PDF/ocr_dataset_1084.jsonl

{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 13, "total_pages": 24, "image_filename": "19930085991_p13.jpg", "text": "NACA RM L9I28\n\n11\n\nREFERENCES\n\n1. Scher, Stanley H.: Wind-Tunnel Investigation of the Stability of Jettisoned Nose Sections of the D-558 Airplane - Phases I and II. NACA RM L7K10, 1948.\n\n2. Bauer, C., and Kleckner, H.: Calculation of the Stability of the Detachable Nose Section of the MX-656 Fuselage. Rep. No. SM-13163, Douglas Aircraft Co., Inc., Jan. 8, 1948.\n\n3. Upson, Ralph H., and Klikoff, W. A.: Application of Practical Hydrodynamics to Airship Design. NACA Rep. 405, 1931.\n\n4. Tosti, Louis P.: Low-Speed Static Stability and Damping-in-Roll Characteristics of Some Swept and Unswept Low-Aspect-Ratio Wings. NACA TN 1468, 1947.\n\n5. Murray, Harry E.: Wind-Tunnel Investigation of End-Plate Effects of Horizontal Tails on a Vertical Tail Compared with Available Theory. NACA TN 1050, 1946.\n\n6. Lundstrom, Reginald R., and O'Kelly, Burke R.: Flight Investigation of the Jettisonable-Nose Method of Pilot Escape Using Rocket-Propelled Models. NACA RM L9D11, 1949.", "timestamp": "2026-07-22T07:00:29.919612+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 58, "total_pages": 92, "image_filename": "19930085870_p58.jpg", "text": "NACA RM No. L9D07\n59\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O CL, □ Cm\nWedge L.E. {△ CL, ◇ Cm\n.16\n.08\nCL\n0\n-.08\n-.16\n-.24\n.01\nCm\n0\n-.01\n\n.06\nElliptical L.E. {O CD, □ L/D\nWedge L.E. {△ CD, ◇ L/D\n.04\nCD\n.02\n0\n-8\n-6\n-4\n-2\n0\n2\n4\n6\n8\nα, deg\n6\n4\nL/D\n2\n0\n[NACA logo]\n\n(q) Wing 7.w=1.371; R=670,000.\nFigure 7 -- Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:00:34.976801+00:00"}
{"citation_id": "19930086083", "source_url": "https://ntrs.nasa.gov/api/citations/19930086083/downloads/19930086083.pdf", "page_number": 2, "total_pages": 48, "image_filename": "19930086083_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T07:00:35.585458+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 105, "total_pages": 114, "image_filename": "19930086061_p105.jpg", "text": "NACA RM L59J07\n\nWing leading edge\nQuarter chord\nLocal center of pressure, percent root chord\n$\\alpha$, deg $C_L$\n$\\circ$ 4.1 0.14\n$\\circ$ 8.1 0.32\n$\\diamond$ 14.1 0.54\n$\\frac{y}{b/2}$, percent\n(a) Angles of attack: 4.1°, 8.1°, 14.1°.\n\nWing leading edge\nQuarter chord\nLocal center of pressure, percent root chord\n$\\alpha$, deg $C_L$\n$\\triangle$ 24.1 0.85\n$\\square$ 32.1 1.06\n$\\square$ 36.1 1.16\nNACA\n$\\frac{y}{b/2}$, percent\n(b) Angles of attack: 24.1°, 32.1°, 36.1°.\n\nFigure 50.- Effect of angle of attack on the local center-of-pressure location of wing 2; $\\psi = 0^\\circ$.\n\n101", "timestamp": "2026-07-22T07:00:37.119951+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 10, "total_pages": 20, "image_filename": "19930085999_p10.jpg", "text": "8\nNACA RM E9I07\n\nits inherent damping and because of its tendency to change the\neffective length of a turbine blade, dependent on the individual\nfit of the base mount in the wheel dovetail.\n\nThe vibratory-stress levels were not excessively high in either\nthe loosely mounted or tightly mounted blades. As a corollary, the\nexcitation forces may be assumed to be low. The results of a recent\ninvestigation have shown that the magnitude of the excitation force\nproducing a vibration has an important bearing on whether the level\nof vibratory stress is affected by looseness of blade mounting\n(reference 5). Some of the data obtained during that investigation\nare presented in figure 7. The blade was a cantilever type, with a\nbase mount similar in design to that employed in axial-flow-compressor\nblades. The data presented are all for the first bending mode.\n\nIn figure 7, the stress in the tightly mounted blade is propor-\ntional to the excitation force regardless of turbine speed. For the\nloosely mounted blade, the stress is dependent on both turbine speed\nand exciting force. At 7950 rpm, for example, the blade is effec-\ntively tight until the excitation force is greater than 1 pound.\nGreater excitation forces apparently produce some rubbing action in\nthe mount and the stress for a given excitation force is therefore\nlower than in the case of the tightly mounted blade. At a speed of\n10,150 rpm, the centrifugal force is sufficient to hold the blade\ntight until the excitation force approaches 2 pounds; thereafter,\nthe stress for a given excitation force is lower than that induced\nin the tight blade but higher than the stress that occurred in the\nblade with the same degree of looseness when operated at a lower\nrotor speed.\n\nThe additional damping inherent in a loose blade mount may not\nbecome effective unless the exciting force is of sufficient magni-\ntude to offset the action of centrifugal force. A loose mounting\nwould conceivably be of value as a vibration damper.\n\nSUMMARY OF RESULTS\n\nHigh-temperature resistance-wire strain gages were used to\ndetermine the vibration characteristics of loosely mounted turbine\nblades. The experimental blades were excited into vibration by\nforces present in a turbojet engine during service operation. The\nblades were 4 inches long; one had a tip amplitude of 0.03 inch in\nthe plane of the turbine wheel; the other had a tip amplitude of\n0.06 inch. For both blades, the majority of the vibrations observed\nwere in first bending and first torsional modes. Some complex-mode", "timestamp": "2026-07-22T07:00:38.129346+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 37, "total_pages": 92, "image_filename": "19930085951_p37.jpg", "text": ".24\n.22\n.20\n.18\n.16\n.14\n.12\n.10\n.08\n.06\n.04\n.02\n0\n\nPower coefficient, $C_P$\n\n$\\beta_{0.75R}$ 20° 25° 30° 35° 40° 45°\n\nNACA\n\nAdvance ratio, J\n\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0 3.2 3.4 3.6 3.8\n\n(b) Power coefficient.\n\nFigure 11.— Continued. Rotational speed, 1600 rpm.\n\nCONFIDENTIAL\n\nNACA RM L9D29\n\n35", "timestamp": "2026-07-22T07:00:40.149750+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 25, "total_pages": 30, "image_filename": "19930085970_p25.jpg", "text": "NACA RM A59E09\nCONFIDENTIAL\n\nM\n0 0.50\n0 0.70\n◇ 0.80\n△ 0.90\n△ 0.95\n△ 1.09\n△ 1.14\n△ 1.24\n△ 1.51\n\nLift coefficient, $C_L$\n.6\n.4\n.2\n0\n-.2\n\nDrag coefficient, $C_D$ (for M = .50)\n0\n.02\n.04\n.06\n\n[Figure: Graph showing variations of drag coefficient with lift coefficient at various test Mach numbers, with data points marked by different symbols corresponding to Mach numbers from 0.50 to 1.51.]\n\nFigure 7.— Variations of drag coefficient with lift coefficient at the various test Mach numbers.\n\nCONFIDENTIAL\n23", "timestamp": "2026-07-22T07:00:42.329597+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 24, "total_pages": 25, "image_filename": "19930085979_p24.jpg", "text": "NACA RM E9E12\n\n[Figure: A schematic diagram showing a curved duct or inlet with multiple contour lines representing temperature distribution. The vertical axis is labeled \"Inlet-lip temperature, °F\" with tick marks at 0, 250, and 500. A legend indicates \"Hot-gas bleedback (percent)\" with values 3.73, 5.75, and 9.10 corresponding to different contour lines.]\n\nFigure 10. - Variation of inlet-lip temperature distribution with hot-gas bleedback. Plenum-chamber-gas temperature, $1000^\\circ$ F; tunnel-air velocity, 205 feet per second; tunnel total temperature, $0^\\circ$ F; angle of attack, $0^\\circ$.\n\nNACA\n\n23", "timestamp": "2026-07-22T07:00:43.406443+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 68, "total_pages": 149, "image_filename": "19930083192_p68.jpg", "text": "64\nNACA TN 1976\n\n# APPENDIX A\n\n## COOPERATING AIRLINES AND AGENCIES\n\nThe following airlines and agencies have cooperated extensively with the NACA in obtaining much of the data used in the preparation of this paper:\n\n*   American Airlines, Inc.\n*   Eastern Air Lines, Inc.\n*   Pan American World Airways System\n*   Trans World Airline, Inc.\n*   United Air Lines, Inc.\n*   Department of Commerce\n    *   U. S. Weather Bureau\n    *   Civil Aeronautics Administration\n*   U. S. Air Force\n    *   Air Materiel Command\n    *   Air Weather Service\n*   U. S. Navy, Bureau of Aeronautics, Structures Unit", "timestamp": "2026-07-22T07:00:45.154909+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 20, "total_pages": 132, "image_filename": "19930085990_p20.jpg", "text": "18\nCONFIDENTIAL\nNACA RM A9I01\n\nWing-Fuselage Combination with Bracket for Mounting Tail Above Fuselage\n\n| Results presented | Flap deflection | Mach number | Reynolds number | Figure number |\n| :--- | :--- | :--- | :--- | :--- |\n| $C_L$ vs $\\alpha$ | $0^\\circ$ | 0.20 to 0.95 | $2 \\times 10^6$ | 16 |\n| $C_L$ vs $C_D$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 17 |\n| $C_L$ vs $C_m$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 18 |\n| $\\alpha$, $C_D$ & $C_m$ vs $C_L$ | $\\delta_n=30^\\circ, \\delta_F=50^\\circ$ | 0.20 | $2 \\times 10^6$ to $10 \\times 10^6$ | 19 |\n\nWing, Fuselage, and Horizontal Tail Above Extended Wing-Chord Plane\n\n| Results presented | Flap deflection | Stabilizer angle | Mach number | Reynolds number | Figure number |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| $C_L$ vs $\\alpha$ | $0^\\circ$ | $4^\\circ$ to $-6^\\circ$ | 0.20 to 0.95 | $2 \\times 10^6$ | 20(a) to 20(h) |\n| $C_L$ vs $C_D$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 21(a) to 21(h) |\n| $C_L$ vs $C_m$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 22(a) to 22(h) |\n| $^1C_L$ vs $C_m$ | $\\downarrow$ | $0^\\circ$ | 0.20, 0.90<br>0.92, 0.93 | $2 \\times 10^6$ | 23 |\n| $C_L$ vs $\\alpha$ | $\\delta_n=30^\\circ, \\delta_F=50^\\circ$ | $4^\\circ, 0^\\circ$, & $-8^\\circ$ | 0.20 | $2 \\times 10^6$ to $10 \\times 10^6$ | 24 |\n| $C_L$ vs $C_D$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 25 |\n| $C_L$ vs $C_m$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 26 |\n\n$^1$Shows the effect of modifying the rear of the fuselage.\n\nFLOW CONDITIONS IN THE REGION OF THE HORIZONTAL TAIL\n\nCharacteristics of Wing-Fuselage Wake\n\n| Results presented | Flap deflection | Mach number | Reynolds number | Figure number |\n| :--- | :--- | :--- | :--- | :--- |\n| Location of wake vs $\\alpha$ | $0^\\circ$ | 0.20 to 0.95 | $2 \\times 10^6$ | 27(a) to 27(g) |\n| $\\downarrow$ | $\\delta_n=30^\\circ, \\delta_F=60^\\circ$ | 0.20 | $2 \\times 10^6$ to $10 \\times 10^6$ | 28 |\n| Pressure loss in wake vs distance from wing-chord plane | $0^\\circ$ | 0.85 | $2 \\times 10^6$ | 29 |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:00:49.024827+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 16, "total_pages": 40, "image_filename": "19930085997_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T07:00:50.419902+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 13, "total_pages": 32, "image_filename": "19930085992_p13.jpg", "text": "NACA RM L9E17\n11\n\nREFERENCES\n\n1. Runyan, Harry L., and Sewall, John L.: Experimental Investigation\nof the Effects of Concentrated Weights on Flutter Characteristics\nof a Straight Cantilever Wing. NACA TN 1594, 1948.\n\n2. Runyan, Harry L., and Watkins, Charles E.: Flutter of a Uniform\nWing with an Arbitrarily Placed Mass According to a Differential-\nEquation Analysis and a Comparison with Experiment.\nNACA TN 1848, 1949.\n\n3. Woolston, Donald S., and Runyan, Harry L.: Appraisal of Method of\nFlutter Analysis Based on Chosen Modes by Comparison with\nExperiment for Cases of Large Mass Coupling.\nNACA TN 1902, 1949.", "timestamp": "2026-07-22T07:00:52.296472+00:00"}
{"citation_id": "19930086020", "source_url": "https://ntrs.nasa.gov/api/citations/19930086020/downloads/19930086020.pdf", "page_number": 5, "total_pages": 22, "image_filename": "19930086020_p5.jpg", "text": "NACA RM A9J06 CONFIDENTIAL 3\n\n$y_c$ distance between camber line and chord line, feet\n\n$\\alpha$ angle of attack, degrees\n\n$\\alpha_t$ angle of twist, positive for washin, degrees\n\n$\\mu$ air viscosity, slugs per foot-second\n\n$\\rho$ mass density of air, slugs per cubic foot\n\nMODELS\n\nDimensions of the models used in this investigation are presented in figures 1 and 2. The two wings were made of steel and had identical plan forms: an aspect ratio of 3.5, a taper ratio of 0.25, and $63^\\circ$ of leading-edge sweepback. The untwisted wing was composed of NACA 64A006 airfoil sections in the streamwise direction. The cambered and twisted wing had the NACA 64A005 thickness distribution in combination with a = 1 mean camber lines. Distribution of wing twist and spanwise camber variation are presented in figure 2.\n\nIn addition to the wing-alone configuration, the untwisted symmetrical wing was tested alternately with a chordwise fence fitted near the wing root parallel to the stream direction, and with a half-fuselage of circular cross section having a fineness ratio of 6-1/4. These modifications are illustrated in figure 3.\n\nMETHODS AND EQUIPMENT\n\nThe majority of the data was obtained by placing the semispan models in a region of accelerated air flow over a special built-up test station on an airplane wing. The model was mounted on a three-component recording balance which was rotated to vary the angle of attack. A general view of the test station with the model installed is shown in figure 4. For certain of the tests the balance was installed in the side wall of the Ames 1- by 3-1/2-foot high-speed wind tunnel as illustrated in figure 5.\n\nA detailed description of the wing-flow test station and the force-measuring equipment used in this investigation is presented in reference 6, including discussions of the horizontal and vertical Mach number gradients, boundary-layer characteristics, and the three-component balance. The ratio of test-station boundary-layer-displacement thickness to model span for the wing-flow tests was 0.0075, nearly the same as\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:00:55.389844+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 27, "total_pages": 55, "image_filename": "19930085966_p27.jpg", "text": "26\nNACA RM L9B17\n\nCONFIDENTIAL\nSym. about $\\phi$\n\n6\n$1\\frac{7}{32}$\n3\n18 R\nRef. line $\\perp$ inlet & outlet faces\n30\n$4R$\n$12\\frac{3}{8}R$\n$1\\frac{7}{32}$\n+\nNACA\n\nAir flow\nSym. about $\\phi$\n\nFigure 4.- Diagrammatic sketch of inlet diffuser.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:00:55.588728+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 21, "total_pages": 46, "image_filename": "19930085983_p21.jpg", "text": "```markdown\nNACA RM A9I27\n\nCONFIDENTIAL\n\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n\nLift coefficient, $C_L$\n\n$\\delta_u$, deg\no 0\n□ -5\n◇ -10\n▽ -20\n△ -25\n\n-8 -4 0 4 8 12 16\nAngle of attack, $\\alpha$, deg\n\nfor $\\delta_u = 0^\\circ$\n\n.16 .12 .08 .04 0 -.04 -.08\nPitching-moment coefficient, $C_m$\n\nCONFIDENTIAL\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 5.- The effect of elevon deflection on the aerodynamic characteristics of the wing-fuselage combination and on the elevon hinge-moment coefficients at a Mach number of 0.60.\n\n19\n```", "timestamp": "2026-07-22T07:00:59.379119+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 95, "total_pages": 98, "image_filename": "19930086073_p95.jpg", "text": "NACA RM A9H04\n93\n\nIncrement of rolling-moment coefficient, $\\Delta C_l$\n0.02\n0\n-0.02\n0 .2 .4 .6 .8 1.0 1.2\n\nIncrement of yawing-moment coefficient, $\\Delta C_n$\n0.06\n0.04\n0.02\n0\n-0.02\n0 .2 .4 .6 .8 1.0 1.2\n$\\beta, 15.9^\\circ$\n$12.0^\\circ$\n$6.0^\\circ$\n$0.0^\\circ$\n\nIncrement of side-force coefficient, $\\Delta C_Y$\n0\n-.1\n0 .2 .4 .6 .8 1.0 1.2\nLift coefficient, $C_L$\n[Figure: NACA logo]\n\nFigure 25.— Increments of rolling-moment, yawing-moment, and side-force coefficients due to the vertical tail.", "timestamp": "2026-07-22T07:01:10.551008+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 14, "total_pages": 24, "image_filename": "19930085991_p14.jpg", "text": "```markdown\n12\n\nTABLE I.- RÉSUMÉ OF TESTS IN THE LANGLEY 20-FOOT FREE-SPINNING\nTUNNEL OF MODELS SIMULATING AIRPLANE JETTISONABLE\nNOSE SECTIONS WITHOUT STABILIZING FINS\n\n<!-- Table (83, 268, 880, 868) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|c|p{6cm}|}\n\\hline\n\\multirow{2}{*}{Model} & \\multirow{2}{*}{n} & \\multirow{2}{*}{Sketch} & \\multicolumn{5}{c|}{Mass characteristics} & \\multirow{2}{*}{Behavior of model} \\\\\n\\cline{4-8}\n& & & Weight (lb) & Center-of-gravity location (percent L) & $k_x/L$ & $k_y/L$ & $k_z/L$ & \\\\\n\\hline\n1 & 1.90 & & 0.704 & 69.2 & 0.141 & 0.256 & 0.256 & Tumbled end over end about lateral or normal axis with axis in an approximately horizontal attitude \\\\\n\\hline\n2 & 1.90 & & 0.683 & 57.0 & 0.179 & 0.321 & 0.321 & Tumbled end over end about lateral or normal axis with axis in an approximately horizontal attitude \\\\\n\\hline\n3 & 2.79 & & 0.460 & 67.0 & 0.141 & 0.316 & 0.305 & Rolled about longitudinal axis with axis in an approximately horizontal attitude or Tumbled about normal axis with axis in an approximately horizontal attitude \\\\\n\\hline\n4 & 3.80 & & 0.300 & 55.0 & ----- & ----- & ----- & Rotated or oscillated about various model axes in inconsistent manner \\\\\n\\hline\n5 & 5.85 & & 0.354 & 71.9 & 0.066 & 0.334 & 0.334 & Rolled about longitudinal axis with nose approximately 35° up from horizontal; at same time, rotated about wind axis \\\\\n\\hline\n\\end{tabular}\n\nNACA RM L9188\n```", "timestamp": "2026-07-22T07:01:11.514159+00:00"}
{"citation_id": "19930086083", "source_url": "https://ntrs.nasa.gov/api/citations/19930086083/downloads/19930086083.pdf", "page_number": 3, "total_pages": 48, "image_filename": "19930086083_p3.jpg", "text": "NACA RM L9F10\nRESTRICTED\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nRESEARCH MEMORANDUM\nPRELIMINARY AERODYNAMIC INVESTIGATION OF THE EFFECT\nOF CAMBER ON A 60° DELTA WING WITH\nROUND AND BEVELED LEADING EDGES\nBy John M. Riebe and Joseph E. Fikes\n\nSUMMARY\n\nAn exploratory investigation to determine the aerodynamic effects\nof camber on a 60° apex delta-wing model has been conducted in the\nLangley 300 MPH 7- by 10-foot tunnel. Camber variation was accom-\nplished through the deflection of full-span round and 25° beveled\nleading-edge flaps on a flat-sided triangular plan-form wing.\n\nThe maximum lift-drag ratio for the delta wing with no flap\ndeflection was about 8.2 for both leading-edge shapes at a Reynolds\nnumber of $3 \\times 10^6$. An increase up to 28 percent in lift-drag ratio\noccurred in the 0.2 to 0.3 lift-coefficient range for 20° flap\ndeflections. Flap deflections resulted in small wing trim changes but\ndid not greatly affect the longitudinal stability through the stall.\n\nDeflecting the leading-edge flap to 30° reduced the maximum\neffective dihedral to about half the flap-neutral value and resulted in\nnegative effective dihedral below lift coefficients of 0.3.\n\nINTRODUCTION\n\nIt has been indicated from recent research on wing plan forms\nsuitable for moderate supersonic flight that the low-speed longitudinal\nstability problems of the triangular plan-form wing appear to be less\nsevere than for the conventional sweptback wing; because of its high\ntaper and low aspect ratio, the triangular wing also has definite\nstructural advantages over the conventional swept wings.\n\nRESTRICTED", "timestamp": "2026-07-22T07:01:14.424553+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 38, "total_pages": 92, "image_filename": "19930085951_p38.jpg", "text": "36\n\n[CONFIDENTIAL]\n\nEfficiency, $\\eta$\n\n1.0\n.9\n.8\n.7\n.6\n.5\n.4\n.3\n.2\n.1\n0\n\n$\\beta_{0.75R}$ 20° 25° 30° 35° 40° 45°\n\nHelical-tip Mach number\n\nAir-stream Mach number\n\nMach number, M\n\n14\n12\n10\n8\n6\n4\n2\n0\n\nAdvance ratio, J\n\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0 3.2 3.4 3.6 3.8\n\n[NACA logo]\n\n(c) Efficiency.\n\nFigure 11.— Concluded. Rotational speed, 1600 rpm.\n\n[CONFIDENTIAL]\n\nNACA RM L9D29", "timestamp": "2026-07-22T07:01:19.668339+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 106, "total_pages": 114, "image_filename": "19930086061_p106.jpg", "text": "```markdown\n102\n\nWing leading edge\nQuarter chord\n$\\alpha$, deg $C_L$\n$\\circ$ 4.1 0.13\n$\\square$ 8.1 0.28\n$\\diamond$ 14.1 0.50\nLocal center of pressure, percent root chord\n$\\frac{y}{b/2}$, percent\n(a) Angles of attack: 4.1°, 8.1°, 14.1°.\n\nWing leading edge\nQuarter chord\n$\\alpha$, deg $C_L$\n$\\triangle$ 24.1 0.83\n$\\nabla$ 34.1 1.12\n$\\square$ 39.1 1.17\nLocal center of pressure, percent root chord\n$\\frac{y}{b/2}$, percent\nNACA\n(b) Angles of attack: 24.1°, 34.1°, 39.1°.\n\nFigure 51.- Effect of angle of attack on the local center-of-pressure location of wing 3; $\\psi = 0^\\circ$.\n\nNACA RM L9J07\n```", "timestamp": "2026-07-22T07:01:23.958216+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 25, "total_pages": 25, "image_filename": "19930085979_p25.jpg", "text": "```markdown\n24\nNACA RM E9E12\n\nPlenum-chamber-gas temperature, °F\n\n| Curve | A | B |\n| :--- | :--- | :--- |\n| Model-air total temperature (°F) | 45.9 | 43.1 |\n| Tunnel-air velocity (ft/sec) | 435 | 200 |\n| Liquid-water content (grams/cu meter) | 1.4 | 0.7 |\n\n| Tunnel-air velocity (ft/sec) | |\n| :--- | :--- |\n| O | 200 |\n| □ | 275 |\n| ◇ | 355 |\n| △ | 435 |\n\n[Figure: Graph showing two curves, A and B, plotting Plenum-chamber-gas temperature against Bleedback percent. Curve A is to the right of Curve B. Data points marked with symbols O, □, ◇, and △ are plotted along the curves.]\n\nNACA\n\nBleedback, percent\n\nFigure 11. - Bleedback required for ice prevention as function of plenum-chamber-gas temperature for tunnel total temperature of 0° F.\n\nNACA-Langley - 8-4-49 - 425\n```", "timestamp": "2026-07-22T07:01:24.177929+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 69, "total_pages": 149, "image_filename": "19930083192_p69.jpg", "text": "NACA TN 1976\n\nAPPENDIX B\n\nSYMBOLS\n\n| Symbol | Definition |\n|--------|------------|\n| A      | aspect ratio $(b^2/S)$ |\n| B      | arbitrary constant |\n| b      | wing span, feet |\n| C      | arbitrary constant |\n| $\\bar{c}$ | mean geometric chord of wing, feet |\n| c      | mean aerodynamic chord, feet |\n| $C_L$  | lift coefficient |\n| $C_{Lg}$ | variation of lift coefficient when penetrating a sharp-edge gust, expressed as a fraction of final lift (Küssner's function, reference 2) |\n| $C_{La}$ | variation of lift coefficient following a sudden change in angle of attack, expressed as a function of final lift (Wagner's function, reference 2) |\n| $C_{mcg}$ | pitching-moment coefficient about center of gravity |\n| D      | differential operator $\\left(\\frac{d}{dt}\\right)$ |\n| $dC_{mcg}/dC_L$ | static margin |\n| f      | frequency of occurrence |\n| $f_w$  | wing frequency, cycles per second |\n| g      | acceleration due to gravity, feet per second per second |", "timestamp": "2026-07-22T07:01:24.675557+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 26, "total_pages": 30, "image_filename": "19930085970_p26.jpg", "text": "```markdown\n24\n\nCONFIDENTIAL\n\n| Configuration | Reynolds number | Source |\n| :--- | :--- | :--- |\n| — Wing and body | $.35-.52 \\times 10^6$ | 1-by $3\\frac{1}{2}$-ft wind tunnel |\n| -- Wing alone | $2.35 \\times 10^6$ | Reference 3 |\n| ---- Wing and body | $.35-.52 \\times 10^6$ | Calculated |\n| o Wing and body | $.69 \\times 10^6$ | Reference 2 |\n\nMinimum drag coefficient, $C_{D_{min}}$\n\n[Graph plotting Minimum drag coefficient vs Mach number. The graph contains lines corresponding to the legend above. Annotations on the graph indicate \"fully turbulent boundary layer\" and \"fully laminar boundary layer\". The NACA logo is present in the bottom right corner of the plot area.]\n\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6\nMach number, M\n\nFigure 8.- Effect of Mach number on the minimum drag coefficient.\n\nCONFIDENTIAL\n\nNACA RM A59E09\n```", "timestamp": "2026-07-22T07:01:24.713966+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 17, "total_pages": 40, "image_filename": "19930085997_p17.jpg", "text": "NACA RM A9I29 CONFIDENTIAL 15\n\n[Figure: Side-scoop model employing boundary-layer suction. The image shows a pointed aerodynamic model with two small rectangular scoops on its surface, positioned next to a ruler marked in inches (1, 2, 3). A label “NACA A-12880-1” is visible near the base of the model.]\n\nFigure 2.- Side-scoop model employing boundary-layer suction.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:01:25.618137+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 32, "total_pages": 34, "image_filename": "19930085913_p32.jpg", "text": "```markdown\nNACA RM L9F24\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T07:01:26.655896+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 11, "total_pages": 20, "image_filename": "19930085999_p11.jpg", "text": "NACA RM E9I07\n\nvibrations were also observed. The sources of excitation were of frequencies related to the number of nozzle blades and combustion chambers and to multiples of the first-order turbine speed.\n\nComparison was made between the results of this investigation and those obtained during similar research on tightly mounted turbine blades. The results indicated that loose blade mounting had little effect on frequency of vibration and vibratory-stress levels. The loosely mounted blades were observed to be more susceptible to vibration in the first bending mode.\n\nThe results presented were obtained from vibration in turbine blades that were subjected to low excitation forces; in the presence of more severe excitation forces, however, the increased damping attributable to loose mounting of a turbine blade may be a factor of importance in reducing the level of vibratory stress.\n\nLewis Flight Propulsion Laboratory,  \nNational Advisory Committee for Aeronautics,  \nCleveland, Ohio.\n\nREFERENCES\n\n1. Voysey, R. G.: Some Vibration Problems in Gas Turbine Engines. Development of the British Gas Turbine Jet Unit. Inst. Mech. Eng. (London). (Reprinted for distribution in the United States by A.S.M.E., Jan. 1947, pp. 483-495.)\n\n2. Farmer, J. Elmo, Deutsch, George C., and Sikora, Paul F.: Cyclic Engine Test of Cast Vitallium Turbine Buckets - II. NACA RM E7J24, 1948.\n\n3. Shannon, J. F.: Vibration Problems in Gas Turbines, Centrifugal and Axial Flow Compressors. R. & M. No. 2226, British A.R.C., March 1945.\n\n4. Morgan, W. C., Kemp, R. H., and Manson, S. S.: Vibration of Turbine Blades in a TurboJet Engine during Operation. NACA RM E7L18, 1948.\n\n5. Manson, S. S., Meyer, A. J., Jr., Calvert, H. F., and Hanson, M. P.: Factors Affecting Vibration of Axial-Flow Compressor Blades. Proc. Soc. Exp. Stress Analysis, vol. 7, no. 2, 1948.", "timestamp": "2026-07-22T07:01:31.031123+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 22, "total_pages": 46, "image_filename": "19930085983_p22.jpg", "text": "20\nCONFIDENTIAL\nNACA RM A9I27\n\n<!-- Image (219, 109, 735, 497) -->\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n$\\delta_u, deg$\n$\\circ$ 0\n$\\square$ -5\n$\\diamond$ -10\n$\\nabla$ -20\n$\\blacktriangleright$ -25\n\n.20 for $\\delta_u=0^\\circ$\n\n<!-- Image (219, 511, 735, 880) -->\n\nHinge-moment coefficient, $C_h$\nAngle of attack, $\\alpha, deg$\n\nNACA\n\n(b) $C_L$ vs $C_D$, $C_h$ vs $\\alpha$.\nFigure 5.- Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:01:33.852723+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 46, "total_pages": 52, "image_filename": "19930085911_p46.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 45\n\n1152\n\n| | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | 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| |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | | |\n| | | | | | | | | | |", "timestamp": "2026-07-22T07:01:34.579174+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 96, "total_pages": 98, "image_filename": "19930086073_p96.jpg", "text": "94\nNACA RM A9H04\n\n<!-- Image (129, 109, 781, 856) -->\n\nFigure 26.— Influence of separation vortices on the angle of attack of the vertical tail surface.", "timestamp": "2026-07-22T07:01:36.410596+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 15, "total_pages": 24, "image_filename": "19930085991_p15.jpg", "text": "NACA RM L91C8\n\nL\nSp\nSF\nLT\nC.G. of nose\nCentroid of SF\nSp, side area of model\nSF, smallest projected area of fins\nPlane of smallest projected fin area\nNACA\nFigure 1.- Illustration of method of computing fin-stabilization factor $\\frac{S_F L_T}{S_P L}$.\n\n13", "timestamp": "2026-07-22T07:01:39.223099+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 21, "total_pages": 132, "image_filename": "19930085990_p21.jpg", "text": "NACA RM A9I01 CONFIDENTIAL 19\n\nDynamic Pressure Ratio, Mach Number, and Effective Angle of Downwash at the Tail\n\n| Results presented | Flap deflection | Mach number | Reynolds number | Figure number |\n|---|---|---|---|---|\n| $q_t/q$ vs $\\alpha$ | $0^\\circ$ | 0.20 to 0.95 | $2 \\times 10^6$ | 30(a) & 30(b) |\n| | $\\delta_n=30^\\circ, \\delta_F=60^\\circ$ | 0.20 | $2 \\times 10^6$ to $10 \\times 10^6$ | 31 |\n| $M_t$ vs $\\alpha$ | $0^\\circ$ | 0.20 to 0.95 | $2 \\times 10^6$ | 32(a) & 32(b) |\n| $\\epsilon$ vs $\\alpha$ | $0^\\circ$ | 0.20 to 0.95 | $2 \\times 10^6$ | 33(a) & 33(b) |\n| $\\epsilon$ vs $\\alpha$ | $\\delta_n=30^\\circ, \\delta_F=60^\\circ$ | 0.20 | $2 \\times 10^6$ to $10 \\times 10^6$ | 34(a) & 34(b) |\n| $\\eta(q_t/q)$ vs M | $0^\\circ$ | 0.20 to 0.95 | $2 \\times 10^6$ | 35(a) & 35(b) |\n| $\\eta(q_t/q)$ vs $\\alpha$ | $\\delta_n=30^\\circ, \\delta_F=60^\\circ$ | 0.20 | $2 \\times 10^6$ to $10 \\times 10^6$ | 36(a) & 36(b) |\n\nSUMMARY CURVES\n\nThe Effects of Compressibility on the Characteristics of the Model\n[Flap deflection, $0^\\circ$; Reynolds number, $2 \\times 10^6$]\n\nLift and drag:\n\n| Results presented | Lift coefficient | Stabilizer angle | Mach number | Figure number |\n|---|---|---|---|---|\n| $\\alpha$ vs M | 0 to 0.6 | $0^\\circ$ | 0.20 to 0.95 | 37(a) & 37(b) |\n| $C_D$ vs M | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 38(a) & 38(b) |\n\nLongitudinal stability and control characteristics:\n\n| Results presented | Lift coefficient | Stabilizer angle | Mach number | Figure number |\n|---|---|---|---|---|\n| $C_m$ vs M | 0 to 0.6 | $0^\\circ$ | 0.20 to 0.95 | 39(a) & 39(b) |\n| $\\epsilon$ vs M | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 40(a) & 40(b) |\n| $\\partial\\epsilon/\\partial\\alpha$ vs M | | $0^\\circ$ to $-4^\\circ$ | | 41 |\n| $C_L$ for $C_m=0$ vs M | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 42(a) & 42(b) |\n| $^2C_L$ vs M | --- | --- | 0.50 to 0.95 | 43 |\n| $i_t$ for $C_m=0$ vs M | --- | --- | 0.20 to 0.95 | 44 |\n| $C_{L_e}$ vs $\\delta_e$ | --- | --- | 0.20 to 0.94 | 45 |\n| $\\delta_e$ for $C_m=0$ vs M | --- | --- | 0.20 to 0.95 | 46 |\n\nLongitudinal Characteristics with the Flaps Deflected\n\n| Results presented | Flap deflection | Mach number | Reynolds number | Figure number |\n|---|---|---|---|---|\n| $i_t$ for $C_m=0$ & $C_D$ vs $C_L$ | $\\delta_n=30^\\circ, \\delta_F=50^\\circ$ | 0.20 | $10 \\times 10^6$ | 47 |\n| L/D vs $C_L$ for $C_m=0$ | $\\downarrow$ | $\\downarrow$ | $\\downarrow$ | 48 |\n| $^3$Sinking speed vs Gliding speed | | | | 49 |\n\n$^2$Lift requirements of hypothetical airplane with a wing loading of 100 pounds per square foot in flight at an altitude of 10,000 feet.\n\n$^3$Sinking speed for hypothetical airplane with a wing loading of 100 pounds per square foot in flight at sea level. Power off.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:01:39.458950+00:00"}
{"citation_id": "19930086020", "source_url": "https://ntrs.nasa.gov/api/citations/19930086020/downloads/19930086020.pdf", "page_number": 6, "total_pages": 22, "image_filename": "19930086020_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM A9J06\n\nthe value of 0.0083 measured in the Ames 12-foot pressure wind tunnel\nduring the tests reported in reference 1. The ratio for the wing-flow\nmodel mounted on the side of the Ames 1- by 3-1/2-foot wind tunnel was\n0.0330.\n\nTESTS\n\nThe wing-flow data were recorded in the form of time histories of\nan oscillation of the model from -5° to +8° angle of attack at various\nconstant Mach numbers from 0.52 to 1.11. The corresponding Reynolds\nnumbers are presented in figure 6. Tests were conducted on the following\nconfigurations:\n\n1. Symmetrical untwisted wing\n2. Cambered and twisted wing\n3. Symmetrical untwisted wing plus fuselage\n4. Symmetrical untwisted wing plus chordwise boundary-layer\n   fence\n\nIn addition, the wing-flow balance was mounted on the wall of the\nAmes 1- by 3-1/2-foot high-speed wind tunnel so that the top of the\nbalance was flush with the inside of the tunnel wall. The Mach number\nrange in these tests was 0.75 to 0.92, with an approximate range of\nReynolds number of 0.69 million to 0.78 million. Tests were conducted\non the symmetrical untwisted wing at constant Mach numbers, both by\noscillating the model over the angle-of-attack range and by recording\nat various fixed angles of attack.\n\nPRECISION\n\nThe precision of the physical measurements made during these tests\nhas been evaluated as described in reference 6. The following table\nshows representative values of the test data and the physical uncertainty\nin each, at the lowest and highest Mach numbers at a lift coefficient\nof 0.30:\n\n| Quantity | M = 0.52 | M = 1.11 |\n| :--- | :--- | :--- |\n| Mach number M | 0.52 $\\pm$0.01 | 1.11 $\\pm$0.02 |\n| Angle of attack $\\alpha$, degrees | 7.8 $\\pm$0.4 | 6.8 $\\pm$0.4 |\n| Lift coefficient $C_L$ | 0.3 $\\pm$0.01 | 0.3 $\\pm$0.006 |\n| Pitching-moment coefficient $C_{m_{0.25\\bar{c}}}$ | 0.003 $\\pm$0.0009 | 0.002 $\\pm$0.0001 |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:01:39.677739+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 59, "total_pages": 92, "image_filename": "19930085870_p59.jpg", "text": "60\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O CL\n                {□ Cm\nWedge L.E. {△ CL\n           {◇ Cm\n.16\n.08\nCL\n0\n-.08\n-.16\n-.24\n.01\nCm\n0\n-.01\n\n.06\nElliptical L.E. {O CD\n                {□ L/D\nWedge L.E. {△ CD\n           {◇ L/D\n.04\nCD\n.02\n0.8\n-6\n-4\n-2\n0\n2\n4\n6\n8\nα, deg\n6\n4\nL/D\n2\n0\n[NACA logo]\n\n(h) Wing 8. w=1.540; R = 620,000.\nFigure 7. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:01:39.917277+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 40, "total_pages": 47, "image_filename": "19930085919_p40.jpg", "text": "NACA RM No. A9C21\nCONFIDENTIAL\n\nLift coefficient, $C_L$\nAngle of attack, $\\alpha$, deg\nPitching-moment coefficient, $C_m$\n\n| $\\delta_f$, deg | $\\delta_{dn}$, deg | Drooped-nose flap | $\\delta_f$, deg | $\\delta_{dn}$, deg | Drooped-nose flap |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| $\\square$ | 0 | 40 | $\\diamond$ | 0 | 40 |\n| $\\square$ | 40 | 40 | $\\diamond$ | 40 | 40 |\n| $\\square$ | 0 | 0 | $\\diamond$ | 0 | 0 |\n| | | Constant-chord aileron | | | Constant-chord aileron |\n| | | | | | -0.25c split flap |\n| | | | | | $\\delta_f = 45^\\circ$ |\n\n0.06c\n0.06c\n1.00c\n\nNACA\n\nFigure 14.- Effect of the drooped-nose flap of full span on the lift and pitching-moment characteristics of the model with short fuselage. $R, 4.2 \\times 10^6$.\n\n39", "timestamp": "2026-07-22T07:01:41.323229+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 39, "total_pages": 92, "image_filename": "19930085951_p39.jpg", "text": "NACA RM L9D29\n37\n\n[Figure: Graph plotting Thrust coefficient, $C_T$ (y-axis, 0 to .24) against Advance ratio, J (x-axis, 0 to 1.8). Four curves are shown for $\\beta_{0.75R} = 20^\\circ, 25^\\circ, 30^\\circ, 35^\\circ$. A \"NACA\" logo is visible on the right side of the plot area.]\n\n(a) Thrust coefficient.\n\nFigure 12.- Characteristics of NACA 10-(3)(062)-045A propeller.\nRotational speed, 2000 rpm.", "timestamp": "2026-07-22T07:01:45.482386+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 107, "total_pages": 114, "image_filename": "19930086061_p107.jpg", "text": "NACA RM L9J07\n\nSpanwise center-of-pressure location, percent semispan\n48\n40\n32\n24\n16\n8\n0\n\nPlan form Wing A\n1 3.46\n2 2.31\n3 1.73\n\n0 4 8 12 16 20 24 28 32 36 40 44 48 52\n$\\alpha$, deg\n\nFigure 52.- Variation of wing spanwise center-of-pressure location with angle of attack for the wing investigated; $\\psi = 0^\\circ$.\n\nNACA\n103", "timestamp": "2026-07-22T07:01:49.374786+00:00"}
{"citation_id": "19930086083", "source_url": "https://ntrs.nasa.gov/api/citations/19930086083/downloads/19930086083.pdf", "page_number": 4, "total_pages": 48, "image_filename": "19930086083_p4.jpg", "text": "2\nNACA RM L9F10\n\nTheoretical calculations and experimental studies have shown that such wing plan forms will develop lift-drag ratios at supersonic Mach numbers which are sufficiently high for flight, being, however, generally lower than those of other plan forms. Low-speed research has also indicated poor landing characteristics of delta plan forms because of the relatively low lift-drag ratios, particularly in the high-lift condition.\n\nThe results of a theoretical study of triangular wings (reference 1) and a pressure-distribution investigation (reference 2) show that high loadings occur along the leading edge of triangular wings. These high loadings and their associated adverse pressure gradients result in separation over the leading-edge portion of the wing and develop vortices that flow back over the wing (references 1 to 5) even at relatively low lifts.\n\nThe present investigation made in the Langley 300 MPH 7- by 10-foot tunnel is a preliminary study aimed at increasing the low-speed lift-drag ratio by incorporating camber into a $60^\\circ$ delta wing. Camber was simulated for this investigation by deflecting full-span leading-edge flaps in an effort to reduce the high peak pressure along the leading edge and thus retard the separation effects. The configuration tested was chosen as the first attempt at solving this problem because it approaches a conical shape, which was believed to be efficient in unloading the leading edge, and because of its structural simplicity compared to other leading-edge flaps considered.\n\nCOEFFICIENTS AND SYMBOLS\n\nThe results of the tests are presented as standard NACA coefficients of forces and moments about the stability axes. Pitching-, yawing-, and rolling-moment coefficients are given about the wing 25-percent-mean-aerodynamic-chord point as shown in figure 1. The positive directions of forces and moments are shown in figure 2.\n\nThe coefficients and symbols are defined as follows:\n\n| | |\n| :--- | :--- |\n| $C_L$ | lift coefficient (L/qS) |\n| $C_D$ | drag coefficient (D/qS) |\n| $C_Y$ | lateral-force coefficient (Y/qS) |\n| $C_l$ | rolling-moment coefficient (L'/qbS) |", "timestamp": "2026-07-22T07:01:49.850193+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 18, "total_pages": 40, "image_filename": "19930085997_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T07:01:50.932028+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 70, "total_pages": 149, "image_filename": "19930083192_p70.jpg", "text": "```markdown\n66\nNACA TN 1976\n\nH\ngust-gradient distance (distance in which the\nvertical velocity of a gust rises linearly\nfrom zero to its maximum value), feet or chords\n\nK\ngust alleviation factor\n\nL\nlift, pounds\n\nl\ndistance from trailing edge of wing to leading\nedge of tail, chords\n\nM\nMach number\n\nn\nnormal acceleration, g units\n\n$\\Delta n$\nacceleration increment due to a gust, g units\n(n - 1 for most cases)\n\n$\\Delta n_g$\nacceleration increment, g units\n$$ \\left( \\text{computed according to formula } \\frac{\\rho S U V \\frac{dC_L}{d\\alpha}}{2W} \\right) $$\n\n$\\Delta n_r$\nacceleration increment on rigid airplane, g units\n\nP\nprobability\n\nq\ndynamic pressure, pounds per square foot $\\left( \\frac{1}{2} \\rho V^2 \\right)$\n\nS\nwing area, square feet\n\ns\ndistance airplane has penetrated gust at time t,\nchords\n\nt\ntime from start of gust penetration, seconds\n(cs/V)\n\nU\ngust velocity at peak of gust, feet per second\n\n$U_e$\neffective gust velocity, feet per second\n$$ \\left( \\text{computed from recorded acceleration increment} \\right. $$\n$$ \\left. \\text{as } \\frac{2 \\Delta n W/S}{\\rho_o V_e K \\frac{dC_L}{d\\alpha}} \\right) $$\n```", "timestamp": "2026-07-22T07:01:55.571768+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 27, "total_pages": 30, "image_filename": "19930085970_p27.jpg", "text": "NACA RM A52E09\nCONFIDENTIAL\n\n| Variable | Configuration | Reynolds number | Source |\n| :--- | :--- | :--- | :--- |\n| $\\frac{\\Delta C_D}{\\Delta C_L^2}$ | Wing and body | $.35-.52 \\times 10^6$ | 1-by $3\\frac{1}{2}$-ft wind tunnel |\n| $\\frac{1}{180} \\frac{d\\alpha}{dC_L}$ | Wing and body | $.35-.52 \\times 10^6$ | 1-by $3\\frac{1}{2}$-ft wind tunnel |\n| $\\frac{\\Delta C_D}{\\Delta C_L^2}$ | Wing alone | | Calculated |\n| $\\frac{\\Delta C_D}{\\Delta C_L^2}$ | Wing and body | $.69 \\times 10^6$ | Reference 2 |\n\nDrag-rise factor, $\\Delta C_D / \\Delta C_L^2$\n\n[Figure: Graph plotting Drag-rise factor against Mach number. The graph includes curves for \"zero leading-edge thrust\" and \"full leading-edge thrust\". A small NACA logo is visible in the bottom right corner of the plot area.]\n\nMach number, M\n\nFigure 9.- Effect of Mach number on the drag-rise factor.\n\nCONFIDENTIAL\n25", "timestamp": "2026-07-22T07:01:55.660766+00:00"}
{"citation_id": "19930086110", "source_url": "https://ntrs.nasa.gov/api/citations/19930086110/downloads/19930086110.pdf", "page_number": 1, "total_pages": 42, "image_filename": "19930086110_p1.jpg", "text": "NACA RM E9H15\n\nCONFIDENTIAL\n\nCopy 296\nRM E9H15\n\nNACA\n\nRESEARCH MEMORANDUM\n\nDOWNWASH IN VORTEX REGION BEHIND TRAPEZOIDAL-WING\nTIP AT MACH NUMBER 1.91\n\nBy J. L. Cummings, H. Mirels, and L. E. Baughman\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nDATE 9-13-54\nAUTHORITY J. W. CROWLEY\nCHANGE # 2456\nW.H.L.\n\nCLASSIFIED DOCUMENT\n\nThis document contains 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 so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officers and employees of the Federal Government who have a need-to-know, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nDATE 9-13-54\nAUTHORITY J. W. CROWLEY\nCHANGE # 2456\nW.H.L.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nNovember 10, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:01:55.844908+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 12, "total_pages": 20, "image_filename": "19930085999_p12.jpg", "text": "10\nNACA RM E9I07\n\n6. Kemp, R. H., Morgan, W. C., and Manson, S. S.: The Application\nof High-Temperature Strain Gages to the Measurement of Vibra-\ntory Stresses in Gas-Turbine Buckets. NACA TN 1174, 1947.", "timestamp": "2026-07-22T07:01:56.859623+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 47, "total_pages": 52, "image_filename": "19930085911_p47.jpg", "text": "46 CONFIDENTIAL NACA RM E9F22\n\n<!-- Image (108, 147, 877, 839) -->\n\n(d) Combustion-chamber-inlet variables.\nFigure 10. - Continued. Time history of flight data and performance of ram-jet unit 16-A-5.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:01:57.683238+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 97, "total_pages": 98, "image_filename": "19930086073_p97.jpg", "text": "NACA RM A9H04\n95\n\n<!-- Image (128, 109, 856, 833) -->\n\nFigure 27. — Increments of rolling-moment, yawing-moment, and side-force coefficients per degree of rudder deflection; $\\delta_r = 10^\\circ$.", "timestamp": "2026-07-22T07:01:59.246862+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 23, "total_pages": 46, "image_filename": "19930085983_p23.jpg", "text": "NACA RM A9I27\nCONFIDENTIAL\n\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n\nLift coefficient, $C_L$\n\n$\\delta_u$, deg\no 0\n$\\square$ -5\n$\\diamond$ -10\n$\\triangle$ -15\n$\\nabla$ -20\n$\\blacktriangledown$ -25\n\n-8 -4 0 4 8 12 16\nAngle of attack, $a$, deg\n\nfor $\\delta_u = 0^\\circ$\n\n.16 .12 .08 .04 0 -.04 -.08\nPitching-moment coefficient, $C_m$\n\nCONFIDENTIAL\nNACA\n21\n\n(a) $C_L$ vs $a$, $C_L$ vs $C_m$.\nFigure 6.- The effect of elevon deflection on the aerodynamic characteristics of the wing-fuselage combination and on the elevon hinge-moment coefficients at a Mach number of 0.80.", "timestamp": "2026-07-22T07:02:00.357618+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 41, "total_pages": 47, "image_filename": "19930085919_p41.jpg", "text": "CONFIDENTIAL\n40\nCONFIDENTIAL\nNACA RM No. A9C21\n\n<!-- Image (104, 197, 905, 714) -->\n\nFigure 15.- Effect of the extended-nose flap of 50-percent span on the lift and pitching-moment characteristics of the model with short fuselage. R, 4.2x10$^6$.", "timestamp": "2026-07-22T07:02:05.235484+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 60, "total_pages": 92, "image_filename": "19930085870_p60.jpg", "text": "```markdown\nNACA RM No. L9D07\n61\n\nCONFIDENTIAL\n\nElliptical L.E. { (O) $C_L$\n                  { (□) $C_m$\nWedge L.E. { (△) $C_L$\n             { (◇) $C_m$\n\n$C_L$\n.24\n.16\n.08\n0\n-.08\n-.16\n-.24\n\n$C_m$\n.01\n0\n-.01\n\nElliptical L.E. { (O) $C_D$\n                  { (□) $L/D$\nWedge L.E. { (△) $C_D$\n             { (◇) $L/D$\n\n$C_D$\n.06\n.04\n.02\n0\n\n$L/D$\n6\n4\n2\n0\n\n-8 -6 -4 -2 0 2 4 6 8\n$\\alpha$, deg\n\n[NACA logo]\n\n(i) Wing 9. $w = 1.705$; $R = 560,000$.\nFigure 7.- Continued.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T07:02:07.246661+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 40, "total_pages": 92, "image_filename": "19930085951_p40.jpg", "text": "38\nNACA RM L9D29\n\n[Figure: Graph of Power coefficient, $C_P$ vs. Advance ratio, $J$. The y-axis ranges from 0 to .24. The x-axis ranges from 0 to 1.8. Four curves are plotted, labeled $\\beta_{0.75R} = 20^\\circ$, $25^\\circ$, $30^\\circ$, and $35^\\circ$. A NACA logo is present on the graph.]\n\n(b) Power coefficient.\nFigure 12.— Continued. Rotational speed, 2000 rpm.", "timestamp": "2026-07-22T07:02:07.963430+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 22, "total_pages": 132, "image_filename": "19930085990_p22.jpg", "text": "20 CONFIDENTIAL NACA RM A9I01\n\nREFERENCES\n\n1. Johnson, Ben H., Jr.: Investigation of a Thin Wing of Aspect Ratio 4 in the Ames 12-Foot Pressure Wind Tunnel. I – Characteristics of a Plain Wing. NACA RM A3D07, 1948.\n\n2. Johnson, Ben H., Jr., and Bandettini, Angelo: Investigation of a Thin Wing of Aspect Ratio 4 in the Ames 12-Foot Pressure Wind Tunnel. II – The Effect of Constant-Chord Leading- and Trailing-Edge Flaps on the Low-Speed Characteristics of the Wing. NACA RM A3F15, 1948.\n\n3. Johnson, Ben H., and Demele, Fred A.: Investigation of a Thin Wing of Aspect Ratio 4 in the Ames 12-Foot Pressure Wind Tunnel. III – The Effectiveness of a Constant-Chord Aileron. NACA RM A3I17, 1948.\n\n4. Johnson, Ben H., Jr., and Reed, Verlin D.: Investigation of a Thin Wing of Aspect Ratio 4 in the Ames 12-Foot Pressure Wind Tunnel. IV – The Effect of a Constant-Chord Leading-Edge Flap at High Subsonic Speeds. NACA RM A3K19, 1949.\n\n5. Rathert, George A., Hanson, Carl M., and Rolls, L. Stewart: Investigation of a Thin Straight Wing of Aspect Ratio 4 by the NACA Wing-Flow Method. – Lift and Pitching-Moment Characteristics of the Wing Alone. NACA RM A3L20, 1949.\n\n6. Bandettini, Angelo, and Reed, Verlin D.: The Aerodynamic Characteristics Throughout the Subsonic Speed Range of a Thin, Sharp-Edged Horizontal Tail of Aspect Ratio 4 Equipped with a Constant-Chord Elevator. NACA RM A9E05, 1949.\n\n7. Sivells, James C., and Deters, Owen J.: Jet-Boundary and Plan-Form Corrections for Partial-Span Models with Reflection Plane, End Plate, or No End Plate in a Closed Circular Wind Tunnel. NACA Rep. 843, 1946.\n\n8. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnels, with Consideration of the Effect of Compressibility. NACA RM A7B28, 1947.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:02:10.414353+00:00"}
{"citation_id": "19930086020", "source_url": "https://ntrs.nasa.gov/api/citations/19930086020/downloads/19930086020.pdf", "page_number": 7, "total_pages": 22, "image_filename": "19930086020_p7.jpg", "text": "NACA RM A9J06 CONFIDENTIAL 5\n\nRESULTS AND DISCUSSION\n\nBasic Data\n\nThe typical variations of angle of attack and pitching-moment coefficient with lift coefficient are illustrated in figure 7 by the basic test data for the symmetrical untwisted wing. These same curves for all the test configurations were equally linear and indicated no obvious irregularities.\n\nComparison with Larger-Scale Tests\n\nThe characteristics of both the symmetrical wing alone and the cambered and twisted wing alone are summarized in figure 8, which shows the lift-curve slopes and the locations of the aerodynamic center as a function of Mach number. Also included in figure 8 are corresponding data up to 0.925 Mach number and at a Reynolds number of approximately 2 million from tests in the Ames 12-foot pressure wind tunnel (references 1 and 4). The comparison for the symmetrical wings is based upon tests using the same type of model and mounting; that is, semispan model on a flat reflection plate. In the case of the cambered and twisted wings the wind-tunnel model was full span and sting mounted;¹ whereas the wing-flow model again was semispan.\n\nThe comparison in figure 8 between wing-flow and wind-tunnel results for the symmetrical wing indicates fair agreement for the variation of lift-curve slope with Mach number up to the limit of the wind-tunnel tests. The pitching-moment-curve slopes, however, reveal a considerable discrepancy. The aerodynamic-center location as determined from the wing-flow tests would be about 18 percent of the mean aerodynamic chord forward of the position indicated by the wind-tunnel tests. The comparison for the cambered and twisted wing shows the wing-flow model had a lower lift-curve slope which decreased rather than increased with increasing Mach number. The pitching-moment characteristics show the same sizable differences, as in the case of the symmetrical wings.\n\nAdditional Tests\n\nThe noted discrepancies cast serious doubt on the validity of the wing-flow data on the test wings, particularly in regard to the pitching-moment characteristics.² Since quite satisfactory correlation between\n\n---\n\n¹The sting mount necessitated the addition of a fuselage; thus these results are for the wing-fuselage combination.\n\n²The effects of aeroelasticity were considered but found to be within the experimental scatter of the measurements.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:02:13.035729+00:00"}