PDF/ocr_dataset_1045.jsonl

{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 28, "total_pages": 51, "image_filename": "19930085962_p28.jpg", "text": "```markdown\nNACA RM A9E05\n\nCONFIDENTIAL\n\n12\n10\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n-1.0\n-12 -8 -4 0 4 8 12 .12 .08 .04 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28 -.32 -.36\nAngle of attack, $\\alpha$, deg Pitching-moment coefficient, $C_m$\n\n$\\delta_e$\n(deg)\n$\\circ$ 0\n$\\square$ 2\n$\\diamond$ 4\n$\\triangle$ 6\n$\\nabla$ 10\n$\\blacktriangledown$ 20\n$\\blacktriangle$ 30\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 10.—The effect of elevator deflection on the aerodynamic characteristics of the tail at a Mach number of 0.92.\n\nCONFIDENTIAL\n\nNACA\n\n27\n```", "timestamp": "2026-07-22T05:43:19.674408+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 1, "total_pages": 21, "image_filename": "19930091993_p1.jpg", "text": "AERO. & ASTRO. LIBRARY\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\ncopy #3\nREPORT 928\n\nANALYSIS OF PERFORMANCE OF JET ENGINE FROM\nCHARACTERISTICS OF COMPONENTS\nII—INTERACTION OF COMPONENTS AS DETERMINED\nFROM ENGINE OPERATION\n\nBy ARTHUR W. GOLDSTEIN, SUMNER ALPERT, WILLIAM BEEDE\nand KARL KOVACH\n\n[Figure: Seal of the National Advisory Committee for Aeronautics]\n\n1949\n\nFor sale by the Superintendent of Documents, U. S. Government Printing Office, Washington 25, D. C. Yearly subscription, $3; foreign, $4.50;\nsingle copy price varies according to size - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - 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- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -", "timestamp": "2026-07-22T05:43:19.880248+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 31, "total_pages": 36, "image_filename": "19930085912_p31.jpg", "text": "NACA RM No. E9C16\n\nRam-pressure recovery, $\\eta$\n\nRise in model-air total temperature, $^\\circ$F\n\n$\\eta = \\eta_{\\text{calc}} - \\left(1 - \\frac{460 + T_0}{460 + T_{\\text{av}}}\\right)$\n\n[Figure: Graph showing variation of ram-pressure recovery with rise in model-air total temperature, including data points and a calculated line labeled $\\eta_{\\text{calc}}$, with an equation annotation.]\n\nNACA\n\nFigure 10. - Variation of ram-pressure recovery with rise in model-air total temperature for plenum-chamber gas temperature of $1000^\\circ$ F. Angle of attack, $0^\\circ$; free-stream total temperature, $0^\\circ$ F.\n\n29", "timestamp": "2026-07-22T05:43:25.902732+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 57, "total_pages": 78, "image_filename": "19930082618_p57.jpg", "text": "```markdown\nNACA TN 1945\n\nSection lift coefficient, $c_l$\nMoment coefficient, $c_{m_{c/4}}$\nSection angle of attack, $\\alpha_0$, deg.\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\triangledown$ 6.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(b) Section lift and pitching-moment characteristics of the NACA 4412 airfoil section with a\n0.20c simulated split flap deflected 60°.\n\nFigure 12.— Continued.\n\n55\n```", "timestamp": "2026-07-22T05:43:26.811051+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 19, "total_pages": 46, "image_filename": "19930085972_p19.jpg", "text": "NACA RM L9B18\n17\n\n[Figure: A black and white photograph showing a front view of a model aircraft mounted on a single-support strut. The model has a pointed nose, swept wings, and external airfoil flaps. A label in the bottom right corner of the photo reads \"NACA L-52659.1\".]\n\n(a) Front view.\n\nFigure 5.- Variable-sweep model mounted on single-support strut with faired cutout and external airfoil flaps. $\\Lambda = 15^\\circ$.", "timestamp": "2026-07-22T05:43:29.275972+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 2, "total_pages": 34, "image_filename": "19930082472_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:43:29.337256+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 29, "total_pages": 42, "image_filename": "19930085914_p29.jpg", "text": "28\n\n20\nFlagged symbols denote\nrerun.\n16\nLift-drag ratio, L/D\n12\n8\n4\nM=.20\nM=.40\nM=.60\nM=.80\nM=.25\nM=.89\nM=.92\nM=.93\nNACA\n0\n0\n.2\n.4\n.6\n.8\n1.0\nfor M=.20\nLift coefficient, $C_L$\n\nFigure 9.- The variation of lift-drag ratio with lift coefficient of the wing-fuselage\ncombination for several Mach numbers at a Reynolds number of 2,000,000.\n\nNACA RM A9D25", "timestamp": "2026-07-22T05:43:30.028773+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 8, "total_pages": 28, "image_filename": "19930086092_p8.jpg", "text": "6\nCONFIDENTIAL\nNACA RM A9F14\n\n$$\n\\Delta\\alpha = 0.48 \\text{ C}_L\n$$\n\n$$\n\\Delta\\text{C}_D = 0.0084 \\text{ C}_L^2\n$$\n\nNo corrections have been applied for the drag and interference of the struts. With the exception of the effects on the drag results, these corrections are felt to be negligible. The effect on drag is of the order of $\\Delta\\text{C}_D = 0.008$ at zero lift, but is not known with sufficient accuracy to warrant application.\n\nRESULTS\n\nThe aerodynamic characteristics of the model with the tail off are shown in figure 4; the characteristics with the tail on and the rudder undeflected are shown in figure 5. At angles of attack from $0^\\circ$ to $12^\\circ$, the effectiveness of the vertical tail (fig. 5) was maintained to an angle of sideslip of $25^\\circ$ (the highest tested). At an angle of attack of $21^\\circ$, however, effectiveness was maintained only to an angle of sideslip of $7^\\circ$; beyond $7^\\circ$ the directional stability was irregular. This is the same attitude at which the directional stability of the model with tail off (fig. 4) became irregular.\n\nThe variations with angle of attack of the directional-stability derivative $\\text{C}_{n_\\beta}$ and of the effective-dihedral derivative $\\text{C}_{l_\\beta}$ are shown in figure 6. The values of the derivatives shown in figure 6 are the values at zero sideslip angle. Above an angle of attack of $7^\\circ$ with the vertical tail on, the value of $\\text{C}_{n_\\beta}$ gradually decreased until at an angle of attack of $21^\\circ$ it was about 70 percent of the value at low angles of attack. The vertical tail increased the value of $\\text{C}_{l_\\beta}$ up to an angle of attack of approximately $9^\\circ$. Beyond $9^\\circ$, the value of $\\text{C}_{l_\\beta}$ was lower with the tail on than with the tail off.\n\nThe yawing-moment and the rudder hinge-moment characteristics of the model with the tail on and the rudder deflected to various angles are shown in figures 7 and 8. The rudder was effective throughout the range of angles of attack and angles of sideslip tested. At an angle of attack of $21^\\circ$, however, at angles of sideslip greater than about $9^\\circ$, rudder effectiveness was considerably less than at the lower angles.\n\nA summary of various aerodynamic characteristics$^2$ of the model with and without the vertical tail is given in the following tabulation:\n\n$^2$Except for $\\text{C}_{l_\\beta_{\\text{max}}}$, these are average values at zero angle of attack, zero angle of sideslip, and, where pertinent, zero angle of rudder deflection.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:43:31.357239+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 17, "total_pages": 33, "image_filename": "19930085977_p17.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:43:33.447036+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 2, "total_pages": 29, "image_filename": "19930093789_p2.jpg", "text": "NACA RM No. E9I21\nCONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nSOME EFFECTS OF STATOR CONE ANGLE AND BLADE-TIP LEAKAGE\nON 40-PERCENT-REACTION TURBINES HAVING ROTOR-BLADE CAPS\n\nBy Robert E. English, Robert J. McCready\nand John S. McCarthy\n\nSUMMARY\n\nFor an investigation of the effects of stator cone angle and\ntip leakage on turbine performance, a single-stage turbine having\n40-percent reaction was operated with two stators and two stationary\nshrouds: (1) a stator having a cone angle of $70^\\circ$ and a stator-\nblade height 0.92 of the rotor-blade height; (2) a stator having\na $0^\\circ$ cone angle and a blade height equal to the rotor-blade height;\n(3) a labyrinth, no-leakage shroud; and (4) a cylindrical stationary\nshroud having a radial clearance 0.016 of the blade height. In all\ncases, the rotor blades were equipped with caps that formed a con-\ntinuous, cylindrical rotating shroud.\n\nThe turbine was operated at an entrance temperature of $660^\\circ$ R\nwith total-pressure ratios from 1.25 to 3.70 and equivalent mean\nblade speeds from 188 to 855 feet per second. For this range of\nconditions, the over-all performance of the turbine was determined.\n\nWith the $0^\\circ$-cone-angle stator, the peak brake efficiency was\napproximately 0.04 higher than with the $70^\\circ$-cone-angle stator. With\nthe $70^\\circ$-cone-angle stator, separation probably occurred at the tips\nof the rotor blades; the separation was eliminated by changing the\ncone angle to $0^\\circ$ and making the stator-blade height equal to the\nrotor-blade height. When the labyrinth shroud was replaced by the\ncylindrical shroud, the efficiency was not changed by a measurable\namount.\n\nINTRODUCTION\n\nAs part of a program to develop techniques for designing tur-\nbines of high efficiency, an investigation is being conducted at\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:43:34.844044+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 2, "total_pages": 39, "image_filename": "19930093769_p2.jpg", "text": "NACA RM No. E6L10a CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nCOMPARISON OF PERFORMANCE OF AN-F-58 FUEL AND GASOLINE\n\nIN J34-WE-22 TURBOJET ENGINE\n\nBy Harry W. Dowman and George G. Younger\n\nSUMMARY\n\nAs part of an investigation of the performance of AN-F-58 fuel in various types of turbojet engine, the performance of this fuel in a 3000-pound-thrust turbojet engine has been investigated in an altitude test chamber together with the comparative performance of 62-octane gasoline.\n\nThe investigation of normal engine performance, which covered a range of engine speeds at altitudes from 5000 to 50,000 feet and flight Mach numbers up to 1.00, showed that both the net thrust and average turbine-outlet temperatures were approximately the same for both fuels. The specific fuel consumption and the combustion efficiency at the maximum engine speeds investigated were approximately the same for both fuels at altitudes up to 35,000 feet, but at an altitude of 50,000 feet the specific fuel consumption was about 9 percent higher and the combustion efficiency was correspondingly lower with the AN-F-58 fuel than with gasoline. The low-engine-speed blow-out limits were about the same for both fuels. Ignition of AN-F-58 fuel with the standard spark plug was possible only with the spark plug in a clean condition; ignition was impossible at all flight conditions investigated when the plug was fouled by an accumulation of liquid fuel from a preceding false start. Use of an extended-electrode spark plug provided satisfactory ignition over a slightly smaller range of altitudes and flight Mach numbers than for gasoline with the standard spark plug.\n\nRadial temperature gradients at the turbine outlet were about the same for both fuels at an altitude of 20,000 feet. At an altitude of 50,000 feet, the difference in average temperature between the blade tip and the root was about $240^\\circ$ F greater for AN-F-58 fuel than for gasoline and the spread between the maximum and minimum temperatures at a given radial location was from $150^\\circ$ to $300^\\circ$ F less for AN-F-58 fuel than for gasoline. After 30 hours and 11 minutes of operation with AN-F-58 fuel, 240 grams of hard carbon deposits were built up on the combustor basket near the fuel nozzles.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:43:35.038565+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 20, "total_pages": 42, "image_filename": "19930085938_p20.jpg", "text": "NACA RM No. L9B04\n19\n\n[Figure: Bottom view of an aircraft model]\n\n(b) Bottom view.\nFigure 2.- Concluded.\n\nNACA\nL-56519", "timestamp": "2026-07-22T05:43:36.052961+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 2, "total_pages": 21, "image_filename": "19930082474_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:43:39.403133+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 12, "total_pages": 50, "image_filename": "19930086076_p12.jpg", "text": "10\nNACA RM E9F09\n\n2 percent from the mean pressure. In addition, motion pictures\nwere taken of the combustion process at 2400 frames per second\nduring the investigation of flame holder 13. No indication of com-\nbustor pulsations could be observed from study of these photo-\ngraphs. These results verified audible observations that combustor\npulsations were absent or negligible in flame holders combining\nsurfaces heated to incandescence by flame immersion with a continu-\nous low stream velocity flame path and fuel injection at the flame\nholder.\n\nSUMMARY OF RESULTS\n\nFrom an investigation of the design factors of a 4- by 8-inch\nram-jet combustor, 24 inches in length, the following results were\nobtained:\n\n1. Combustion was stable up to an inlet-air velocity of\n325 feet per second with flame holder having flame-immersed, incan-\ndescent surfaces.\n\n2. A continuous flame path of low stream velocity from the\ndownstream portion of a flame holder to the upstream portion\nimproved combustion efficiency.\n\n3. Silicon-coated molybdenum used as flame holder components\nimmersed in the ram-jet combustion zone had a service life greater\nthan 47 minutes; under similar conditions Inconel lasted less than\n5 minutes.\n\n4. Combustion pulsations were absent or negligible in flame\nholders employing a continuous flame path and fuel injection at\nthe flame holder.\n\n5. Duct-pulsation effects were eliminated by maintaining\nsonic flow in a variable-area inlet to the test unit.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.\n\nREFERENCE\n\n1. Breitwieser, Roland: Performance of a Ram-Jet-Type Combustor\nwith Flame Holders Immersed in the Combustion Zone. NACA\nRM E8F21, 1948.", "timestamp": "2026-07-22T05:43:41.902852+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 13, "total_pages": 98, "image_filename": "19930086073_p13.jpg", "text": "NACA RM A9H04\n\nthroughout the lift range. Directional stability, however, decreased with increasing lift and the model became directionally unstable at high lift coefficients. Rudder effectiveness, on the other hand, remained nearly constant throughout the lift range. The contribution of the vertical tail to the directional stability and the rudder yawing effectiveness could be predicted with reasonable accuracy at zero wing lift.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nREFERENCES\n\n1. Anderson, Adrien E.: An Investigation at Low Speed of a Large-Scale Triangular Wing of Aspect Ratio Two.- I. Characteristics of a Wing Having a Double-Wedge Section With Maximum Thickness at 20-Percent Chord. NACA RM A7F06, 1947.\n\n2. Anderson, Adrien E.: An Investigation at Low Speed of a Large-Scale Triangular Wing of Aspect Ratio Two.- II. The Effect of Airfoil Section Modifications and the Determination of the Wake Downwash. NACA RM A7H28, 1947.\n\n3. Anderson, Adrien E.: Chordwise and Spanwise Loadings Measured at Low Speed on Large Triangular Wings. NACA RM A9B17, 1949.\n\n4. Stephenson, Jack D., and Amuedo, Arthur R.: Tests of a Triangular Wing of Aspect Ratio 2 in the Ames 12-Foot Pressure Wind Tunnel. II - The Effectiveness and Hinge Moments of a Constant-Chord Plain Flap. NACA RM A8E03, 1948.\n\n5. DeYoung, John: Theoretical Additional Span Loading Characteristics of Wings with Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA TN 1491, 1947.", "timestamp": "2026-07-22T05:43:43.854784+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 7, "total_pages": 36, "image_filename": "19930086097_p7.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 5\n\n$\\alpha$ angle of attack\n\n$\\beta$ airfoil trailing-edge angle, measured between chord line and airfoil tangent line at the base\n\n$\\gamma$ ratio of specific heats (1.400 for air)\n\n$\\eta$ ratio of trailing-edge thickness to maximum thickness (h/t)\n\n$\\theta$ local angle of inclination of element on airfoil surface measured relative to the free-stream direction\n\n$\\rho$ mass density\n\nSubscripts\n\nb base of airfoil\n\nu upper surface of airfoil\n\nl lower surface of airfoil\n\n$\\infty$ free stream\n\nTHEORETICAL ANALYSIS\n\nGeneral Considerations\n\nThe high drag of blunt-trailing-edge airfoils at very low Mach numbers is easily explained from existing knowledge of subsonic flows. At these low velocities the minimum drag of a well-designed airfoil consists primarily of skin friction. Any increase in trailing-edge thickness will not significantly alter the skin friction, but will increase the total drag through the addition of base drag and through the elimination of some of the pressure recovery normally obtained over the rear portion of a conventional airfoil. Thus, the usefulness of blunt-trailing-edge sections appears to be restricted to applications where a low drag at subsonic speeds is not of importance.\n\nAt supersonic speeds an increase in trailing-edge thickness will not necessarily lead to a drag increase, as some simple physical considerations will show. Before presenting these considerations, though, it will be advantageous to clarify one particular concept. Throughout this report the shape of a blunt-trailing-edge airfoil will be thought of qualitatively as being formed from a sharp-trailing-edge airfoil by increasing the trailing-edge thickness while maintaining the same chord length, rather than by simply cutting off the trailing edge. This latter viewpoint (removing part of the trailing edge) would needlessly complicate\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:43:44.781672+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 29, "total_pages": 51, "image_filename": "19930085962_p29.jpg", "text": "28\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (181, 110, 768, 906) -->\n\nFigure 10. — Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:43:45.242047+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 32, "total_pages": 36, "image_filename": "19930085912_p32.jpg", "text": "30\nNACA RM No. E9C16\n\n<!-- Image (119, 166, 861, 820) -->\n\nFigure 11. - Maximum inlet-lip temperature observed in investigation.\nBleedback, 8.65 percent; plenum-chamber gas temperature, 1000° F;\ntunnel velocity, 218 feet per second.", "timestamp": "2026-07-22T05:43:52.515166+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 20, "total_pages": 46, "image_filename": "19930085972_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:43:52.804814+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 30, "total_pages": 42, "image_filename": "19930085914_p30.jpg", "text": "NACA RM A9D25\n29\n\nMaximum lift-drag ratio, $(L/D)_{max}$\n18\n14\n10\n0 .2 .4 .6 .8 1.0\nMach number, M\n$R=2.0 \\times 10^6$\n\n18\n14\n10\n0 2 4 6 8 10\nReynolds number, $R \\times 10^{-6}$\n$M=2.0$\n[NACA logo]\n\nFigure 10.— The variation of the maximum lift-drag ratio of the wing-fuselage combination with Mach number and Reynolds number.", "timestamp": "2026-07-22T05:43:58.603000+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 2, "total_pages": 21, "image_filename": "19930091993_p2.jpg", "text": "623.742\nUS8+\n\n# AERONAUTIC SYMBOLS\n\n## 1. FUNDAMENTAL AND DERIVED UNITS\n\n| | Symbol | Metric | | English | |\n| :--- | :---: | :--- | :--- | :--- | :--- |\n| | | **Unit** | **Abbreviation** | **Unit** | **Abbreviation** |\n| Length<br>Time<br>Force | $l$<br>$t$<br>$F$ | meter<br>second<br>weight of 1 kilogram | m<br>s<br>kg | foot (or mile)<br>second (or hour)<br>weight of 1 pound | ft (or mi)<br>sec (or hr)<br>lb |\n| Power<br>Speed | $P$<br>$V$ | horsepower (metric)<br>kilometers per hour<br>meters per second | kph<br>mps | horsepower<br>miles per hour<br>feet per second | hp<br>mph<br>fps |\n\n## 2. GENERAL SYMBOLS\n\n$W$ Weight=$mg$\n$g$ Standard acceleration of gravity=9.80665 m/s$^2$ or 32.1740 ft/sec$^2$\n$m$ Mass=$\\frac{W}{g}$\n$I$ Moment of inertia=$mk^2$. (Indicate axis of radius of gyration $k$ by proper subscript.)\n$\\mu$ Coefficient of viscosity\n\n$\\nu$ Kinematic viscosity\n$\\rho$ Density (mass per unit volume)\nStandard density of dry air, 0.12497 kg-m$^{-4}$s$^2$ at 15$^\\circ$ C and 760 mm; or 0.002378 lb-ft$^{-4}$ sec$^2$\nSpecific weight of \"standard\" air, 1.2255 kg/m$^3$ or 0.07651 lb/cu ft.\n\n## 3. AERODYNAMIC SYMBOLS\n\n$S$ Area\n$S_w$ Area of wing\n$G$ Gap\n$b$ Span\n$c$ Chord\n$A$ Aspect ratio, $\\frac{b^2}{S}$\n$V$ True air speed\n$q$ Dynamic pressure, $\\frac{1}{2}\\rho V^2$\n$L$ Lift, absolute coefficient $C_L=\\frac{L}{qS}$\n$D$ Drag, absolute coefficient $C_D=\\frac{D}{qS}$\n$D_p$ Profile drag, absolute coefficient $C_{Dp}=\\frac{D_p}{qS}$\n$D_i$ Induced drag, absolute coefficient $C_{Di}=\\frac{D_i}{qS}$\n$D_p$ Parasite drag, absolute coefficient $C_{Dp}=\\frac{D_p}{qS}$\n$C$ Cross-wind force, absolute coefficient $C_C=\\frac{C}{qS}$\n\n$i_w$ Angle of setting of wings (relative to thrust line)\n$i_t$ Angle of stabilizer setting (relative to thrust line)\n$Q$ Resultant moment\n$\\Omega$ Resultant angular velocity\n$R$ Reynolds number, $\\rho \\frac{Vl}{\\mu}$, where $l$ is a linear dimension (e.g., for an airfoil of 1.0 ft chord, 100 mph, standard pressure at 15$^\\circ$ C, the corresponding Reynolds number is 935,400; or for an airfoil of 1.0 m chord, 100 mps, the corresponding Reynolds number is 6,865,000)\n$\\alpha$ Angle of attack\n$\\epsilon$ Angle of downwash\n$\\alpha_\\infty$ Angle of attack, infinite aspect ratio\n$\\alpha_i$ Angle of attack, induced\n$\\alpha_0$ Angle of attack, absolute (measured from zero-lift position)\n$\\gamma$ Flight-path angle", "timestamp": "2026-07-22T05:44:02.371319+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 2, "total_pages": 12, "image_filename": "19930091987_p2.jpg", "text": "# AERONAUTIC SYMBOLS\n\n## 1. FUNDAMENTAL AND DERIVED UNITS\n\n| Symbol | Metric | | English | |\n| :--- | :--- | :--- | :--- | :--- |\n| | **Unit** | **Abbreviation** | **Unit** | **Abbreviation** |\n| Length . . . . . . . . . . | $l$ | meter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T05:44:08.815813+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 18, "total_pages": 33, "image_filename": "19930085977_p18.jpg", "text": "NACA RM L9H22 CONFIDENTIAL 17\n\n[Figure: A pictorial view showing sponge-wiper-seal installation on a model with a sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section. The model includes a vertical tail fin mounted on a circular base, with three additional vertical fins in the background. A star-shaped shadow is cast beneath the base. The NACA logo and identifier \"L-61939\" are visible in the lower right corner of the image.]\n\nFigure 5.— A pictorial view showing sponge-wiper-seal installation on the model with $0^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:44:09.219787+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 9, "total_pages": 28, "image_filename": "19930086092_p9.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 7\n\n| Parameter | Value | |\n| :--- | :--- | :--- |\n| | Tail off | Tail on |\n| $C_{n_\\beta}$ | $-0.0012$ | $0.0025$ |\n| $C_{l_\\beta_{max}}$ | $-.0033 (\\Gamma_e \\cong 15^\\circ)$ | $-.0026 (\\Gamma_e \\cong 12^\\circ)$ |\n| $dC_{l_\\beta}/d\\alpha$ | $-.00022$ | $-.00027$ |\n| $C_{Y_\\beta}$ | $-.0008$ | $-.0061$ |\n| $C_{n_{\\delta_r}}$ | $---$ | $-.0011$ |\n| $C_{h_{\\delta_r}}$ | $---$ | $-.0029$ |\n| $C_{h_{\\alpha_t}}$ | $---$ | $-.0006$ |\n\nDISCUSSION\n\nIn order to enable some generalizations to be made of the characteristics of the swept-back vertical tail and provide a basis for estimating the effects of geometric changes of the tail surface, various characteristics of the vertical tail have been computed theoretically. In the following discussion, the various factors involved in the theoretical computations are discussed and the experimental and theoretical results are compared.\n\nDirectional Stability\n\nThe stabilizing effect of a vertical tail can be estimated by use of the following equation:\n\n$$ \\Delta C_{n_{\\beta_t}} = \\left( \\frac{dC_N}{d\\alpha} \\right)_t \\times \\left( 1 + \\frac{d\\alpha}{d\\beta} \\right) \\times \\frac{q_t}{q} \\times \\frac{S_t}{S} \\times \\frac{l_t}{b} \\quad (1) $$\n\nA way of estimating the tail normal-force-curve slope $\\left( \\frac{dC_N}{d\\alpha} \\right)_t$ is by the use of the data in reference 9. These data, however, are for complete wings and, therefore, in applying these data, the end-plate effect of the fuselage is assumed to be total. This procedure necessarily will require a correction to account for the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:44:14.787573+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 13, "total_pages": 50, "image_filename": "19930086076_p13.jpg", "text": "Atmospheric\nexhaust\nStation 3\nQuartz\nobservation\nwindow\nStation 2\nStation 1\nWater spray\nFlame holder\nWater jacket\nFuel injector\nCombustion-air supply\n24\"\n8\"\n4\"\nLucite\nobservation\nwindow\nIgnitor\nCenter\nplate\nNeedle-type\nvariable-\narea\ndiffuser\nGlass observation\nwindow\nMovable-wall\ndiffuser\nSurge tank\nNACA\nFigure 1. - Experimental setup of 4- by 8-inch combustor and auxiliary ducting.\nNACA RM E9F09\n11", "timestamp": "2026-07-22T05:44:16.115030+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 3, "total_pages": 21, "image_filename": "19930082474_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1799\n\nON THE FLYING QUALITIES OF HELICOPTERS\n\nBy John P. Reeder and F. B. Gustafson\n\nSUMMARY\n\nThe flying-qualities problems of current helicopters as observed during flight are discussed. These problems have been found to be (1) instability with angle of attack in forward flight, (2) control sensitivity in hovering (particularly for the smaller helicopters), and (3) control forces following control movement in maneuvers. Some discussion is also given of tentative remedies for the most outstanding deficiencies.\n\nINTRODUCTION\n\nExperience indicates that, in its present stage of development, the helicopter is different from and more difficult to fly than most airplanes. The difficulty seems to arise from several sources: (1) The helicopter has one additional control (collective pitch) to be operated. (2) The power controls (collective pitch and throttle) must be used almost continuously in conjunction with the flight controls during operations near the ground, chiefly because of the rapid variation of power required with airspeed in the speed range normally used in these operations. (3) The helicopter has undesirable stability characteristics in forward flight which would not be acceptable in an airplane. (4) Undesirable control characteristics exist in both forward flight and hovering. Hovering flight also introduces a new and unique problem of apparent lag in control response which is, however, somewhat analogous to formation flying with airplanes.\n\nThe National Advisory Committee for Aeronautics has long been interested in stability and control problems and in setting up requirements for the satisfactory stability and control characteristics for airplanes. This work is now being extended to the helicopter because the helicopter eventually must meet requirements parallel to those for the airplane in order to reach its potential capabilities. Although airplane requirements may not be applicable to helicopters in a specific manner, the underlying reason for setting up the requirements applies to both airplane and helicopter.", "timestamp": "2026-07-22T05:44:16.905777+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 3, "total_pages": 29, "image_filename": "19930093789_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM No. E8I21\n\nthe NACA Lewis laboratory to determine the performance of a\nsingle-stage, cold-air turbine in which the relative importance of\ndesign variables is evaluated by incorporating systematic changes\nin the blade design. This investigation should provide data for\nthe design of turbines and indicate where further studies may most\nprofitably be concentrated.\n\nThe first phase of this program is an investigation of the\neffects of stator cone angle and tip leakage on turbine performance.\nIn order to observe the effect of stator cone angle on over-all\nturbine performance, two stators were used: a stator with a $70^\\circ$\ncone angle and a stator-blade height 0.92 of the rotor-blade height;\nand a stator with a $0^\\circ$ cone angle and a stator-blade height equal\nto the rotor-blade height. The effect of tip leakage on turbine\nperformance was investigated by operating the turbine with a laby-\nrinth no-leakage shroud and a cylindrical shroud.\n\nOf the possible combinations of these variables, three con-\nfigurations of the turbine were studied: (1) the $70^\\circ$-cone-angle\nstator and the labyrinth shroud; (2) the $0^\\circ$-cone-angle stator and\nthe labyrinth shroud; and (3) the $0^\\circ$-cone-angle stator and the\ncylindrical shroud. For the investigation of these three config-\nurations, the turbine was operated at total-pressure ratios of 1.25\nto 3.70 and equivalent mean blade speeds of 188 to 855 feet per\nsecond using air at a temperature of $660^\\circ$ R.\n\nSYMBOLS\n\nThe following symbols are used in this report:\n\n| | |\n| :--- | :--- |\n| a | velocity of sound, (ft/sec) |\n| E | turbine-shaft work, (Btu/lb) |\n| g | standard acceleration due to gravity, 32.17 (ft/sec$^2$) |\n| h | specific enthalpy, (Btu/lb) |\n| J | mechanical equivalent of heat, 778.3 (ft-lb/Btu) |\n| p | absolute pressure, (lb/sq ft) |\n| T | absolute temperature, ($^\\circ$R) |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:44:17.120605+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 30, "total_pages": 51, "image_filename": "19930085962_p30.jpg", "text": "NACA RM A9E05\nCONFIDENTIAL\n\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\nLift coefficient, $C_L$\n\n$\\delta_e$\n(deg)\n$\\circ$ 0\n$\\square$ 2\n$\\diamond$ 4\n$\\triangle$ 6\n$\\nabla$ 10\n$\\triangleright$ 20\n\n-12 -8 -4 0 4 8 12\nAngle of attack, $a$, deg\n\n.08 .04 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28 -.32\nPitching-moment coefficient, $C_m$\n\n(a) $C_L$ vs $a$, $C_L$ vs $C_m$.\n\nFigure 11. — The effect of elevator deflection on the aerodynamic characteristics of the tail at a Mach number of 0.94.\n\nCONFIDENTIAL\n29", "timestamp": "2026-07-22T05:44:21.306985+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 21, "total_pages": 46, "image_filename": "19930085972_p21.jpg", "text": "NACA RM L9B18\n\n19\n\n[Figure: A model aircraft viewed from the rear, mounted on a stand. A label in the upper right corner reads \"NACA I-52658\".]\n\n(b) Rear view.\n\nFigure 5.— Concluded.", "timestamp": "2026-07-22T05:44:24.010258+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 8, "total_pages": 36, "image_filename": "19930086097_p8.jpg", "text": "6 CONFIDENTIAL NACA RM A9H11\n\nmatters because of the accompanying changes in the reference area on which force coefficients are based.\n\nA double-wedge airfoil and a corresponding blunt-trailing-edge airfoil of equal chord are illustrated in figure 1. For simplicity, a common angle is used between all flat surfaces of the blunt-trailing-edge airfoil and the chord line. In comparison with a double-wedge airfoil of the same thickness ratio, the surfaces of the airfoil with a thick trailing edge are inclined at a smaller angle with respect to the chord line, thereby reducing the pressure drag of the profile contour forward of the base. On the other hand, it is apparent that considerable base drag may be introduced by employing a blunt trailing edge. Since the skin-friction drag is essentially the same for blunt- and sharp-trailing-edge airfoils, it follows that the profile drag will be lowered if the increase in base drag is less than the afore-mentioned reduction in pressure drag.\n\nIt will be illustrative to consider a particular example in order to demonstrate that, in certain cases at least, the profile drag can be reduced by increasing the trailing-edge thickness. A comparison of the drag of a 10-percent-thick double-wedge airfoil and a 10-percent thick wedge airfoil at a Mach number of 5 will serve to establish this point. By employing the customary shock-expansion method to calculate the wave-drag components of these two profiles, and by making the obviously conservative assumption that a vacuum exists at the base of the wedge, the following drag coefficients are obtained:\n\n| Airfoil | Wave drag of profile forward of base | Base drag | Profile drag |\n| :--- | :--- | :--- | :--- |\n| Double wedge | 0.0091 | 0 | $0.0091 + c_{df}$ |\n| Wedge (blunt trailing edge) | .0024 | .0057 (vacuum at base) | $.0081 + c_{df}$ |\n\nThis simple example showing lower drag for the wedge airfoil clearly illustrates two facts: First, at relatively high Mach numbers the conventional double-wedge section is not the optimum section for a given thickness ratio, and, second, the use of a blunt trailing edge can reduce profile drag by a substantial amount in this Mach number range. By using a more indirect method and by considering the characteristics at a Mach number of 8, these same two results have previously been pointed out by Ivey in reference 4.\n\nIf the above calculation were performed for an airfoil with thickness ratio much less than 10-percent or for a Mach number much less than 5, then the overly crude approximation of a vacuum at the base would indicate a higher drag for the wedge profile. In order for the use of a thick trailing edge to reduce the profile drag of thinner airfoils, or of airfoils in the lower supersonic Mach number range, the base drag must necessarily be considerably less than that represented by a vacuum.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:44:26.448251+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 14, "total_pages": 98, "image_filename": "19930086073_p14.jpg", "text": "```markdown\n12\nNACA RM A9H04\n\nTABLE I.- GEOMETRIC DATA OF MODEL USED IN THE INVESTIGATION\n\n| Item | Wing alone | Wing with body and vertical tail |\n| :--- | :--- | :--- |\n| **Wing** | | |\n| Span, feet | 25.00 | 25.00 |\n| Area, square feet | 307 | 307 |\n| Area exposed outside of fuselage, square feet | --- | 211 |\n| Mean aerodynamic chord, feet | 16.37 | 16.37 |\n| Angle of incidence, degrees | --- | 0 |\n| Aspect ratio | 2.04 | 2.04 |\n| **Body** | | |\n| Length, feet | --- | 56.16 |\n| Maximum diameter, feet | --- | 4.49 |\n| Fineness ratio | --- | 12.50 |\n| Ratio of maximum diameter to wing span | --- | 0.18 |\n| **Split-flap-type controls** | | |\n| Semispan, feet | 10.83 | 8.70 |\n| Total area, square feet | 57.80 | 46.46 |\n| Total wing area affected by control surface, square feet | 301.5 | 205.5 |\n| **Vertical tail** | | |\n| Total area to body center line, square feet | --- | 66.90 |\n| Aspect ratio (total) | --- | 1.00 |\n| Rudder area (exposed), square feet | --- | 11.50 |\n| Rudder area (total), square feet | --- | 13.62 |\n| Tail length ($\\bar{c}/4$ to tail center of pressure), feet | --- | 13.41 |\n| Tail length ($\\bar{c}/4$ to rudder hinge line), feet | --- | 18.52 |\n\nNACA\n```", "timestamp": "2026-07-22T05:44:27.146248+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 3, "total_pages": 39, "image_filename": "19930093769_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM No. E8L10a\n\nINTRODUCTION\n\nThe need of the armed forces for a fuel for turbojet engines that would be available in greater quantities than those now in use has led to the consideration of a new fuel specification, AN-F-58, which has much wider limits than the specifications of current turbojet fuels. An extensive program to determine the suitability of fuel conforming to the AN-F-58 specification is being conducted at the NACA Lewis laboratory on several types of turbojet engine and combustor.\n\nAs part of this general program, the performance of AN-F-58 fuel was investigated during August and September 1948 in a 3000-pound-thrust turbojet engine in an altitude test chamber and the results are reported. This investigation included, for comparison, the performance of the engine with 62-octane gasoline. The normal performance of the engine was investigated over a range of engine speeds at altitudes from 5000 to 50,000 feet and for flight Mach numbers from 0.22 to 1.00. The low-engine-speed blow-out limits, the altitude starting limits, and the turbine-outlet gas-temperature gradients for both fuels were determined at several simulated flight conditions. The use of AN-F-58 fuel for cold-weather starts at zero-ram, sea-level conditions and the amount of carbon deposited in the combustor with this fuel were also investigated.\n\nFUELS\n\nTwo types of fuel conforming to the AN-F-58 specification were used in the investigation. These two fuels are designated NACA numbers 48-206 and 48-210 and the analysis of each, with the AN-F-58 specifications, is given in table I. The gasoline that was used for comparative performance data was a clear, 62-octane gasoline and its analysis is also included in table I.\n\nAPPARATUS AND INSTRUMENTATION\n\nThe investigations were conducted with a modified experimental model of the J34 engine and a J34-WE-22 engine (hereinafter designated engine A and engine B, respectively) having a rated speed of 12,500 rpm and a thrust rating of 3000 pounds at zero-ram, sea-level conditions. The main components of the engines include an 11-stage axial-flow compressor, a double-annulus combustor, and a two-stage turbine. The two engines were similar in general design and differed only in minor details; both engines incorporated the manufacturer's recent modifications to the combustor basket and the eleventh-stage compressor blading. The combustor basket has 3/32-inch antilocking holes on both the inner and outer rings.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:44:28.645183+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 31, "total_pages": 42, "image_filename": "19930085914_p31.jpg", "text": "```markdown\n30\nNACA RM A9D25\n\n<!-- Image (169, 90, 806, 806) -->\n\nFigure 11.- The variation of the aerodynamic center\n($C_L=0$) of the wing-fuselage combination with\nMach number and Reynolds number.\n```", "timestamp": "2026-07-22T05:44:29.602514+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 3, "total_pages": 21, "image_filename": "19930091993_p3.jpg", "text": "REPORT 928\n\nANALYSIS OF PERFORMANCE OF JET ENGINE FROM  \nCHARACTERISTICS OF COMPONENTS  \nII—INTERACTION OF COMPONENTS AS DETERMINED  \nFROM ENGINE OPERATION\n\nBy ARTHUR W. GOLDSTEIN, SUMNER ALPERT, WILLIAM BEEDE  \nand KARL KOVACH\n\nFlight Propulsion Research Laboratory  \nCleveland, Ohio", "timestamp": "2026-07-22T05:44:30.420625+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 33, "total_pages": 36, "image_filename": "19930085912_p33.jpg", "text": "```markdown\nNACA RM No. E9C16\n31\n\n11.05\n\n| Curve | A | B |\n| :--- | :--- | :--- |\n| Model-air total temperature, $^\\circ$F | 41.3 | 39.4 |\n| Tunnel velocity, ft/sec | 410 | 220 |\n| Liquid-water content, gram/cu m | 1.0 | 0.7 |\n\n| Tunnel velocity (ft/sec) | Angle of attack (deg) |\n| :--- | :--- |\n| $\\circ$ 200 | 0 |\n| $\\square$ 280 | 0 |\n| $\\diamond$ 360 | 0 |\n| $\\triangle$ 410 | 0 |\n| $\\triangledown$ 200 | 8 |\n| $\\nabla$ 280 | 8 |\n\nPlenum-chamber gas temperature, $^\\circ$F\nBleedback, percent\n\n[Figure: Graph showing Plenum-chamber gas temperature vs. Bleedback percent with curves A and B and various data points]\n\nNACA\n\nFigure 12. - Bleedback required for ice prevention as function of plenum-chamber gas temperature for free-stream total temperature of $0^\\circ$ F.\n```", "timestamp": "2026-07-22T05:44:30.629207+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 19, "total_pages": 33, "image_filename": "19930085977_p19.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:44:32.987747+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 3, "total_pages": 12, "image_filename": "19930091987_p3.jpg", "text": "REPORT 922\n\nCHARACTERISTICS OF LOW-ASPECT-RATIO WINGS\nAT SUPERCRITICAL MACH NUMBERS\n\nBy JOHN STACK and W. F. LINDSEY\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.", "timestamp": "2026-07-22T05:44:35.205033+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 21, "total_pages": 42, "image_filename": "19930085938_p21.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:44:40.659767+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 31, "total_pages": 51, "image_filename": "19930085962_p31.jpg", "text": "30\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (155, 109, 792, 906) -->\n\n(b) $C_L$ vs $C_D$.\nFigure 11.- Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:44:44.106786+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 4, "total_pages": 21, "image_filename": "19930082474_p4.jpg", "text": "2\nNACA TN No. 1799\n\nDuring the past several years rotor-performance investigations have been conducted at the Langley Laboratory using the two-place utility aircraft shown in figure 1. The rotors flown during this period differed in solidity, airfoil section, twist, and blade-surface rigidity. During the course of these investigations attention was drawn to certain stability characteristics of the helicopter which have long been considered unacceptable for airplanes, some interesting control characteristics and flight phenomena were revealed, and some very limited measurements of stability and control characteristics were made. Recently, moreover, the status of the performance investigations was such that instrumentation could be installed to get more detailed information on flight characteristics. Some of the data obtained with this instrumentation, experience with other helicopter types, knowledge of British experiments (an example is reference 1), and information from translations of German papers (references 2 and 3) have been used in formulating the ideas presented in this paper. A further valuable source of experience concerning the characteristics of the helicopter in maneuvers has been afforded by pull-up tests for load-factor determination; these tests have been made by the CAA with the assistance of the NACA. With this background, the present paper should help to indicate the most fruitful lines for immediate further study.\n\nOBSERVATIONS OF FLYING QUALITIES\n\nLongitudinal Stability in Forward Flight\n\nDuring the course of the performance investigations, considerable flying was done at relatively high speeds approaching the limits imposed by blade stalling. Steady conditions were difficult to hold because of a strong tendency of the machine to diverge in pitch, this divergence creating the impression of balancing on a ball. This characteristic seemed far more pronounced with some of the rotors tested than with others but was always troublesome. Upward pitching was most troublesome as it frequently precipitated or intensified stalling, which increased the tendency to pitch up and was accompanied by rather violent periodic stick forces and vibration. The forward displacement of the control from trim necessary to check some of these pitching motions suggested that a short delay in applying corrective control would allow a maneuver severe enough that control would be lost. Although there seemed ample control to stop downward pitching, an excessive amount of forward control was again required in order to check the subsequent upward pitching. These characteristics indicated a pronounced type of longitudinal instability.\n\nThe tendency of the helicopter to depart from the trim speed and the necessity of applying appreciable control deflection against a pitching maneuver involving acceleration, initiated either by control or by gusty air, is apparent throughout the speed range normally used in forward flight. This tendency becomes much less pronounced, however, at the lower speeds.", "timestamp": "2026-07-22T05:44:45.326925+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 22, "total_pages": 46, "image_filename": "19930085972_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:44:45.601900+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 4, "total_pages": 29, "image_filename": "19930093789_p4.jpg", "text": "```markdown\nNACA RM No. E8I21 CONFIDENTIAL 3\n\nU mean rotor-blade speed, (ft/sec)\n\nV absolute velocity of gas, (ft/sec)\n\nW gas velocity relative to rotor, (ft/sec)\n\nw weight flow of air, (lb/sec)\n\n$\\alpha$ angle of absolute velocity measured from direction of rotor-blade motion, (deg)\n\n$\\beta$ angle of relative velocity measured from direction of rotor-blade motion, (deg)\n\n$\\Delta$ prefix denoting change\n\n$\\eta$ brake efficiency\n\n$\\rho$ mass density, (slug/cu ft)\n\nSubscripts:\n\n0 NACA sea-level air; 2116 (lb/sq ft) and $518.4^\\circ$ R\n\n1 stator entrance\n\n2 rotor entrance\n\n3 rotor exit\n\nj ideal jet\n\ns isentropic\n\nu tangential component\n\nx axial component\n\nSuperscript:\n\n' total, or stagnation, state\n\nTURBINE DESIGN\n\nThe aerodynamic design of the turbine is for the following conditions:\n\nCONFIDENTIAL\n\n```", "timestamp": "2026-07-22T05:44:47.012004+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 10, "total_pages": 28, "image_filename": "19930086092_p10.jpg", "text": "8 CONFIDENTIAL NACA RM A9F14\n\nimperfect end-plate effect of the fuselage.⁵ At the present time, however, the end-plate effect of a fuselage on a vertical tail has not been determined either theoretically or experimentally; hence, no such correction will be used. Accordingly, the value of normal-force-curve slope of the swept-back tail of 0.041 per degree, obtained from reference 9 assuming the fuselage to exert a total end-plate effect, will be used. (This compares to the lift-curve slope of 0.042 per degree measured for the full-span wing of the same plan form, reference 2.)\n\nIn order to evaluate the sidewash term $1 + \\frac{d\\sigma}{d\\beta}$ and the effective dynamic pressure term $\\frac{q_t}{q}$ for equation (1), surveys were made of the air flow in the region of the tail. The integrated value of the average dynamic pressure $\\frac{q_t}{q}$ was found to be very nearly equal to 0.9. This is consistent with unswept configurations. No consistent deviation from zero sidewash was measurable, that is, $\\frac{d\\sigma}{d\\beta} = 0$.\n\nBy the use of the above values of $\\left(\\frac{dC_N}{d\\alpha}\\right)_t$, $\\frac{d\\sigma}{d\\beta}$, and $\\frac{q_t}{q}$, together with the dimensional characteristics of the model, the stabilizing effect of the vertical tail (equation (1)) was computed which can be compared with the experimental result:\n\nTheoretical $\\Delta C_{n\\beta_t} = 0.0048$\n\nExperimental $\\Delta C_{n\\beta_t} = 0.0037$\n\nRudder Effectiveness\n\nThe rudder effectiveness can be estimated from the following equation:\n\n$$\nC_{n\\delta_r} = -\\left(\\frac{dC_N}{d\\alpha}\\right)_t \\times \\alpha_\\delta \\times \\frac{q_t}{q} \\times \\frac{S_t}{S} \\times \\frac{l_t}{b}\n\\tag{2}\n$$\n\nThe same value of the tail normal-force-curve slope will be used, 0.041 per degree (reference 9, assuming a total end-plate effect of the fuselage).\n\nThe flap-effectiveness parameter $\\alpha_\\delta$ can be estimated by applying a correction for the effect of sweep to the value of $\\alpha_\\delta$ of an equivalent unswept, flapped airfoil. The parameter $\\alpha_\\delta$ for an equivalent\n\n⁵This is in contrast to the usual design procedure for unswept tails in which the starting point is zero end-plate effect, a correction factor then being applied to increase the aspect ratio to account for the end-plate effect of the horizontal tail.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:44:56.011601+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 9, "total_pages": 36, "image_filename": "19930086097_p9.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 7\n\nCalculation of Profile Drag for Various Base Pressure Coefficients\n\nBecause very little is known about the base pressure in two-dimensional flow, the subsequent theoretical analysis will consist primarily of calculating the drag reduction that is possible for various base pressure coefficients. By comparing these results with the small amount of experimental data that are available, some indication can be obtained of the actual profile-drag reductions that may be expected.\n\nIn order to obtain a simple expression for the wave drag of the profile forward of the base, the linearized supersonic airfoil theory will be employed at present. If desired, a slightly more refined drag analysis could be made by using the conventional second-order theory of Busemann. Such an analysis, however, is not necessary for drag calculations since the drag coefficients of blunt- and sharp-trailing-edge airfoils differ even when only first-order terms are considered. The local pressure coefficient is then\n\n$$\nP = C_1 \\theta \\tag{1}\n$$\n\nwhere\n\n$$\nC_1 = \\frac{2}{\\sqrt{M_\\infty^2 - 1}} \\tag{2}\n$$\n\nand $\\theta$, the local angle of inclination, is measured positive for elements facing the oncoming wind. The particular airfoil sections that will be used in the drag calculations consist of straight-side symmetrical contours, as illustrated in figure 1. At present only conditions at zero angle of attack will be considered, so that by symmetry $\\theta_u = \\frac{dy}{dx} = -\\theta_l$ and the pressure drag of the contour forward of the base becomes\n\n$$\nc_{d_w} = 2C_1 \\int_0^1 \\left( \\frac{dy}{dx} \\right)^2 d\\left( \\frac{x}{c} \\right) \\tag{3}\n$$\n\nThe profile drag is the sum of the base drag, skin-friction drag, and pressure drag of the profile forward of the base. For an airfoil of thickness ratio $t/c$ and trailing-edge thickness $h = \\eta t$, the pressure drag from equation (3) becomes\n\n$$\nc_{d_w} = \\frac{(t/c)^2}{\\sqrt{M_\\infty^2 - 1}} (2 - \\eta)^2 \\tag{4}\n$$\n\nand the profile drag is\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:44:57.615174+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 32, "total_pages": 42, "image_filename": "19930085914_p32.jpg", "text": "NACA RM A9D25\n31\n\n<!-- Image (159, 139, 903, 733) -->\n\n(a) $C_L$ vs $C_D$.\n\nFigure 12.- The effect of Reynolds number on the aerodynamic characteristics of the wing-fuselage combination at a Mach number of 0.20.", "timestamp": "2026-07-22T05:44:58.852419+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 15, "total_pages": 98, "image_filename": "19930086073_p15.jpg", "text": "NACA RM A9H04\n13\n\nTABLE II.- BODY COORDINATES\n[Stations and radii are in percent\nof the total length.]\n\n| Station | | Radius |\n| :--- | :--- | :--- |\n| 0 | 100.00 | 0 |\n| .625 | 99.375 | .26 |\n| 1.25 | 98.75 | .42 |\n| 2.50 | 97.50 | .70 |\n| 5.00 | 95.00 | 1.15 |\n| 7.50 | 92.50 | 1.54 |\n| 10.00 | 90.00 | 1.86 |\n| 15.00 | 85.00 | 2.41 |\n| 20.00 | 80.00 | 2.86 |\n| 25.00 | 75.00 | 3.22 |\n| 30.00 | 70.00 | 3.51 |\n| 35.00 | 65.00 | 3.73 |\n| 40.00 | 60.00 | 3.88 |\n| 45.00 | 55.00 | 3.97 |\n| 50.00 | - - - | 4.00 |\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T05:45:00.914255+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 3, "total_pages": 34, "image_filename": "19930082472_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1797\n\nA STUDY OF STALL PHENOMENA ON A $45^\\circ$  \nSWEPT-FORWARD WING\n\nBy Gerald M. McCormack and Woodrow L. Cook\n\nSUMMARY\n\nAn investigation has been made to determine the underlying causes of the undesirable longitudinal characteristics of a $45^\\circ$ swept-forward wing in the moderate and high lift-coefficient range. Three-component force data, pressure-distribution data, tuft studies, and boundary-layer measurements were obtained to enable a detailed correlation between separation phenomena and the longitudinal characteristics of the swept-forward wing.\n\nIn the moderate lift range, the occurrence of turbulent separation caused a chordwise redistribution of load over the inboard sections. This, in turn, caused increases in drag and a rearward shift of aerodynamic center but caused no loss of lift. In the high lift range, the occurrence of leading-edge separation caused a loss of section lift that occurred first over the inboard sections and traveled outward as angle of attack was increased. This caused very large increases in drag, a decreased lift-curve slope, and, due to the changes in spanwise loading, caused an extremely large forward shift of aerodynamic center.\n\nIn order to improve the longitudinal characteristics of the swept-forward wing, both forms of separation must be postponed. The evidence indicates that effort should be directed first toward postponing leading-edge separation. Only after leading-edge separation is adequately postponed should control of the turbulent boundary layer be attempted.\n\nINTRODUCTION\n\nOne of the most difficult problems connected with the design of an airplane employing swept wings is the improvement of the poor", "timestamp": "2026-07-22T05:45:04.937035+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 4, "total_pages": 39, "image_filename": "19930093769_p4.jpg", "text": "NACA RM No. E3L10a CONFIDENTIAL 3\n\nBecause of frequent failures of the pump and the governor assembly in the fuel system, this assembly was replaced by a standard J33 fuel pump, throttle, and barometric control.\n\nIn the interest of expediency, an existing setup, which incorporated an afterburner, was used for all performance runs conducted with engine A. The afterburner was inoperative for all investigations reported herein and functioned merely as an engine tail pipe. The adjustable nozzle of this afterburner was fixed in a position that provided approximately limiting turbine-inlet gas temperature at maximum rated speed, zero-ram, sea-level conditions and was held locked in this position throughout the investigation. Engine B, which replaced engine A after a turbine failure, was equipped with a nonafterburning tail pipe and an NACA adjustable-area nozzle. The adjustable-area nozzle of this engine was also locked in position to provide approximately limiting turbine-inlet gas temperature at zero-ram, sea-level conditions.\n\nThe general arrangement of the engine setup in the altitude test chamber is shown in figures 1 and 2. The test chamber is 10 feet in diameter and 57 feet long and includes air-inlet and exhaust-outlet piping, an inlet-air honeycomb, blow-out patches, cooling-air duct, and necessary instrumentation bulkheads. The engine is mounted on a thrust stand and is flexibly connected to a forward baffle (fig. 2), which confines the air flow to the engine inlet and provides a means of maintaining a pressure difference across the engine. A rear baffle is installed around the engine tail pipe to serve as a heat barrier and to prevent recirculation of exhaust gases around the engine. The engine thrust is balanced and measured through a lever arrangement by a null-type air-pressure diaphragm. The forces introduced into the thrust-measuring system by the pressure differences across the two baffles were determined by calibration.\n\nFor the investigation of the ignition characteristics of the engine, three different spark plugs, which are shown in figure 3, were used. The spark plug shown in figure 3(a) is a standard spark plug for the J34 engine and that shown in figure 3(b) is a standard plug that was modified by milling longitudinal slots in the outer electrode shell to allow fuel drainage. The spark plug shown in figure 3(c) is an extended-electrode type, which was fabricated at this laboratory. The spark gap for this spark plug was from 0.090 to 0.100 inch. The standard engine ignition system, which incorporates two high-tension coils, was used with all three spark plugs.\n\nAir flow, fuel flow, engine speed, and temperature and pressure measurements at various stations in the engine and test\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:45:06.385175+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 34, "total_pages": 36, "image_filename": "19930085912_p34.jpg", "text": "32\nNACA RM No. E9C16\n\n[Figure: A graph showing inlet-lip temperature distributions. The graph has two curves labeled \"Dry\" and \"Wet\". The vertical axis is labeled \"Temperature, °F\" with markings at 200 and 400. The NACA logo is present at the bottom right of the graph.]\n\nFigure 13. - Inlet-lip temperature distributions for dry air and icing conditions. Bleedback, 4.4 percent; plenum-chamber gas temperature, $1000^\\circ$ F; icing condition, liquid-water content of 0.5 gram per cubic meter at free-stream total temperature of $0^\\circ$ F.", "timestamp": "2026-07-22T05:45:06.767292+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 14, "total_pages": 50, "image_filename": "19930086076_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:45:09.427453+00:00"}