PDF/ocr_dataset_1079.jsonl

{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 3, "total_pages": 32, "image_filename": "19930085992_p3.jpg", "text": "NACA RM L9E17\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nPRELIMINARY EXPERIMENTAL INVESTIGATION OF EFFECTS OF\n\nAERODYNAMIC SHAPE OF CONCENTRATED WEIGHTS ON\n\nFLUTTER OF A STRAIGHT CANTILEVER WING\n\nBy John L. Sewall and Donald S. Woolston\n\nSUMMARY\n\nResults are presented to show the effect on flutter characteristics of variation of the aerodynamic shape of concentrated weights rigidly mounted on a simplified wing structure. The model was mounted as a rigid cantilever and tested with weights that were $7\\frac{1}{2}$ percent and 5 percent of the weight of the wing. In regard to shape, two general types of weights, having similar mass and moment-of-inertia properties, were employed: one a streamlined body resembling in shape an external wing fuel tank and the other a chosen nonstreamlined, or blunt, body. Approximately 20 flutter tests were conducted in a preliminary program at low Mach numbers with weights varied over a wide range of spanwise positions; an additional chordwise position was included at the wing tip. Results show only small changes in flutter speed and flutter frequency due to radical changes in the aerodynamic shape of concentrated weights. A large reduction in flutter speed is shown as relatively light concentrated weights are moved nearer the tip, with only a small change in flutter frequency. Results also demonstrate, experimentally, a considerable influence of moment of inertia on flutter speed and flutter frequency.\n\nINTRODUCTION\n\nThe installation of large external fuel tanks on airplane wings has caused attention to be directed to the possible influence of these tanks on certain aeroelastic properties of the wing. For example, an investigation of the cause of wing failure for a certain airplane having an external fuel tank at the wing tip (with the tank in an", "timestamp": "2026-07-22T06:52:17.277243+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 12, "total_pages": 30, "image_filename": "19930085970_p12.jpg", "text": "10\nCONFIDENTIAL\nNACA RM A9E09\n\n4. McCormack, Gerald M., and Walling, Walter C.: Aerodynamic Study\nof a Wing-Fuselage Combination Employing a Wing Swept Back\n63°.- Investigation of a Large-Scale Model at Low Speed.\nNACA RM A8D02, 1949.\n\n5. Goldstein, S., and Young A. D.: The Linear Perturbation Theory\nof Compressible Flow With Applications to Wind Tunnel Inter-\nference. R.& M. 1909, British A.R.C., 1943.\n\n6. DeYoung, John: Theoretical Additional Span Loading Character-\nistics of Wings with Arbitrary Sweep, Aspect Ratio, and Taper\nRatio. NACA TN 1491, 1947.\n\n7. Cohen, Doris: The Theoretical Lift of Flat Swept-Back Wings at\nSupersonic Speeds. NACA TN 1555, 1948.\n\n8. Theordorsen, Theodore, and Regier, Arthur: Experiments on Drag\nof Revolving Disks, Cylinders, and Streamline Rods at High\nSpeeds. NACA Rep. 793, 1944.\n\n9. Kleissas, John: Charts of the Zero-Lift Drag of Supersonic\nSweptback Wings for Various Taper Ratios. Northrup Aircraft,\nInc., Rep. GML09, 1947.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:20.325269+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 23, "total_pages": 92, "image_filename": "19930085951_p23.jpg", "text": "```markdown\nNACA RM L9D29\n\nCONFIDENTIAL\n~~UNCLASSIFIED~~\n\nRAKE CONTAINING BOTH TOTAL\nAND STATIC PRESSURE TUBES\n\nPROPELLER HUB AND SPINNER\nARE FREE TO MOVE AXIALLY, RESTRAINED\nBY PNEUMATIC THRUST CAPSULE\n\nFAIRING MAY BE EITHER FIXED\nOR FREE TO MOVE AXIALLY, RESTRAINED\nBY PNEUMATIC FAIRING-DRAG CAPSULE\n\n[Figure: Diagram showing a propeller configuration with dimensions and labels]\n\nTUNNEL\nFLATS\n\nTENSION\nSTRUT\n\nTUNNEL WALL\n\n16 FT.D. 10 FT.\n\n26\"D. 36\"D.\n\n36\" D.\n\n303\"\n\nSUPPORT STRUT FASTENED\nRIGIDLY TO TUNNEL WALL\n\nNACA\n\nFigure 3.- Configuration of 2000-horsepower dynamometer for tests of propellers in the Langley 16-foot\nhigh-speed tunnel.\nCONFIDENTIAL\n\n21\n```", "timestamp": "2026-07-22T06:52:22.137814+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 10, "total_pages": 46, "image_filename": "19930085983_p10.jpg", "text": "8 CONFIDENTIAL NACA RM A9I27\n\ncf angle of attack for differential elevon deflections of $\\pm 10^\\circ$, $\\pm 20^\\circ$, and $\\pm 30^\\circ$ at Mach numbers ranging from 0.20 to 0.93. Also presented in figure 12 are elevon-hinge-moment coefficients for the left elevon only (the deflection of which was positive) over the same range of elevon deflections and Mach numbers. These data indicate that the effectiveness of the elevons in producing rolling moment was maintained throughout the test range of angle of attack and Mach number. The effectiveness was nearly constant at angles of attack between $-1^\\circ$ and $+8^\\circ$ for a Mach number of 0.20, and between $-1^\\circ$ and $+6^\\circ$ for the higher Mach numbers. The angles of attack at which the rolling-moment effectiveness of the elevons began to decrease rapidly coincide with those at which the rearward movement of the aerodynamic center is noted in the pitching-moment data. The variation of elevon-hinge-moment coefficient with angle of attack remained fairly uniform over the same angle-of-attack range for which the maximum rolling-moment effectiveness was maintained. At angles of attack just beyond these ranges the variation of hinge-moment coefficient with angle of attack became considerably greater, and at the larger positive angles of attack became erratic.\n\nThe variation of rolling-moment coefficient with total elevon deflection (the arithmetic sum of the positive and negative deflections) was smooth to the largest deflection, as may be seen in figure 13. Increasing the Mach number from 0.20 to 0.93 reduced the effectiveness by roughly 10 percent for an angle of attack of $6^\\circ$ and by about 25 percent for an angle of attack of $10^\\circ$ at the largest elevon deflection $\\delta_H = \\pm 30^\\circ$. The effect of Mach number on the rolling-moment effectiveness of the elevons is summarized in figure 14 for angles of attack of $0^\\circ$ and $4^\\circ$. The rolling moment produced by a given elevon deflection was generally reduced slightly with increasing Mach number, the effect becoming greater with increasing deflections.\n\nHelix angle.— On the basis of the methods presented in reference 10, helix angles generated by the wing tip in a steady roll have been calculated utilizing the data of figure 12. For the purposes of the calculations no torsional deflection and $0^\\circ$ of sideslip were assumed. Values of the damping-moment coefficient $C_{l_p}$, calculated by the method of reference 11, varied from $-0.226$ at a Mach number of 0.20 to $-0.231$ at a Mach number of 0.93.\n\nThe variation of the predicted wing-tip helix angle with total elevon deflection $\\delta_H$ is presented in figure 15 for various Mach numbers at a lift coefficient of 0.20. As anticipated from the decrease in rolling effectiveness above an angle of attack of $8^\\circ$, calculations of $pb/2V$ at a lift coefficient of 0.40 indicated a considerable decrease from its value at a lift coefficient of 0.20. No such calculations are presented herein, however, since above a Mach number of 0.20 the test angle-of-attack range was insufficient to evaluate corrections to the rolling-moment coefficient in roll. The variation of $pb/2V$ with $\\delta_H$ was fairly linear throughout the range of elevon deflections considered.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:22.337003+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 89, "total_pages": 98, "image_filename": "19930086073_p89.jpg", "text": "```markdown\nNACA RM A9E04\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\nLift coefficient, $C_L$\n\n$\\beta$, deg\n$\\circ$ 0.0\n$\\square$ 12.0\n\n-.08 -.06 -.04 -.02 0 .02\nRolling-moment coefficient, $C_l$\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n\n-.06 -.04 -.02 0 .02\nYawing-moment coefficient, $C_n$\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n\n-.12 -.08 -.04 0 .04 .08\nSide-force coefficient, $C_Y$\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 19. – Concluded.\n\nNACA\n\n87\n```", "timestamp": "2026-07-22T06:52:22.667171+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 3, "total_pages": 24, "image_filename": "19930085991_p3.jpg", "text": "NACA RM L9I28\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAN EMPIRICAL CRITERION FOR FIN STABILIZING  \nJETTISSONABLE NOSE SECTIONS OF AIRPLANES  \nBy Stanley H. Scher\n\nSUMMARY\n\nInvestigations in the Langley 20-foot free-spinning tunnel of models  \nof five jettisonable nose sections have shown that airplane nose sections  \nare inherently unstable but can be stabilized by the addition of suitable  \nfins. An empirical criterion has been developed which indicates the  \nfin area required for stabilizing an airplane jettisonable nose  \nsection.\n\nINTRODUCTION\n\nA proposed method of providing for emergency pilot escape from high-  \nspeed airplanes consists of jettisoning the nose section of the fuselage  \nclear of the remainder of the airplane, with the break-off station just  \nrearward of the pilot's station; the pilot leaves the nose section after  \nit has decelerated to a safe speed. Recently, the low-speed behaviors of  \nfive models of possible jettisonable nose configurations for single-seat  \ntransonic airplanes have been investigated in the Langley 20-foot free-  \nspinning tunnel, and it has been noted that each model descended in the  \nvertically rising air stream with some type of rotary motion (refer-  \nence 1 and unpublished data). More recent results (data unpublished)  \nhave indicated that the rotary motion of a jettisoned unstable nose at  \nhigh speeds may not necessarily be similar to that indicated at low  \nspeed, but that even if the nose does not rotate it will tend to trim  \naway from a nose-first flight attitude which may cause decelerations  \ndangerous to the pilot. Analysis indicates that if a nose jettisoned at  \ntransonic speeds could be made to continue flying in a nose-first  \nattitude, the deceleration would not be excessive and, in addition, the  \ndeceleration would act on the pilot's body in the direction (transverse)  \nin which human tolerance to acceleration is highest.", "timestamp": "2026-07-22T06:52:24.138666+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 25, "total_pages": 47, "image_filename": "19930085919_p25.jpg", "text": "24\n\nUpper surface of extended-nose flap\ntangent to upper surface of airfoil\nat each angle setting.\n\nChord line\n\n0.15c\n\n1/8\" plate\n\nSection parallel to air stream\n\n3/8\" diameter\nconstant across\nspan\n\nδn\n\nChord line\n\nAdjustable clamp bracket\n\nSection perpendicular to wing leading edge\n\nExtended-nose flap\n\nCONFIDENTIAL\n\nCONFIDENTIAL\n\nChord line\n\n0.04c\n\nδd\n\n0.004c\n\nPoint of rotation corresponds\nto 15-percent-chord line\n\nSection perpendicular to the 15-percent chord line\n\nDrooped-nose flap\n\n1/8\"\n\n1/16\"\n\nChord line\n\nSharp leading edge\n\nNACA\n\nSection parallel to air stream\n\nSharp leading edge\n\nFigure 5.—The geometry of the extended-nose flap, the drooped-nose flap, and the sharp leading edge.\n\nNACA RM No. 49C21", "timestamp": "2026-07-22T06:52:24.321536+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 61, "total_pages": 72, "image_filename": "19930085491_p61.jpg", "text": "60\nCONFIDENTIAL\nNACA RM No. A8J04\n\n<!-- Image (111, 109, 875, 869) -->\n\nFigure 11.- Variations of parameters affecting drag due to lift for WF-63.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:24.582446+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 9, "total_pages": 25, "image_filename": "19930085979_p9.jpg", "text": "8\nNACA RM E53E12\n\nof plenum-chamber-gas pressures higher than the optimum resulted in\nhigher temperatures near the accessory housing with lower air tem-\nperatures near the outer duct wall. The use of plenum-chamber-gas\npressures lower than optimum resulted in low air temperatures near\nthe accessory-housing skin with increased model-air temperatures\nnear the duct wall.\n\nThe effect of angle of attack on the temperature distribution\ninside the model was negligible.\n\nIncreasing the plenum-chamber-gas temperature at a fixed\nangle of attack had very little effect on the temperature distri-\nbution at the inlet screen aside from increasing the temperature\nlevel.\n\n**Mass-flow loss.** - A reduction in mass flow through the model\noccurred with increasing model-air temperature at fixed tunnel-air\nvelocity, tunnel-air temperature, and angle of attack. The\ndecrease in mass flow with increasing model-air temperature for\nseveral values of tunnel-air velocity and an angle of attack of 0°\nis shown in figure 8. The decrease in mass flow is linearly related\nto the model-air-temperature rise, but does not vary directly with\ndecreasing model-air density associated with increasing temperature,\nas shown in figure 8. Thus, the model did not operate as a constant-\nvolume machine as did the models previously investigated (refer-\nences 1 and 2); instead the experimental mass-flow curve falls\nbetween the constant-mass-flow and constant-volume curves.\n\n**Ram-pressure recovery.** - The ram-pressure loss associated\nwith the addition of heat by jets directed perpendicularly to a\nmoving air stream consists of two components. The first component\narises from the momentum pressure loss associated with changing\nthe direction of the jets. The other component arises from the\nchange in density of the air due to the addition of heat. The\neffect of the first component is illustrated in figure 4.\n\nThe variation of ram-pressure recovery with model-air-\ntemperature rise is shown in figure 9. The ram-pressure recovery\ndecreased linearly with increasing model-air-temperature rise. A\ncomparison of figures 4 and 9 with data from references 1 and 2\nshowed that the ram-pressure recoveries observed with heat addition\nduring this investigation were considerably higher than those\nobtained in previous investigations and were caused by the model\nhaving operating characteristics between those of a constant-mass-\nflow and a constant-volume machine. As a result, the ram-pressure\nrecoveries with heat addition approached those obtained with cold-\ngas bleedback.", "timestamp": "2026-07-22T06:52:26.953491+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 4, "total_pages": 40, "image_filename": "19930085997_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM A9I29\n\nthe scoop side walls immediately behind the inlet and contiguous to the model forebody. From these results, it appeared reasonable to assume that boundary-layer control by means of slots could be replaced by suction scoops and that equally high pressure recovery would result. The latter method of removing the boundary layer possibly would have advantages of arrangement, reduced external drag, and reduced effects of boundary-layer shock-wave interaction ahead of the inlets.\n\nIt is the purpose of the present report to describe tests of a specific configuration which employed boundary-layer control by means of suction scoops and to investigate modifications in the model designed to improve the maximum pressure recovery. No study was made of the external drag contributed by the inlets. In selecting an optimum inlet design, this important variable would have to be considered further.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| A | area, square inches |\n| H | total pressure, pounds per square foot |\n| M | Mach number |\n| m | rate of mass flow, pounds per second |\n| R | Reynolds number |\n| $\\alpha$ | angle of attack, degrees |\n| $\\delta$ | boundary-layer thickness, inches |\n| $\\overline{H}_2/H_0$ | ratio of the average total pressure at position 2 to the free-stream total pressure $\\left[ \\overline{H}_2/H_0 = \\frac{1}{A} \\sum_{n=1}^{n=7} (H_2/H_0)_n \\Delta A_n \\right]$, where n refers to area divisions and tube locations |\n| $(\\overline{H}_2/H_0)_e$ | equivalent total-pressure ratio at position 2, the difference between $\\overline{H}_2/H_0$ and the energy expended in boundary-layer removal |\n| $m_1/m_0$ | ratio of the mass flow entering the main scoops to that which would flow through a tube of the same inlet area in the free stream |\n| $m_4/m_0$ | ratio of the mass flow entering the boundary-layer scoops to that which would flow through a tube of the same inlet area in the free stream |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:27.951394+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 13, "total_pages": 55, "image_filename": "19930085966_p13.jpg", "text": "12 CONFIDENTIAL NACA RM L9B17\n\nFor a given altitude, flight Mach number, burner efficiency, and temperature ratio $T_t$, it is possible to determine from figures 15 and 16 the air flow and fuel flow required to obtain a given thrust condition chosen from figure 12. Figure 15 is a plot of the air-flow parameter against simulated flight Mach number, for constant values of temperature ratio $T_t$. The parameter includes the quantity $\\theta_{t5}$, which can be determined from the temperature ratio $T_t$, the flight Mach number, and the altitude. The maximum deviation of the data from the curves was less than 1 percent on the basis of curves of constant temperature ratio. The air-flow parameter $\\frac{W_a}{\\delta_0} \\sqrt{\\theta_{t5}}$ for a ram-jet unit with a hypothetical combustion chamber of zero pressure losses, is not a function of the temperature ratio (see reference 1), but for an actual unit with appreciable combustion-chamber gas velocities the momentum pressure losses increase with increasing values of temperature ratio, and it is necessary to take into account variations in temperature ratio in correlating the air-flow requirements.\n\nFigure 16 is a plot of the fuel-flow parameter against simulated flight Mach number. This parameter includes the burner efficiency, which must either be determined or assumed in order to make use of the curve. The parameter differs from the one of reference 1 in that the temperature ratio $T_t$ is included to correct for the combustion-chamber losses. With the inclusion of this term the maximum scatter of the data was 12 percent.\n\nThe over-all efficiency was calculated from\n\n$$\n\\eta = \\frac{FV_0}{W_f h_c J}\n$$\n\nThis is the ratio of the thrust power to the combustion energy in the fuel and can be shown to be the product of the burner, cycle, and propulsive efficiencies. The ratio of the heat received by the air to the heat of combustion of the gas is the combustion efficiency,\n\n$$\n\\eta_b = \\frac{(W_a + W_f) c_{p3-5} (T_{t5} - T_{t3})}{W_f h_c}\n$$\n\nThe ratio of the over-all efficiency to the burner efficiency is equal to the product of the cycle and propulsive efficiencies and can be considered as the efficiency with which the heat received by the air is converted to thrust power.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:28.614868+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 21, "total_pages": 34, "image_filename": "19930085913_p21.jpg", "text": "```markdown\n20\nNACA RM L9F24\n\n| 1 | 2 | 3 | 4 |\n| :--- | :--- | :--- | :--- |\n| 30 | 50 | 20 | 20 |\n| | | | |\n| 14 | | | |\n\n| 1 | 2 | 3 | 4 |\n| :--- | :--- | :--- | :--- |\n| 30 | 50 | 20 | 20 |\n| | | | |\n| 15 | | | |\n\n| 1 | 2 | 3 | 4 |\n| :--- | :--- | :--- | :--- |\n| 30 | 50 | 20 | 20 |\n| | | | |\n| 16 | | | |\n\n| 50 | 50 | 20 | 20 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 17 | | | |\n\n| 50 | 50 | 20 | 20 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 18 | | | |\n\n| 50 | 50 | 20 | 20 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 19 | | | |\n\n| 50 | 50 | 20 | 20 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 20 | | | |\n\n| 50 | 50 | 15 | 15 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 21 | | | |\n\n| 50 | 50 | 15 | 15 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 22 | | | |\n\n| 50 | 50 | 15 | 15 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 23 | | | |\n\n| 50 | 50 | 15 | 15 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 24 | | | |\n\n| 50 | 50 | 15 | 15 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 25 | | | |\n\n| 50 | 50 | 15 | 15 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 26 | | | |\n\n| 20 | 20 | 15 | 15 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 27 | | | |\n\n| 20 | 20 | 15 | 15 |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 28 | | | |\n\nNACA\n\n(b) Model A; $\\Lambda = 0^\\circ$; $e_w = 0$.\nFigure 1.— Continued.\n```", "timestamp": "2026-07-22T06:52:30.069233+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 8, "total_pages": 132, "image_filename": "19930085990_p8.jpg", "text": "6\nCONFIDENTIAL\nNACA RM A9I01\n\nCORRECTIONS TO DATA\n\nThe data have been corrected for the effects of tunnel-wall interference, of constriction due to the tunnel walls, and of model-support tare forces. The method of reference 7 was used in computing the corrections to the data for tunnel-wall interference. The following corrections were added:\n\n$$\n\\begin{aligned}\n\\Delta\\alpha &= 0.363 \\text{ C}_L \\\\\n\\Delta\\text{C}_D &= 0.0056 \\text{ C}_L^2 \\\\\n\\Delta\\text{C}_m &= 0\n\\end{aligned}\n$$\n\nCorrections to the data for the constriction effects of the tunnel walls have been evaluated by the method of reference 8. The magnitudes of these corrections as applied to Mach number and to dynamic pressure (measured with the tunnel empty) are illustrated by the following table:\n\n| Corrected Mach number | Uncorrected Mach number | | $\\frac{q_{corrected}}{q_{uncorrected}}$ | |\n| :--- | :--- | :--- | :--- | :--- |\n| | Wing alone | Wing and fuselage | Wing alone | Wing and fuselage |\n| 0.95 | 0.937 | 0.917 | 1.005 | 1.036 |\n| .92 | .915 | .896 | 1.003 | 1.027 |\n| .90 | .897 | .881 | 1.002 | 1.023 |\n| .85 | .848 | .838 | 1.002 | 1.016 |\n| .80 | .799 | .792 | 1.001 | 1.012 |\n| .70 | .700 | .696 | 1.001 | 1.008 |\n| .50 | .500 | .499 | 1.001 | 1.005 |\n| .20 | .200 | .200 | 1.001 | 1.005 |\n\nThe theoretical choking Mach number for the wing-fuselage combination was 0.96.\n\nTare corrections due to the air forces exerted on the turntable were obtained from force measurements made with the model removed from the tunnel. Possible interference effects between the model and the turntable were not evaluated. The magnitude of the measured tare-drag coefficient, based on the wing area, was independent of Mach number and varied with Reynolds number as follows:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:30.451219+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 10, "total_pages": 30, "image_filename": "19930085975_p10.jpg", "text": "8\nCONFIDENTIAL\nNACA RM L9E10\n\nIt will be noted that for the range tested $C_{l_p}$ increases appreciably with angle of attack, particularly at the higher Mach numbers. A similar effect was noted in reference 4. The linearized theory, however, does not predict any variation of $C_{l_p}$ with angle of attack because it is evaluated in terms of lift rather than resultant force and does not consider any nonlinear variation of lift with angle of attack. The section data for these wings are not available but a value of $2\\pi$ was assumed for the lift-curve slope in the theoretical calculations. A study of the effect of nonlinear lift characteristics has been made in reference 7 and has indicated that this effect alone can cause large changes in $C_{l_p}$.\n\nCONCLUSIONS\n\nOn the basis of an investigation of the damping characteristics of three wings of aspect ratio 4 and taper ratio 0.6 having quarter-chord line sweep angles of $3.6^\\circ$, $32.6^\\circ$, and $46.7^\\circ$ in the Mach number range from 0.40 to 0.91, the following conclusions can be drawn:\n\n1. The damping-in-roll coefficient $C_{l_p}$ increased in magnitude with Mach number and decreased with sweep angle at low angles of attack ($0.30^\\circ$ and $3.45^\\circ$) in the same manner as that predicted by theory.\n\n2. The magnitude of the damping-in-roll coefficient $C_{l_p}$ increased markedly with angle of attack (in the test range from $0.30^\\circ$ to $6.5^\\circ$) particularly at the higher Mach numbers.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:31.480912+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 34, "total_pages": 52, "image_filename": "19930085911_p34.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 33\n\nNet acceleration, $a_n$, g's\nFree-stream Mach number, $M_0$\n\nTime after release, $\\tau$, sec\n\n(a) Resultant flight conditions.\n\nFigure 8. - Time history of flight data and performance of ram-jet unit 16-A-3.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:31.668522+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 46, "total_pages": 92, "image_filename": "19930085870_p46.jpg", "text": "NACA RM No. L9D07\n47\n\nCONFIDENTIAL\n\nElliptical L.E. $\\begin{cases} \\bigcirc C_L \\\\ \\square C_m \\end{cases}$\nWedge L.E. $\\begin{cases} \\triangle C_L \\\\ \\diamond C_m \\end{cases}$\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. $\\begin{cases} \\bigcirc C_D \\\\ \\square L/D \\end{cases}$\nWedge L.E. $\\begin{cases} \\triangle C_D \\\\ \\diamond L/D \\end{cases}$\n\n$C_D$\n.06\n.04\n.02\n\n$L/D$\n6\n4\n2\n0\n\n$\\alpha$, deg\n-8\n-6\n-4\n-2\n0\n2\n4\n6\n8\n\n[NACA logo]\n\n(f) Wing 6. w=0.940; R = 900,000.\nFigure 6.- Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:32.365635+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 92, "total_pages": 114, "image_filename": "19930086061_p92.jpg", "text": "88\nNACA RM L9J07\n\n$$\n\\frac{c_l c}{C_L C_{av}}\n$$\nTheoretical load distribution\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 4.1 | 0.15 |\n| 8.1 | 0.33 |\n\n$$\n\\frac{y}{b/2} \\text{ , percent}\n$$\nLeft Right\n\n(a) Angles of attack: 4.1°, 8.1°\n\n$$\n\\frac{c_l c}{C_L C_{av}}\n$$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 14.1 | 0.60 |\n| 24.1 | 0.82 |\n| 34.1 | 0.98 |\n\n$$\n\\frac{y}{b/2} \\text{ , percent}\n$$\nLeft Right\n\n(b) Angles of attack: 14.1°, 24.1°, 34.1°\n\n$$\n\\frac{c_l c}{C_L C_{av}}\n$$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 39.1 | 0.91 |\n| 44.1 | 0.67 |\n\n$$\n\\frac{y}{b/2} \\text{ , percent}\n$$\nLeft Right\nNACA\n\n(c) Angles of attack: 39.1°, 44.1°\n\nFigure 37.- Span load distribution of wing 1 at various angles of attack; $\\psi = 0^\\circ$. All data are taken over left semispan.", "timestamp": "2026-07-22T06:52:32.519630+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 55, "total_pages": 149, "image_filename": "19930083192_p55.jpg", "text": "```markdown\nNACA TN 1976\n51\n\n# OPERATING STATISTICS\n\nA knowledge of gust structure and methods for computing airplane reactions is insufficient for gust-load calculations without knowing the conditions for which load calculations are required. Thus, in order to solve the practical problem of load prediction, the gusts that airplanes encounter and the operating conditions (speed, weight, altitude, and, perhaps, center-of-gravity position) that are likely to exist at the time the gust is encountered must be known. The fact that the source of load, the atmospheric gust, is of a fairly random character and that the operating conditions vary widely make it impractical to predict the exact loads and the associated conditions that are experienced by a given airplane during its lifetime.\n\n## METHOD\n\nThe general method of obtaining the desired information has been to install instruments in aircraft and to record their experiences and operating conditions without interfering with routine practices. Three procedures have been used: the installation of simple instruments such as the NACA V-G recorder giving broad coverage as to route, airplane type, and time but no detailed information, the installation of special instrumentation for a very limited period of time to obtain information on a particular quantity in detail, and, finally, the installation of fairly elaborate instrumentation controlled by an observer to obtain as much detailed data as possible on several quantities during one or more flights.\n\nThe NACA V-G recorder (reference 4) records on a smoked glass plate the normal acceleration as a function of the airspeed at which it was imposed. In use, a record plate is left in the instrument during operations and the resulting clear area on the glass represents an envelope of accelerations and airspeeds experienced by the airplane.\n\nStandard NACA photographically recording instruments or motion pictures of indicating instruments are used for more detailed measurements.\n\n## SCOPE OF DATA\n\nThe characteristics and pertinent dimensions of transport airplanes for which gust statistics have been collected are given in table XVIII. The listed values of the limit load-factor increment due to gusts, the high speed in level flight, and the placard do-not-exceed speed were obtained\n```", "timestamp": "2026-07-22T06:52:32.943169+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 13, "total_pages": 30, "image_filename": "19930085970_p13.jpg", "text": "Equation for body ordinates:\n$$\\frac{r}{6} = \\left[1 - \\left(1 - \\frac{2x}{l}\\right)^2\\right]^{\\frac{3}{4}}$$\n\nAspect ratio, $A$, = 3.42\nTaper ratio, $\\lambda$, = .258\n\nFineness ratio: $\\frac{l}{2r_6} = 12.5$\n\n[Figure: Sketch of model illustrating principal dimensions of the wing and body. The figure includes a side view of a fuselage and a top view of a wing attached to it. Various dimensions are labeled with arrows and values. Angles and ratios are also indicated.]\n\nDimensions labeled on the figure:\n- $x$\n- $63^\\circ$\n- $2.350$\n- $1.925$\n- $2r_6$\n- $.680$\n- $4.250$\n- $\\frac{l}{2}$\n- $6.715$\n- m.a.c.\n- $1.648$\n- $1.016$\n- $5.050$\n- $.605$\n\nAll dimensions are in inches\n\nStreamwise airfoil section: NACA 64A-006\n\nNACA [Logo]\n\nFigure 1.- Sketch of model illustrating principal dimensions of the wing and body.\n\nCONFIDENTIAL\n\nNACA RM A9E09\n\nCONFIDENTIAL\n\n11", "timestamp": "2026-07-22T06:52:35.493213+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 4, "total_pages": 32, "image_filename": "19930085992_p4.jpg", "text": "2\nNACA RM L9EL7\n\nalmost empty condition at the time of failure) suggested the possi-\nbility of wing flutter, with a lower flutter speed resulting from the\naerodynamic forces acting on the tank. No analytical treatment is\navailable for predicting the oscillatory forces on such bodies, and\nthus the effect on flutter characteristics of a change in body shape\ncannot be directly calculated. A systematic experimental study of\neffects of concentrated weights on flutter characteristics was\nreported in reference 1, and analytical studies of these effects were\nmade in references 2 and 3. Throughout the studies, however, no\nparticular attention was given to the aerodynamic contours of the con-\ncentrated weights employed. The primary objective of this paper is to\npresent experimental results on some effects on the flutter speed and\nflutter frequency of a wing carrying concentrated weights having\nwidely different aerodynamic shapes.\n\nThis paper presents results of a limited study that is part of a\nbroader investigation. For the flutter tests, a straight untapered\nuniform cantilever wing was used. The concentrated weights employed\nhad similar mass and moment-of-inertia properties, but were of\ndifferent aerodynamic shapes. One of these weights was a streamlined\nbody resembling in shape an external fuel tank, whereas the other was\nof a nonstreamlined shape.\n\nThe concentrated weights were selected so that the ratio of their\nweights to that of the wing was comparable to the ratio of the weights\nof an empty external fuel tank and the wing of a typical airplane. In\nview of the relatively low ratio of the weight of an empty external\nfuel tank to the weight of a wing, it seemed unlikely that its mass\nmight exert much influence on the flutter characteristics of the wing.\nBecause of the shape of the tank, however, the moment of inertia of\nthe tank is usually high in comparison with its weight. Since this\nappreciable moment of inertia may quite conceivably exert strong\ninfluence on flutter characteristics, a study of the effect on these\ncharacteristics of a variation in the moment of inertia of the external\ntank has been included.\n\nFlutter tests were conducted with the weights mounted rigidly at\nthe wing tip and at various spanwise positions. In the present pre-\nliminary study, the testing was done only at low Mach numbers and only\na few wing-weight configurations were used.", "timestamp": "2026-07-22T06:52:36.192167+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 26, "total_pages": 47, "image_filename": "19930085919_p26.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 25\n\n[Figure: (a) Full-span drooped-nose flap. NACA A-12382]\n\n[Figure: (b) Full-span extended-nose flap. NACA A-12384]\n\nFigure 6.— The model with leading-edge flaps mounted in one of the Ames 7- by 10-foot wind tunnels.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:39.360750+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 5, "total_pages": 17, "image_filename": "19930085988_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM L9H30\n\nThe error in the results is believed to be within the following limits:\n\n| Quantity | Error | |\n| :--- | :--- | :--- |\n| | M = 1.0 | M = 1.5 |\n| $C_D$ (referred to wing-plan-form area): | | |\n| Total . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 57, "total_pages": 60, "image_filename": "19930085862_p57.jpg", "text": "NACA RM No. L9A07\n\n$$\n\\begin{array}{c}\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } \\\\\n\\text{ } 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{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 62, "total_pages": 72, "image_filename": "19930085491_p62.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 61\n\n(a) WF-63, $C_L = 0$\n\n(b) WF-57, $C_L = 0.27$\n\nLaminar flow area\nTurbulent flow area\nSeparated flow area\nFuselage area\nDirection of flow\n\n[Figure: Typical liquid-film test results.]\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:41.831591+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 11, "total_pages": 46, "image_filename": "19930085983_p11.jpg", "text": "NACA RM A9I27 CONFIDENTIAL 9\n\nIncreasing Mach number generally reduced the helix angle. While the predicted wing-tip helix angle is large enough to insure high rolling velocities, it must be emphasized that the present calculations are for a rigid wing and that deflection of the wing could cause serious reductions in the magnitude of the rolling velocity.\n\nLongitudinal Control of a Hypothetical Airplane\n\nData from the tests have been used in the calculation of the stability, maneuverability, elevon hinge moments, and power-off sinking speed of a hypothetical tailless airplane, geometrically similar to the model tested. Dimensions of the airplane were assumed to be as follows:\n\n| Wing span, feet | 50 |\n| --- | --- |\n| Wing area, square feet | 714.3 |\n| Total elevon area, square feet | 89.14 |\n\nThe center of gravity was assumed to be at 25 percent of the mean aerodynamic chord, and a wing loading of 40 pounds per square foot was assumed.\n\nFigure 16 presents elevon hinge moment, elevon deflection, and lift coefficient as functions of Mach number calculated for the airplane in level flight and as affected by normal acceleration at an altitude of 25,000 feet. The variation of elevon deflection with Mach number and with normal acceleration factor was smooth and uniform. A very large variation of hinge moment with Mach number is noted for normal acceleration factors greater than 1.0. For unaccelerated flight ($n = 1.0$) increasing Mach number would require a gradually increasing push force up to a Mach number of 0.90. For a normal acceleration factor of 2.0, increasing Mach number is accompanied by a gradually decreasing push force. For constant-speed maneuvers with varying normal acceleration there are large and erratic changes in the hinge moment.\n\nPower-off sinking speed, elevon deflection for balance, elevon hinge moment, and angle of attack are presented in figure 17 as functions of power-off gliding speed for sea-level operation. (Data at a Mach number of 0.20 were used in calculating the performance parameters shown in this figure.) The minimum power-off sinking speed is 22 feet per second and it occurs at a forward speed of approximately 215 miles per hour. The variation of elevon deflection required for longitudinal balance with gliding speed was stable for gliding speeds greater than 180 miles per hour. No computations are shown for gliding speeds less than 180 miles per hour, since the data indicated that the airplane would be longitudinally unstable at the required lift coefficients.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:43.337501+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 90, "total_pages": 98, "image_filename": "19930086073_p90.jpg", "text": "```markdown\n88\n\n.4\n.2\n0\n-.2\n-30 -20 -10 0 10 20 30 40 50\nFlap deflection, $\\delta_f$, deg\nIncrement of lift coefficient, $\\Delta C_L$\n\n0.0° Sideslip\nAngle of attack, $\\alpha$, deg\n0\n8\n19\n24\n\n12.1° Sideslip\n-30 -20 -10 0 10 20 30 40 50\nFlap deflection, $\\delta_f$, deg\n(a) Wing alone.\n\n.4\n.2\n0\n-.2\n-30 -20 -10 0 10 20 30 40 50\nFlap deflection, $\\delta_f$, deg\nIncrement of lift coefficient, $\\Delta C_L$\n\n0.0° Sideslip\n\n12.0° Sideslip\n-30 -20 -10 0 10 20 30 40 50\nFlap deflection, $\\delta_f$, deg\n(b) Wing + body.\n\n[Figure: NACA logo]\n\nFigure 20.— Increments of lift coefficient for various flap deflections at four angles of attack.\n\nNACA RM A59H04\n```", "timestamp": "2026-07-22T06:52:43.585382+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 4, "total_pages": 24, "image_filename": "19930085991_p4.jpg", "text": "2\nNACA RM L9I28\n\nIn the present investigation, the testing technique and test results for the five models at low speed are briefly reviewed, and an empirical criterion based on consideration of these results has been prepared which relates the effects of fin design and of center-of-gravity location on the stability of airplane jettisonable nose sections. Curved fins (simulating fins which would normally be folded flush with the fuselage and extended during an emergency requiring jettisoning of the nose) were used in some of the tests. Recent NACA work on some of the high-speed aspects of fin stabilization of a jettisonable nose is also discussed. The problem of providing for clean separation between nose and airplane is beyond the scope of the present paper.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| $X, Y, Z$ | longitudinal, lateral, and normal axes, respectively, through center of gravity of nose |\n| $k_x, k_y, k_z$ | radii of gyration of nose about X-, Y-, and Z-axes, respectively, inches |\n| $n$ | fineness ratio of nose, excluding canopy or other protuberance (for circular cross section, Length/Diameter; for non-circular cross section, Length/Maximum cross dimension) |\n| $L$ | length of nose section, feet (All center-of-gravity locations are expressed as a percentage of this length from the front end of the nose section.) |\n| $S_F$ | smallest projected fin area in any plane parallel to longitudinal axis |\n| $S_p$ | projected area of nose (excluding protuberances) in plane of smallest projected fin area |\n| $L_F$ | projected distance between centroid of $S_F$ and center of gravity of nose |\n| $\\frac{S_F L_F}{S_p L}$ | fin-stabilization factor (fig. 1) |\n| $\\alpha$ | angle of attack of nose X-axis, degrees |\n| $C_m$ | pitching-moment coefficient as determined graphically |", "timestamp": "2026-07-22T06:52:52.186290+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 35, "total_pages": 52, "image_filename": "19930085911_p35.jpg", "text": "34\nCONFIDENTIAL\nNACA RM E9F22\n\n<!-- Image (133, 113, 887, 846) -->\n\n(b) Independent test variables.\nFigure 8. - Continued. Time history of flight data and performance of\nram-jet 16-A-3.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:54.849219+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 5, "total_pages": 40, "image_filename": "19930085997_p5.jpg", "text": "NACA RM A9I29 CONFIDENTIAL 3\n\nThe following subscripts indicate the position at which the quantities were measured (fig. 1):\n\n0 free stream  \n1 entrance to main scoop  \n2 survey position immediately downstream of main-scoop entrance  \n3 settling chamber  \n4 entrance to boundary-layer scoops  \n5 survey station in boundary-layer removal duct  \n\nAPPARATUS AND MODELS\n\nThe tests were performed in the Ames 8- by 8-inch supersonic wind tunnel at Mach numbers between 1.36 and 2.01 and Reynolds numbers, based upon the length of the model ahead of the inlets (3.934 in.), of 2.6 to 3.4 million. Flow through the boundary-layer removal ducts was exhausted to the atmosphere through a vacuum pump. A description of the equipment and wind tunnel can be found in references 2 and 3.\n\nThe model (see figs. 1 and 2) was built to simulate the forward portion of the fuselage and the ducts of a possible supersonic airplane designed to fly in the speed range up to a Mach number of 2.0. In designing the scoops two variables were compromised: First, to supply efficiently the air consumed by engines capable of driving the airplane at these speeds, large inlet areas would be required below a Mach number of about 0.5, and small areas would be required in the supersonic range; and, second, large leading-edge radii would be desirable to prevent the flow from stalling on the inside surface of the lips at subsonic speeds, and sharp leading edges would be desirable to decrease the wave drag at supersonic speeds. These situations were compromised by choosing the inlet area large enough so that a normal shock wave would form ahead of the inlet at all supersonic speeds. Thus, by choosing the inlet areas sufficiently large, auxiliary inlets would be unnecessary in the subsonic range. Furthermore, large leading-edge radii would be permissible since the flow behind the normal shock wave always would be subsonic. At supersonic speeds, the large inlet area would result in spilling of the air around the inlets at the expense, of course, of increased external drag.\n\nConical subsonic diffusers commonly used at low speeds have a severe adverse pressure gradient near their entrance when operated at high inlet Mach numbers. It was assumed that decreasing this gradient would reduce the tendency toward separation of the boundary layer; hence the internal shape of the main ducts was designed to have a constant static-pressure gradient from the inlet to a station approximately 20 percent of the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:56.049739+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 47, "total_pages": 92, "image_filename": "19930085870_p47.jpg", "text": "48\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {○ $C_L$\n                {□ $C_m$\nWedge L.E.     {△ $C_L$\n                {◇ $C_m$\n\n.16\n\n.08\n$C_L$\n0\n\n-.08\n\n-.16\n\n-.24\n\n.01\n$C_m$\n0\n\n-.01\n\n.06\nElliptical L.E. {○ $C_D$\n                {□ $L/D$\nWedge L.E.     {△ $C_D$\n                {◇ $L/D$\n\n.04\n$C_D$\n.02\n\n0.8\n\n-6   -4   -2    0    2    4    6    8\n$\\alpha$, deg\n\n6\n4\n$L/D$\n2\n0\n\n[NACA logo]\n\n(g) Wing 7.w=1.030; R = 840,000.\nFigure 6.-Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:59.766473+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 11, "total_pages": 30, "image_filename": "19930085975_p11.jpg", "text": "NACA RM L9E10 CONFIDENTIAL 9\n\nREFERENCES\n\n1. Toll, Thomas A., and Queijo, M. J.: Approximate Relations and Charts for Low-Speed Stability Derivatives of Swept Wings. NACA TN 1581, 1948.\n\n2. Bird, John D.: Some Theoretical Low-Speed Span Loading Characteristics of Swept Wings in Roll and Sideslip. NACA TN 1839, 1949.\n\n3. Fisher, Lewis R.: Approximate Corrections for the Effects of Compressibility on the Subsonic Stability Derivatives of Swept Wings. NACA TN 1854, 1949.\n\n4. Myers, Boyd C., II, and Kuhn, Richard E.: High-Subsonic Damping-in-Roll Characteristics of a Wing with the Quarter-Chord Line Swept Back $35^\\circ$ and with Aspect Ratio 3 and Taper Ratio 0.6. NACA RM L9C23, 1949.\n\n5. 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\n6. Pearson, Henry A., and Jones, Robert T.: Theoretical Stability and Control Characteristics of Wings with Various Amounts of Taper and Twist. NACA Rep. 635, 1938.\n\n7. MacLachlan, Robert, and Letko, William: Correlation of Two Experimental Methods of Determining the Rolling Characteristics of Unswept Wings. NACA TN 1309, 1947.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:59.854618+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 14, "total_pages": 30, "image_filename": "19930085970_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:53:00.892261+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 14, "total_pages": 55, "image_filename": "19930085966_p14.jpg", "text": "NACA RM L9B17 CONFIDENTIAL 13\n\nThe cycle efficiency for a hypothetical ram jet with no pressure losses is given in equation (10) of reference 2 as\n\n$$\n\\eta_{tc} = \\frac{1}{1 + \\frac{2}{(\\gamma - 1)M_0^2}}\n$$\n\nThe actual propulsive efficiency is given in equation (20) of reference 2 as\n\n$$\n\\eta_p = \\frac{1}{1 + \\frac{1}{2}\\left(\\frac{\\sqrt{V_7} - V_0}{V_0}\\right)}\n$$\n\nThe last expression for a ram jet with no pressure losses reduces to\n\n$$\n\\eta_p = \\frac{2}{1 + \\sqrt{\\tau_t}}\n$$\n\nThe over-all efficiency for a ram jet with no pressure losses is therefore a function of three variables: burner efficiency, flight Mach number, and temperature ratio.\n\nThe product of the actual cycle and propulsive efficiencies (as obtained by dividing the over-all efficiency by combustion efficiency) is plotted against simulated flight Mach number in figure 17 for several temperature ratios. A value of 3.6 percent was reached with a temperature ratio of 3 at a simulated flight Mach number of 0.545 which corresponded in this case to an over-all efficiency of 2.04 percent.\n\nThe ratio of the actual cycle-efficiency and propulsive-efficiency product to the product of the cycle and propulsive efficiencies for a no-loss system is plotted against the simulated flight Mach number in figure 18. No reliable trends are indicated by the data but the order of magnitude of the ordinate is 0.80. In this comparison the burner efficiency is not a factor and the 20-percent drop below 100 percent must be charged to internal friction, turbulence losses, and momentum-pressure losses occurring in the diffuser and combustion chamber.\n\nIn order to determine the contribution to this loss chargeable to diffuser losses, pressure measurements were made to determine the diffuser efficiency\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:04.352608+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 10, "total_pages": 25, "image_filename": "19930085979_p10.jpg", "text": "NACA RM E52L2\n\nInlet-lip temperature distribution. - The maximum inlet-lip temperatures were obtained at the highest value of bleedback (9.1 percent) and plenum-chamber-gas temperature ($1000^\\circ$ F) utilized in the investigation. The variation of inlet-lip temperature distribution with hot-gas bleedback is shown in figure 10 for a plenum-chamber-gas temperature of $1000^\\circ$ F and a tunnel-air velocity of 205 feet per second. The inlet-lip temperatures decreased with decreasing bleedback, increasing tunnel-air velocity, or decreasing plenum-chamber-gas temperature.\n\nIcing with Bleedback\n\nIn analyzing the icing data, the pressure-drop coefficients $\\Delta p/q$ across the screen were computed for each icing run where $\\Delta p$ is the static-pressure drop across the screen and $q$ is the dynamic pressure upstream of the screen. The screen was considered iced when the value of $\\Delta p/q$ approached 1.5 times the value for the screen at the beginning of each run. The experimental bleed-back and plenum-chamber-gas temperatures corresponding to this criterion are shown in figure 11 for tunnel-air velocities of 200, 275, 355, and 435 feet per second at an angle of attack of $0^\\circ$.\n\nThe theoretical curves shown in figure 11 are based on the assumption that icing occurs when the minimum adiabatic wall temperature or kinetic temperature (static temperature plus 0.85 times the dynamic temperature) on the screen was $32^\\circ$ F. These curves represent the upper and lower limits of the icing conditions used in the investigation; that is, curve A was calculated for saturated air at $0^\\circ$ F, a tunnel-air velocity of 435 feet per second, and a liquid-water content of 1.4 grams per cubic meter. Curve B was calculated for saturated air at $0^\\circ$ F, a tunnel-air velocity of 200 feet per second, and a liquid-water content of 0.7 gram per cubic meter.\n\nIn order to assure a minimum adiabatic wall temperature of $32^\\circ$ F on the screen, an average adiabatic wall temperature of $42^\\circ$ F and a total temperature of $45.9^\\circ$ F are required at a tunnel-air velocity of 435 feet per second, corresponding to a velocity in the screen of 558 feet per second and a total temperature of $43.1^\\circ$ F at a tunnel-air velocity of 200 feet per second, corresponding to a velocity in the screen of 295 feet per second.\n\nThe data in figure 11 fall within the limits of the two curves. The displacement of some of the high-velocity data from curve A and the low-velocity data from curve B was due to the variation in liquid-water content and the use of other than the optimum amount of bleedback.", "timestamp": "2026-07-22T06:53:04.605155+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 24, "total_pages": 92, "image_filename": "19930085951_p24.jpg", "text": "```markdown\n22\nNACA RM L9D29\n\nUNCONFIDENTIAL\n\nDeveloped plan form\n\nLeading edge\n\nNACA-10-(3)(08)-03\nNACA-10-(3)(08)-03R\nNACA-10-(3)(12)-03\n\nb/D\n\n$\\beta$\n\nh/b\n\n$c_{ld}$\n\nSpinner location\n\nBlade thickness ratio, h/b\nBlade width ratio, b/D\nBlade angle, $\\beta$, deg\nBlade-section design lift coefficient, $c_{ld}$\n\nFraction of tip radius, r/R\n\nNACA\n\nFigure 4.- Blade-form curves for NACA propellers having a solidity of 0.03 per blade at the 0.7 radius.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:53:05.862952+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 27, "total_pages": 47, "image_filename": "19930085919_p27.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:53:07.032790+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 22, "total_pages": 34, "image_filename": "19930085913_p22.jpg", "text": "NACA RM L9F24\n21\n\n| | 1 | 2 | 3 | 4 | | 1 | 2 | 3 | 4 | | 1 | 2 | 3 | 4 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | 50 | 50 | 30 | 30 | | 50 | 50 | 30 | 30 | | 50 | 50 | 30 | 30 |\n| 29 | | | | | 30 | | | | | 31 | | | | |\n| | 50 | 50 | 30 | 30 | | 50 | 50 | 30 | 30 | | 50 | 50 | 30 | 30 |\n| 32 | | | | | 33 | | | | | 34 | | | | |\n| | 50 | 50 | 30 | 30 | | 50 | 50 | 30 | 30 | | 50 | 50 | 30 | 30 |\n| 35 | | | | | 36 | | | | | 37 | | | | |\n| | 50 | 50 | 30 | 30 | | 50 | 50 | 30 | 30 | | 50 | 50 | 30 | 30 |\n| 38 | | | | | 39 | | | | | 40 | | | | |\n| | 30 | 50 | 15 | 15 | | 30 | 30 | 30 | 30 | | 30 | 30 | 30 | 30 |\n| 41 | | | | | 42 | | | | | 43 | | | | |\n| | 20 | 20 | 15 | 15 | | 30 | 30 | 20 | 20 | | 30 | 30 | 20 | 20 |\n| 44 | | | | | 45 | | | | | 46 | | | | |\n| | 20 | 20 | 20 | 20 | | 30 | 30 | 20 | 20 | | | | | |\n| 47 | | | | | 48 | | | | | | | | | |\n\n[Figure: NACA logo]\n\n(c) Model B; $\\Lambda = 45^\\circ$, $\\theta_W = -1$.\nFigure 1.— Continued.", "timestamp": "2026-07-22T06:53:07.855947+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 9, "total_pages": 132, "image_filename": "19930085990_p9.jpg", "text": "```markdown\nNACA RM A9I01 CONFIDENTIAL 7\n\n| Reynolds number | $C_{D_{tare}}$ |\n| :--- | :--- |\n| 2,000,000 | 0.0063 |\n| 6,000,000 | .0057 |\n| 10,000,000 | .0056 |\n\nThe rake of total-pressure tubes and static-pressure tubes used to measure the dynamic pressure at the horizontal tail was calibrated throughout the complete range of Mach numbers, of Reynolds numbers, and of angles of attack of the rake.\n\nTESTS\n\nLift, drag, and pitching-moment data have been obtained for the model and its components in the following combinations: (1) the wing alone; (2) the wing and the fuselage; (3) the wing, the fuselage, and the tail mounted in the extended wing-chord plane; (4) the wing, the fuselage, and the supporting bracket for mounting the tail above the fuselage; and (5) the wing, the fuselage, and the tail mounted above the fuselage.\n\nAt a Reynolds number of 2,000,000 the model was tested at Mach numbers from 0.20 to 0.95. The range of angles of attack for these tests was from $-6^\\circ$ to beyond the stall, except at the higher Mach numbers where the range was reduced by the limitations of wind-tunnel power and of model strength. At a Mach number of 0.20 the effect of leading-edge and trailing-edge flap deflection ($\\delta_n = 30^\\circ$ and $\\delta_f = 50^\\circ$) was investigated at Reynolds numbers of 2,000,000, 6,000,000, and 10,000,000. This combination of flap deflections was selected upon the basis of reference 2 wherein it was shown to be the optimum for maximum lift of the wing alone.\n\nTo determine the longitudinal control which would be provided by an all-movable stabilizer, the model was tested with the angle of the stabilizer varied in $2^\\circ$ increments from $-10^\\circ$ to $4^\\circ$ for the model with the tail mounted in the extended wing-chord plane and from $-6^\\circ$ to $4^\\circ$ for the model with the tail mounted above the fuselage.\n\nThe velocity distribution in the wing-fuselage wake was investigated at a position corresponding longitudinally to the midchord of the horizontal tail (3.508 wing mean aerodynamic chord behind the quarter point of the wing mean aerodynamic chord) and corresponding laterally to the location of the mean aerodynamic chord of the tail (0.428 wing mean aerodynamic chord from the plane of symmetry). The extent of the survey was sufficient to permit the determination of the dynamic pressure at either position of the horizontal tail for a range of angle of attack, of Mach number, and of Reynolds number.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:53:08.801059+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 58, "total_pages": 60, "image_filename": "19930085862_p58.jpg", "text": "56\nNACA RM No. L9A07\n\n<!-- Image (128, 109, 809, 844) -->\n\nFigure 23.- Variation of rolling-moment coefficient with span of 0.10c projection step spoilers on wing equipped with extensible leading-edge flaps and trailing-edge split flaps.", "timestamp": "2026-07-22T06:53:10.439063+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 63, "total_pages": 72, "image_filename": "19930085491_p63.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:53:11.446407+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 93, "total_pages": 114, "image_filename": "19930086061_p93.jpg", "text": "NACA RM L9J07\n89\n\n$$\n\\frac{c_l c}{C_L c_{av}}\n$$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\diamond$ 4.1 | 0.16 |\n| $\\square$ 8.1 | 0.32 |\n\n[Graph showing data points for $\\alpha = 4.1^\\circ$ and $\\alpha = 8.1^\\circ$]\n\n(a) Angles of attack: $4.1^\\circ$, $8.1^\\circ$.\n\n$$\n\\frac{c_l c}{C_L c_{av}}\n$$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\diamond$ 14.1 | 0.52 |\n| $\\triangle$ 24.1 | 0.74 |\n| $\\triangledown$ 34.1 | 0.89 |\n\n[Graph showing data points for $\\alpha = 14.1^\\circ$, $\\alpha = 24.1^\\circ$, and $\\alpha = 34.1^\\circ$]\n\n(b) Angles of attack: $14.1^\\circ$, $24.1^\\circ$, $34.1^\\circ$.\n\n$$\n\\frac{c_l c}{C_L c_{av}}\n$$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\square$ 39.1 | 0.83 |\n| $\\diamond$ 44.1 | 0.64 |\n\n[Graph showing data points for $\\alpha = 39.1^\\circ$ and $\\alpha = 44.1^\\circ$]\n[NACA logo]\n\n(a) Angles of attack: $39.1^\\circ$, $44.1^\\circ$.\n\nFigure 38.- Span load distribution of wing 1 at various angles of attack; $\\psi = 10^\\circ$. Flagged symbols represent data taken with left semispan at $\\psi = -10^\\circ$.", "timestamp": "2026-07-22T06:53:16.533486+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 56, "total_pages": 149, "image_filename": "19930083192_p56.jpg", "text": "52\nNACA TN 1976\n\nwhen possible from official sources. For the older airplanes where the\ndesign conditions differed from modern requirements, the pertinent speeds\nand load factors were recomputed according to reference 1 in order to\nplace all the characteristics on a comparable basis.\n\nTable XIX is a summary of the conditions investigated and includes\nthe routes, periods covered, and the amount of data obtained. The opera-\ntions have been subdivided into prewar, wartime, and postwar periods.\nFor the prewar and wartime periods, information was obtained with the V-G\nrecorder, but for the postwar period sufficient V-G records are not\navailable. Special studies of speed and Mach number variations are\navailable and some of the results have been included. The material\ncovered in table XIX does not include all the data obtained since scat-\ntered V-G data of insufficient scope for analysis have been disregarded.\n\nSTATISTICAL METHODS\n\nThe random character of gusts and the lack of control over flight\nconditions have resulted in the use of statistical methods of analysis\nto smooth data, to eliminate improper weighting, to extrapolate results,\nand to place data from different sources on a comparable basis. The\nbasic data in most cases are a count of the values of a quantity according\nto magnitude. Since the application of statistical methods to gust loads\nis fairly recent, the mass of data collected earlier offers some diffi-\nculties because the requirements of statistical analysis were not\nconsidered in their collection.\n\nPearson type III probability distribution curves (see references 41\nand 42) have been utilized in the analysis on the assumption that they\nadequately represent the data on operating statistics. The curve is\ndetermined by three parameters: the mean value, the standard deviation,\nand the coefficient of skewness. In general, the goodness of fit has\nbeen based on engineering judgment rather than any test procedure.\n\nAn important problem in the study of the frequency of exceeding the\nlarger values of load or speed under operating conditions is the deter-\nmination of whether observed or estimated differences in frequencies\nbetween samples are real or represent limitations of the data. No\nsatisfactory test for this purpose has been found, but a 5:1 ratio of\nfrequencies has been used in reference 43 as an engineering measure. If\nthe differences are less than this criterion, real differences may exist\nbut are considered too small to be determined from the available data.\n\nA serious limitation in the analysis of operating statistics exists\nsince the results are used to predict future operating experience. If\nno extraneous factor enters, such as changes in design rules, operating", "timestamp": "2026-07-22T06:53:18.337012+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 91, "total_pages": 98, "image_filename": "19930086073_p91.jpg", "text": ".2\n0.0° Sideslip\nIncrement of pitching-moment coefficient, $\\Delta C_m$\nAngle of attack, $\\alpha$, deg\n0\n8\n19\n24\n-30 -20 -10 0 10 20 30 40 50\nFlap deflection, $\\delta_f$, deg\n.2\n0.0° Sideslip\nIncrement of pitching-moment coefficient, $\\Delta C_m$\n-30 -20 -10 0 10 20 30 40 50\nFlap deflection, $\\delta_f$, deg\n.2\n12.1° Sideslip\nIncrement of pitching-moment coefficient, $\\Delta C_m$\n-30 -20 -10 0 10 20 30 40 50\nFlap deflection, $\\delta_f$, deg\n.2\n12.0° Sideslip\nIncrement of pitching-moment coefficient, $\\Delta C_m$\n-30 -20 -10 0 10 20 30 40 50\nFlap deflection, $\\delta_f$, deg\n(a) Wing alone.\n(b) Wing + body.\n[Figure: NACA logo]\nFigure 21.- Increments of pitching-moment coefficient for various flap deflections at four angles of attack.\nNACA RM A59E04\n89", "timestamp": "2026-07-22T06:53:22.114565+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 5, "total_pages": 32, "image_filename": "19930085992_p5.jpg", "text": "```markdown\nNACA RM L9E17\n3\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| $W$ | weight of wing, pounds |\n| $W_W$ | weight of concentrated weight, pounds |\n| $l$ | length of wing, feet |\n| $b$ | half-chord of wing, feet |\n| $I_W$ | mass moment of inertia of weight about wing elastic axis, inch-pound-second$^2$ |\n| $I_{CG}$ | mass moment of inertia of wing about center of gravity, inch-pound-second$^2$ |\n| $I_{EA}$ | mass moment of inertia of wing about elastic axis, inch-pound-second$^2$ |\n| $EI$ | bending rigidity of wing, pound-inches$^2$ |\n| $GJ$ | torsional rigidity of wing, pound-inches$^2$ |\n| $\\rho$ | density of testing medium, slugs per cubic foot |\n| $m$ | mass of wing per unit length |\n| $\\frac{1}{\\kappa}$ | mass ratio $\\left( \\frac{m}{\\pi \\rho b^2} \\right)$ |\n| $r_\\alpha$ | nondimensional radius of gyration relative to elastic axis $\\left( \\sqrt{\\frac{I_{EA}}{12lmb^2}} \\right)$ |\n| $e_W$ | distance between elastic axis of wing and center of gravity of weight referred to half-chord |\n| $f_n$ | frequency, cycles per second |\n| $f_{h_1}$ | first bending natural frequency, cycles per second |\n```", "timestamp": "2026-07-22T06:53:24.451693+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 5, "total_pages": 24, "image_filename": "19930085991_p5.jpg", "text": "NACA RM L9I28\n\n$C_{m\\alpha}$ rate of change of pitching-moment coefficient with angle of attack in degrees $\\left(\\frac{dC_m}{d\\alpha}\\right)$\n\n$\\frac{dA}{dx}$ rate of change of nose cross-sectional area with nose length\n\n$\\epsilon$ angle between tangent to nose surface in plane of symmetry and nose X-axis, degrees\n\n$x$ distance from front of nose to any station, feet\n\n$a$ distance from front of nose to center of gravity, feet\n\n$\\Delta F$ normal force per unit length\n\n$q$ dynamic pressure, pounds per square foot $\\left(\\frac{1}{2}\\rho V^2\\right)$\n\n$\\rho$ density of air, slugs per cubic foot\n\n$V$ airspeed, feet per second\n\n$C_{L\\alpha}$ rate of change of lift coefficient with angle of attack in degrees $\\left(\\frac{dC_L}{d\\alpha}\\right)$\n\n$S_{F_t}$ area of one nose fin\n\nMODELS AND METHODS\n\nThe models tested represented $\\frac{1}{10}$ - to $\\frac{1}{23}$-scale models of possible airplane jettisonable nose sections. They were made of balsa and hardwood and ballasted with lead weights to simulate relative mass arrangements of the possible nose configurations at an altitude of 15,000 feet. The models had circular or near-circular cross sections and some had canopy portions or other protuberances. Sketches and mass characteristics of the models are presented in table I.\n\nDuring the tests, each model was held in the air stream of the Langley 20-foot free-spinning tunnel at various angles of attack from $0^\\circ$ to $180^\\circ$ and then released; the model was also launched with rotation applied about each of its three axes with the axes held alternately", "timestamp": "2026-07-22T06:53:26.143855+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 36, "total_pages": 52, "image_filename": "19930085911_p36.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 35\n\n1152\n\n[Figure: Graph showing \"Diffuser total-pressure recovery, $P_4/P_0$\" on the y-axis (ranging from 0.4 to 1.0) versus \"Time after release, $\\tau$, sec\" on the x-axis (ranging from 0 to 50). The plot shows a curve starting near 0.9, dipping slightly, dropping sharply around $\\tau=30$, recovering partially, then dropping again. A vertical dashed line labeled \"Impact\" is shown near $\\tau=45$. The NACA logo is in the bottom right corner of the graph area.]\n\n(c) Diffuser variables.\n\nFigure 8. - Continued. Time history of flight data and performance of ram-jet 16-A-3.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:26.352389+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 15, "total_pages": 30, "image_filename": "19930085970_p15.jpg", "text": "NACA RM A9E09 CONFIDENTIAL 13\n\n[Figure: Side view of a model on sting support, with scale bar and NACA logo labeled A-12059]\n\n(a) Side view.\n\n[Figure: Plan view of the same model on sting support, showing wings and tail, with scale bar and NACA logo labeled A-12060]\n\n(b) Plan view.\n\nFigure 2.— Model on sting support.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:27.460053+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 6, "total_pages": 40, "image_filename": "19930085997_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM A9I29\n\ndiffusor length aft of the inlet. Downstream of this point the cross-sectional areas of each duct were adjusted to arrive at the cross section of the common settling chamber.\n\nThe model dimensions are given in figure 1, and a photograph of the model tested is shown in figure 2. The model forebody was roughly triangular in cross section, and the scoops were located aft of the pilot enclosure in a position to utilize the oblique shock waves originating from the enclosure for external compression of the air stream. Aft of the main inlets, the external shape of the model was faired to adapt it to a cylindrical settling chamber.\n\nThe original design, hereafter designated configuration A, and five modifications to this design were tested. In configurations B, C, and D the model contours in the critical region near the entrance to the scoops were modified as shown in the line drawings of figure 3. In configurations E and F, the contours in the vicinity of the inlet were identical to those of configuration D; however, the model forebody was dropped with respect to the duct inlets. As noted in figure 1, the model with the forebody incidence reduced $2^\\circ$ is designated configuration E and that reduced $6^\\circ$ is configuration F.\n\nTESTS\n\nIn general, an analysis of the performance of the duct inlet design tested is concerned with a study of the following six variables: total-pressure recovery, free-stream Mach number, mass flow through the main scoops, mass flow through the boundary-layer scoops, angle of attack, and the inlet's contribution to the external drag. In the present tests the last variable was neglected, and the total-pressure recovery was chosen as the dependent variable. Thus, the performance of the model was studied by investigating the total-pressure recovery as a function of the remaining four variables.\n\nVariation in the mass flow into the main scoops was obtained by changing the position of the plug at the rear of the settling chamber (fig. 1). The total-pressure ratio across the exit plug was sufficient to maintain a sonic throat at the minimum area at all times. This fact, together with the known stagnation temperature and measurements of the average total pressure in the settling chamber, allowed the rate of mass flow through the scoops to be calculated (reference 1).\n\nVariation in the flow into the boundary-layer scoops was obtained by means of a valve in the line leading to the vacuum pump. The total pressure recovered in the boundary-layer scoops was measured by a pitot tube located at position 5 as shown in figure 1. The rate of mass flow through the boundary-layer scoops was determined by measuring the total pressure at the center and static pressure at the wall of a 3/4-inch pipe located outside the tunnel and by assuming a velocity profile corresponding to that for fully developed turbulent flow. Because it was\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:30.082151+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 12, "total_pages": 30, "image_filename": "19930085975_p12.jpg", "text": "10\nNACA RM L9E10\n\nCONFIDENTIAL\n\n[Figure: Three-view drawing of a wing with various sweep angles and dimensions annotated. Top view shows 50%c, 25%c, 3.6°, 7.1° L.E., -4.3°. Middle view shows 25.7°, 30°, 32.6°. Bottom view shows 35.1°, 40.7°, 45°, 46.7°, 48.4°. A NACA logo and a scale bar (0 to 10 inches) are also present.]\n\nTabulated Data\n\n| Wing | | Aileron | |\n| :--- | :--- | :--- | :--- |\n| Area | 2.25 sq.ft. | Type | True contour, sealed gap |\n| Aspect ratio | 4.0 | Chord | 20 % c |\n| Airfoil section | NACA 65A 006 | Span | 40% b/2 |\n| Span | 3.0 ft. | Inboard station | 55% b/2 |\n| Mean aerodynamic chord | 0.765 ft. | Outboard station | 95% b/2 |\n| Taper ratio | 0.60 | | |\n| Root chord | 11.25 in. | | |\n| Tip chord | 6.75 in. | | |\n\nCONFIDENTIAL\n\nFigure 1.— A drawing of the three wings tested in the present investigation.", "timestamp": "2026-07-22T06:53:30.385135+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 12, "total_pages": 46, "image_filename": "19930085983_p12.jpg", "text": "10 CONFIDENTIAL NACA RM A9I27\n\nSUMMARY OF RESULTS\n\nTests have been made of a cambered and twisted wing with the leading edge swept back $63^\\circ$ in combination with a slender fuselage. The wing was equipped with constant-chord elevons extending over the outer 50 percent of the span. The tests were conducted at a Reynolds number of 2.0 million and at Mach numbers ranging from 0.20 to 0.93. The following results were obtained:\n\n1. At low speed (M = 0.20) negative elevon deflections reduced the lift coefficient at which the loss of static longitudinal stability occurred, while at higher Mach numbers this lift coefficient generally increased with negative elevon deflections greater than $-5^\\circ$. (With the elevons undeflected the loss of static longitudinal stability generally occurred at a lift coefficient of about 0.5.)\n\n2. There was little effect of compressibility on the pitching-moment effectiveness of the elevons at lift coefficients of 0.20 or less. At higher lift coefficients the effectiveness increased with increasing Mach number.\n\n3. The effectiveness of the elevons in producing rolling moment was reduced slightly with increasing Mach number. The effectiveness was nearly constant at angles of attack between $-1^\\circ$ and $+8^\\circ$ for a Mach number of 0.20 and between $-1^\\circ$ and $+6^\\circ$ for the higher Mach numbers.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, California.\n\nREFERENCES\n\n1. Hopkins, Edward J.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.—Effects of Split Flaps, Elevons, and Leading-Edge Devices at Low Speed. NACA RM A9C21, 1949.\n\n2. Reynolds, Robert M., and Smith, Donald W.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.—Subsonic Mach and Reynolds Number Effects on the Characteristics of the Wing and on the Effectiveness of an Elevon. NACA RM A8D20, 1948.\n\n3. McCormack, Gerald M., and Walling, Walter C.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.—Investigation of a Large-Scale Model at Low Speed. NACA RM A8D02, 1949.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:30.861294+00:00"}