15\n0 | -20\n\nSection angle of attack, $\\alpha_o$, deg\n\n(b) $\\delta_f = 20^\\circ$.\nFigure 8.- Continued.", "timestamp": "2026-07-22T04:11:57.577570+00:00"}
{"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 74, "total_pages": 85, "image_filename": "19930085529_p74.jpg", "text": "```markdown\nNACA RM No. L8A30a\n73\n\nTABLE 67\n$$\n[\\Lambda = -30^\\circ, \\delta_{fl} = 5.0^\\circ, \\alpha = 7^\\circ]\n$$\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** | **0.925** | **0.96** | | | **0.60** | **0.80** | **0.85** | **0.89** | **0.925** | **0.96** |\n| A1 | 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 | -0.016 | -0.079 | -0.125 | -0.165 | -0.216 | -0.279 |\n| 6 | 50.0 | -0.507 | -0.597 | -0.596 | -0.570 | -0.589 | -0.750 | 91 | 62.5 | -0.015 | -0.080 | -0.129 | -0.165 | -0.195 | -0.268 |\n| 7 | 59.0 | -0.436 | -0.526 | -0.526 | -0.506 | -0.514 | -0.661 | 92 | 72.5 | -0.012 | -0.065 | -0.117 | -0.137 | -0.131 | -0.183 |\n| 8 | 67.5 | -0.338 | -0.538 | -0.561 | -0.593 | -0.604 | -0.671 | 93 | 84.0 | - - | - - | - - | - - | - - | - - |\n| 9 | 77.5 | - - | - - | - - | - - | - - | - - | 94 | 94.0 | - - | - - | - - | - - | - - | - - |\n| 10 | 87.5 | - - | - - | - - | - - | - - | - - | | | | | | | | |\n| 11 | 96.0 | - - | - - | - - | - - | - - | - - | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| B12 | 2.0 | -0.633 | -0.546 | -0.546 | -0.546 | -1.007 | -0.997 | 95 | 3.0 | -0.746 | -0.751 | -0.718 | -0.709 | -0.698 | -0.698 |\n| 13 | 6.0 | -0.663 | -0.558 | -0.562 | -0.554 | -0.565 | -0.849 | 96 | 10.0 | -0.146 | -0.110 | -0.199 | -0.374 | -0.368 | -0.378 |\n| 14 | 15.0 | -0.677 | -0.556 | -0.560 | -0.539 | -0.744 | -0.892 | 97 | 25.0 | -0.117 | -0.096 | -0.070 | -0.048 | -0.018 | -0.055 |\n| 15 | 27.5 | -0.696 | -0.558 | -0.561 | -0.554 | -0.791 | -0.813 | 98 | 41.0 | -0.012 | -0.056 | -0.091 | -0.127 | -0.154 | -0.144 |\n| 16 | 40.0 | -0.663 | -0.618 | -0.626 | -0.626 | -0.585 | -0.701 | 99 | 52.5 | -0.047 | -0.096 | -0.135 | -0.166 | -0.199 | -0.230 |\n| 17 | 50.0 | -0.596 | -0.626 | -0.611 | -0.770 | -0.777 | -0.751 | 100 | 62.5 | -0.011 | -0.084 | -0.120 | -0.148 | -0.179 | -0.223 |\n| 18 | 59.0 | -0.507 | -0.609 | -0.617 | -0.776 | -0.766 | -0.633 | 101 | 72.5 | -0.004 | -0.042 | -0.072 | -0.088 | -0.099 | -0.132 |\n| 19 | 67.5 | -0.406 | -0.560 | -0.604 | -0.569 | -0.688 | -0.559 | 102 | 86.3 | -0.018 | -0.023 | -0.021 | -0.077 | -0.055 | -0.083 |\n| 20 | 77.5 | -0.304 | -0.478 | -0.509 | -0.450 | -0.450 | -0.597 | 103 | 94.6 | -0.041 | -0.043 | -0.076 | -0.049 | -0.063 | -0.075 |\n| 21 | 88.0 | -0.182 | -0.367 | -0.432 | -0.407 | -0.333 | -0.641 | | | | | | | | |\n| 22 | 96.5 | - - | - - | - - | - - | - - | - - | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| C23 | 2.0 | -0.329 | -0.389 | -0.244 | -0.041 | -0.881 | -0.748 | 104 | 3.0 | -0.727 | -0.722 | -0.701 | -0.687 | -0.681 | -0.684 |\n| 24 | 6.0 | -0.771 | -0.358 | -0.164 | -0.069 | -0.936 | -0.815 | 105 | 10.0 | -0.432 | -0.436 | -0.416 | -0.406 | -0.401 | -0.414 |\n| 25 | 15.0 | -0.991 | -0.040 | -0.075 | -0.037 | -0.933 | -0.867 | 106 | 25.0 | -0.176 | -0.150 | -0.123 | -0.117 | -0.135 | -0.154 |\n| 26 | 27.5 | -0.687 | -0.740 | -0.755 | -0.597 | -0.908 | -0.867 | 107 | 41.0 | -0.033 | -0.013 | -0.014 | -0.040 | -0.053 | -0.040 |\n| 27 | 40.0 | -0.556 | -0.570 | -0.563 | -0.563 | -0.863 | -0.847 | 108 | 52.5 | -0.048 | -0.048 | -0.081 | -0.119 | -0.143 | -0.142 |\n| 28 | 50.0 | -0.484 | -0.548 | -0.515 | -0.615 | -0.760 | -0.840 | 109 | 62.5 | -0.003 | -0.046 | -0.079 | -0.114 | -0.150 | -0.195 |\n| 29 | 59.0 | -0.409 | -0.436 | -0.436 | -0.418 | -0.747 | -0.817 | 110 | 72.5 | -0.024 | -0.001 | -0.011 | -0.059 | -0.096 | -0.098 |\n| 30 | 67.5 | -0.297 | -0.363 | -0.376 | -0.418 | -0.577 | -0.803 | 111 | 85.1 | -0.056 | -0.027 | -0.004 | -0.005 | -0.014 | -0.045 |\n| 31 | 77.5 | -0.206 | -0.323 | -0.359 | -0.406 | -0.509 | -0.768 | 112 | 94.6 | -0.080 | -0.040 | -0.019 | -0.004 | -0.019 | -0.045 |\n| 32 | 88.0 | -0.102 | -0.252 | -0.317 | -0.383 | -0.452 | -0.622 | | | | | | | | |\n| 33 | 96.5 | - - | - - | - - | - - | - - | - - | | | | | | | | |\n| | | | | | | | | | | | | | | | | |\n| D34 | 2.0 | -0.914 | -1.437 | -1.163 | -0.964 | -0.811 | -0.686 | 113 | 3.0 | -0.727 | -0.692 | -0.674 | -0.656 | -0.648 | -0.653 |\n| 35 | 15.0 | -0.962 | -1.311 | -1.122 | -0.910 | -0.855 | -0.786 | 114 | 10.0 | -0.437 | -0.436 | -0.421 | -0.408 | -0.407 | -0.417 |\n| 36 | 27.5 | -0.664 | -1.222 | -1.087 | -0.834 | -0.932 | -0.889 | 115 | 25.0 | -0.184 | -0.167 | -0.150 | -0.131 | -0.153 | -0.169 |\n| 37 | 40.0 | -0.563 | -0.814 | -1.036 | -1.004 | -0.975 | -0.882 | 116 | 41.0 | -0.056 | -0.046 | -0.053 | -0.064 | -0.014 | -0.024 |\n| 38 | 50.0 | -0.465 | -0.659 | -0.806 | -0.806 | -0.887 | -0.877 | 117 | 52.5 | -0.041 | -0.048 | -0.083 | -0.075 | -0.079 | -0.094 |\n| 39 | 59.0 | -0.372 | -0.470 | -0.703 | -0.696 | -0.903 | -0.845 | 118 | 62.5 | -0.014 | -0.001 | -0.033 | -0.072 | -0.087 | -0.084 |\n| 40 | 67.5 | -0.278 | -0.436 | -0.609 | -0.609 | -0.847 | -0.847 | 119 | 72.5 | -0.006 | -0.001 | -0.011 | -0.059 | -0.096 | -0.098 |\n| 41 | 77.5 | -0.156 | -0.135 | -0.138 | -0.311 | -0.558 | -0.741 | 120 | 87.4 | -0.056 | -0.111 | -0.066 | -0.023 | -0.007 | -0.013 |\n| 42 | 87.5 | -0.075 | -0.077 | -0.077 | -0.266 | -0.369 | -0.606 | 121 | 94.6 | -0.0", "timestamp": "2026-07-22T04:11:59.730586+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 44, "total_pages": 92, "image_filename": "19930085930_p44.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n\nNACA RM L9G07\n\n[Figure: Schematic diagram showing a curved wing-like structure with dashed outlines, dimensioned at 2.00\", labeled points A and A, and a cross-section labeled \"Section AA\". The NACA logo is present near the bottom center.]\n\nSection AA\n\nFigure 13.- Schematic setup for second variable-span model.\n\nUNCLASSIFIED\nCONFIDENTIAL\n\n43", "timestamp": "2026-07-22T04:12:01.042481+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 33, "total_pages": 59, "image_filename": "19930085936_p33.jpg", "text": "32\nNACA RM No. E9B03\n\n<!-- Image (149, 117, 852, 852) -->\n\n(a) $\\theta = 0^\\circ$ longitudinal plane.\nFigure 6. - Pressure distributions along longitudinal planes at $0^\\circ$ angle of attack for range of yaw angles.", "timestamp": "2026-07-22T04:12:01.042855+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 15, "total_pages": 32, "image_filename": "19930085982_p15.jpg", "text": "NACA RM E9E13\n\naxial-velocity ratio increases with decreasing flow. These velocity changes, when applied to the equivalent diagram, indicate an increase in angle of attack at the tip and a decrease at the hub.\n\nLarge variations in energy distribution can apparently result from changes in weight flow with blade designs of this type. It is possible that a poor energy distribution resulting from operation at a point other than design in one stage of a multistage compressor may be reduced in succeeding stages. For optimum performance, however, all stages of a multistage compressor should be carefully matched to avoid mixing losses resulting from operation other than design of one or more stages.\n\nVariation of turning angle with angle of attack. - Comparison of two-dimensional static-cascade and three-dimensional rotor data can be made only for the inner radii of the compressor because of the lack of cascade data for the combination of high stagger angles and low solidities encountered in this design at the outer radii. The comparison of experimental performance with design predictions, however, indicated that the turning angles obtained were considerably below those anticipated. In figure 9, the equivalent turning angle $\\Delta \\beta_{e}^{\\prime}$ is plotted against equivalent angle of attack for four radial positions; for comparison, the design values obtained from the equation\n\n$\\Delta \\beta^{\\prime}=K(\\alpha+8.28)$\n\nand including the arbitrary correction for solidity and stagger are shown. The turning angles obtained at all radii were from $4^{\\circ}$ to $9^{\\circ}$ below those predicted for straight-through flow. At the outer radii where the solidity is extremely low (0.566 at the tip) and the stagger angles are high, the arbitrary correction for solidity was insufficient. In the region of the hub, the validity of the value of K and of $\\alpha_{1,0}$ used to determine the blade-angle setting has been verified by two-dimensional cascade data (reference 4). The discrepancy near the hub (station d) can, however, be partly explained by the radial flows due to the large hub taper. If a developed section of the blade is taken along an approximate streamline obtained by taking equal flow increments at the inlet and outlet of the blade row, the angle between the blade trailing edge and the blade chord will be reduced approximately $5^{\\circ}$ as a result of the decrease in effective camber due to hub taper and blade twist. At the intermediate stations, the underturning may result from the combined effects of insufficient allowance for decreased solidity, increased stagger, radial flows resulting from hub taper, and unknown induced velocity effects. For large hub tapers, the blade sections", "timestamp": "2026-07-22T04:12:10.117752+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 16, "total_pages": 36, "image_filename": "19930086003_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:12:11.940724+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 14, "total_pages": 54, "image_filename": "19930086015_p14.jpg", "text": "```markdown\nCONFIDENTIAL\n\n1 control panel\n2 test section\n3 cooling coils\n4 cooling tower\n5 compressor\n6 drive motors\n\n7 dry air storage tank\n8 vacuum pumps and compressors\n\n[Figure: Drawing of the Ames 6- by 6-foot supersonic wind tunnel.]\n\nFigure 1.- Drawing of the Ames 6- by 6-foot supersonic wind tunnel.\n\nNACA\nA-13000.1\n\nNACA RM A9E24\n\nCONFIDENTIAL\n\n13\n```", "timestamp": "2026-07-22T04:12:13.917652+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 37, "total_pages": 47, "image_filename": "19930085918_p37.jpg", "text": "36\nNACA RM A9D29\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nPressure coefficient $P$\nChordwise station, $x/c$\nSpanwise station, $2y/b$\n\n(f) $\\alpha=24.7^\\circ$\n\nFigure 9.—Continued.", "timestamp": "2026-07-22T04:12:17.206045+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 44, "total_pages": 104, "image_filename": "19930085842_p44.jpg", "text": "40\n\n[Graph: A plot showing three curves ($C_L$, $C_D$, $C_m$) against Angle of attack, $\\alpha$, deg. The x-axis ranges from -10 to 90. The left y-axis ranges from -0.2 to 1.4 for Lift coefficient, $C_L$, and Pitching-moment coefficient, $C_m$. The right y-axis ranges from 0 to 1.4 for Drag coefficient, $C_D$. The curves are labeled $C_L$, $C_D$, and $C_m$.]\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 14.- Variation of $C_L$, $C_D$, and $C_m$ with $\\alpha$ of a $\\frac{1}{3}$-scale model of the airplane. Basic model configuration; propellers removed.\n\nNACA RM L9C29", "timestamp": "2026-07-22T04:12:18.916796+00:00"}
{"citation_id": "19930085588", "source_url": "https://ntrs.nasa.gov/api/citations/19930085588/downloads/19930085588.pdf", "page_number": 45, "total_pages": 51, "image_filename": "19930085588_p45.jpg", "text": "44\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:12:19.464835+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 30, "total_pages": 50, "image_filename": "19930085952_p30.jpg", "text": "```markdown\nNACA RM L9C24\n29\n\n<!-- Image (275, 116, 746, 567) -->\n\nFigure 8.- Effect of propeller articulation on the variation of ailavator deflection required for trim with lift coefficient for simulated full-power operation. $\\delta_F = 0^\\circ$.\n\n<!-- Image (299, 669, 711, 896) -->\n\nFigure 9.- Effect of propeller articulation on the variation of $C_{h\\delta_a}$ with $C_L$ for simulated full-power operation. $\\delta_F = 0^\\circ$.\n```", "timestamp": "2026-07-22T04:12:20.447473+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 4, "total_pages": 20, "image_filename": "19930086060_p4.jpg", "text": "```markdown\n2\nCONFIDENTIAL\nNACA RM L9F02\n\ntotal drag for present transonic configurations, is a major factor to\nbe reckoned with in the quest for higher speeds. In order to\ninvestigate and clarify the phenomena of the fuselage drag rise\nassociated with transonic and supersonic speeds, the NACA is conducting\na series of flight investigations on bodies of revolution differing in\nfineness ratio and position of maximum diameter. The tests are\nconducted by means of rocket-propelled models at the Pilotless Aircraft\nResearch Station, Wallops Island, Va. The preliminary investigation\nis presented in this paper and compares experimental and theoretical\ndrag results for fin-stabilized bodies of 6.04 fineness ratio having\nmaximum-diameter stations at 20, 40, and 60 percent of the body length.\n\nThe Mach number range of 0.6 to 1.85 corresponds to a Reynolds\nnumber range of $11 \\times 10^6$ to $52 \\times 10^6$ based on body length.\n\nMODELS AND TESTS\n\nThe general arrangement of the test vehicles is shown in figure 1\nand a photograph of the test configurations is shown as figure 2. The\nprofiles of all the wooden bodies are described by parabolic arcs\ngenerated at the positions of maximum diameter. The equations\ndescribing the profiles of the bodies are given in figure 3. In all\ncases the frontal area (0.307 sq ft), base area (0.0586 sq ft), and\nlength (3.77 ft) remain constant.\n\nAll models were stabilized by three 45° sweptback fins of\n1.69 square feet total exposed area. The dorsal fins were of\n0.0278 thickness ratio in the streamwise direction and so located that\nthe trailing edge of the fins always intersected the body at 9.47 per-\ncent of the body length forward of the model base.\n\nA two-stage propulsion system was employed utilizing a shortened\n3.25-inch-diameter Mk.7 aircraft rocket motor as the sustainer unit\nand a 5-inch HVAR motor for the booster unit. The booster unit was\nstabilized by four fins and was attached to the sustainer motor by\nmeans of a nozzle plug adapter.\n\nData were obtained by the standard drag technique as used in\nreference 1. The technique utilizes a CW Doppler velocimeter, located\nat the launching site, for the purposes of measuring the deceleration\nof the model due to gravity and drag. The data are measured throughout\nthe first 10 or 12 seconds of coasting flight (after burnout of the\nsustainer rocket motor) during which the flight path is virtually a\nstraight line from the point of launching. Atmospheric conditions are\nrecorded by means of radiosonde observations. A more complete\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T04:12:20.634064+00:00"}
{"citation_id": "19930085843", "source_url": "https://ntrs.nasa.gov/api/citations/19930085843/downloads/19930085843.pdf", "page_number": 62, "total_pages": 65, "image_filename": "19930085843_p62.jpg", "text": "```markdown\n60\nNACA RM L9031\n\n<!-- Image (173, 119, 824, 822) -->\n\nFigure 20.- A comparison of pitching-moment coefficients obtained by three different test methods on models of a tailless airplane. Vertical fins on; $\\delta_a = 0^\\circ$.\n```", "timestamp": "2026-07-22T04:12:23.682317+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 60, "total_pages": 82, "image_filename": "19930085551_p60.jpg", "text": "NACA RM No. L8K30\n59\n\nSideslip angle, deg\nLeft\nRight\n0\n10\n\nYawing velocity, deg/sec\nLeft\nRight\nPitching velocity, rad/sec\nUp\nDown\n-10\n0\n10\n-0.04\n0\n0.04\nPitch\nYaw\n\nControl forces, lb\nPush\nPull\nLeft\nRight\n-40\n0\n40\n80\n120\nRudder\nElevator\nAileron\n\nControl position, deg\nLeft\nRight\nDown\nUp\n-10\n0\n10\nRudder\nElevator\nAileron\n\nAltitude, ft\n6000\n6200\n\nIndicated airspeed, mph\n100\n140\nNACA\n\n0\n4\n8\n12\n16\n20\n24\n28\n32\nTime, sec\n\n(j) Final approach condition; flaps full down; gear down; power for level flight; altitude controlled; right turn; 120 miles per hour.\n\nFigure 14.— Concluded.", "timestamp": "2026-07-22T04:12:28.030910+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 94, "total_pages": 122, "image_filename": "19930082090_p94.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:12:29.981322+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 45, "total_pages": 92, "image_filename": "19930085930_p45.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:12:36.280315+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 61, "total_pages": 118, "image_filename": "19930085838_p61.jpg", "text": "NACA RM No. L9B23\n\nAileron section hinge-moment coefficient, $C_{h_a}$\n\nSection angle of attack, $\\alpha_o$, deg\n\n$\\delta_t = 0^\\circ$\n\n$\\delta_a$ (deg)\n\n-20\n\n-15\n\n-10\n\n0\n\n5\n\n10\n\n15\n\n(c) $\\delta_f = 25^\\circ$.\n\nFigure 8.- Continued.\n\n[Figure: Graph showing Aileron section hinge-moment coefficient vs. Section angle of attack for various $\\delta_a$ values, with NACA logo]\n\n59", "timestamp": "2026-07-22T04:12:36.487180+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 11, "total_pages": 34, "image_filename": "19930086022_p11.jpg", "text": "NACA RM L9E24\n\nis estimated that slightly less than $2^\\circ$ elevator travel would be necessary to counteract the pitching moment resulting from a total aileron deflection of $50^\\circ$ for the flap-deflected configurations throughout the angle-of-attack range.\n\nHinge-moment characteristics.— A representative plot of the variation of aileron hinge moment with aileron deflection is shown in figure 10. The effect of high-lift and stall-control devices on the aileron hinge-moment parameters, $C_{h\\delta}$, $C_{h\\alpha}$, $F_{R\\delta}$, and $F_{R\\alpha}$, is shown in figure 9. No appreciable effect on $C_{h\\delta}$ was noted when leading- and trailing-edge flaps were deflected; however, the high-lift devices caused a decrease in $C_{h\\alpha}$ in the higher angle-of-attack range. Except for very low angles of attack, the addition of leading- and trailing-edge flaps increased the value of $F_{R\\delta}$. In the case of $F_{R\\alpha}$, the addition of high-lift devices reduced the values of $F_{R\\alpha}$ in the higher angle-of-attack range and thereby reduced the abrupt increases in $F_{R\\alpha}$ obtained for the plain wing. In the lower angle-of-attack range the addition of flaps served to increase the values of $F_{R\\alpha}$. The addition of upper-surface fences produced variations of the aileron hinge-moment parameters which were more uniform throughout the angle-of-attack range.\n\nThe variations of $C_{h\\alpha}/C_{h\\delta}$ with angle of attack for the various model configurations investigated are presented in figure 11. Large positive values of $C_{h\\alpha}/C_{h\\delta}$ were obtained in the high angle-of-attack range for the plain wing and the wing equipped with leading- and trailing-edge flaps. The addition of upper-surface fences to the latter configuration greatly reduced the large values of $C_{h\\alpha}/C_{h\\delta}$ obtained in the high angle-of-attack range.\n\nBased upon the analysis of reference 8, values of $C_{h\\alpha}/C_{h\\delta}$ in excess of 2.0 are likely to result in large values of the ratio of peak force obtained at the initiation of a roll to the steady force in the roll with the possibility of objectionably high stick forces during the rapid initiation and reversal of an aileron roll. It can be seen from figure 11 that for the plain wing and the wing equipped with leading- and trailing-edge flaps an analysis such as presented in reference 8 would be necessary to insure against the possibility of obtaining excessive stick forces in any particular design.", "timestamp": "2026-07-22T04:12:38.793420+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 1, "total_pages": 42, "image_filename": "19930086078_p1.jpg", "text": "NACA RM L9H04\n\nFILE COPY\nNO. 6\n\nCONFIDENTIAL\n\nCopy 276\nRM L9H04\n\nNACA\n\nRESEARCH MEMORANDUM\n\nINVESTIGATION OF EXTENSIBLE WING-TIP AILERONS ON AN\nUNTAPERED SEMISPAN WING AT 0° AND 45° SWEEPBACK\n\nBy John R. Hagerman and William M. O'Hare\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE FILES\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information on this document may be imparted only to those in the military and naval service of the United States, appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nDATE 8-18-54\nAUTHORITY J.W. CROWLEY\nCHANGE # 2454\nF.E.T.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nSeptember 20, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:12:42.283024+00:00"}