PDF/ocr_dataset_1070.jsonl

{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 34, "total_pages": 149, "image_filename": "19930083192_p34.jpg", "text": "30\nNACA TN 1976\n\nSlope of Lift Curve\n\nInvestigations to determine the proper slope of the lift curve to be used in gust-load calculations have been made mainly in the Langley gust tunnel. The information available for presentation is indirect and in many instances inconclusive because of interrelations of many factors. The factors involved in the computation of the proper slope of the lift curve include the determination of the airfoil section characteristics, aspect-ratio corrections, interference effects, the effects of configuration with reference to sweep and multiplanes and the effects of compressibility and power.\n\nSection characteristics.- As indicated by the results presented in figure 28, the conventional estimates of the lift-curve slope together with the unsteady-lift functions for infinite aspect ratio gave the best results in predicting the acceleration increment due to a sharp-edge gust for ordinary airfoils. The use of the laminar-flow section introduced airfoils having higher slopes of the lift curve. The increase was of considerable concern since it amounted to an increase of 10 percent in the load-factor increment. Brief consideration indicated that the slope of the lift curve depends on the steady-flow relation between the boundary-layer thickness and angle of attack. It was believed that under unsteady flow conditions the relation for steady flow would not apply and, therefore, the boundary layer would not have time to adjust itself during a sudden change in angle of attack. Since no experimental or theoretical information was available in connection with the probable slope of the lift curve, experimental evidence was needed.\n\nFor this purpose a skeleton model with a laminar-flow wing was tested in the Langley gust tunnel. The characteristics of the test model are given in table VII and figure 31. For the first tests the wing had a smooth surface. For the second tests, carborundum grains were glued over the leading edge of the wing back to 7.8 percent of the chord on both the upper and lower surfaces. The slopes of the lift curve for the smooth and rough wings have been included in table VII and are based on experimental data presented in reference 28 corrected to finite aspect ratio according to reference 27. In addition to the lift-curve slopes obtained from low-turbulence wind-tunnel tests, the theoretical lift-curve slope, based on the theory presented in reference 29, is also included in the table.\n\nThe average values of the acceleration increments for the smooth and rough wings and the probable errors are given in table VIII. The acceleration increments computed on the basis of the characteristics in table VII have also been included. Table VIII shows that the maximum acceleration increments for the two test conditions are essentially the same. Comparison of the 0.02g experimental difference with the calculated", "timestamp": "2026-07-22T06:34:42.825677+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 17, "total_pages": 26, "image_filename": "19930085890_p17.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:34:43.377105+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 3, "total_pages": 22, "image_filename": "19930085922_p3.jpg", "text": "2\nNACA RM No. L9023\n\nb wing span (3.093 ft on model)\nc chord parallel to plane of symmetry\n$\\bar{c}$ mean aerodynamic chord (M.A.C.) (1.05 ft on model)\n$C_l$ rolling-moment coefficient $\\left(\\frac{\\text{Rolling moment}}{qSb}\\right)$\n$C_{l_p}$ damping-in-roll coefficient $\\left(\\frac{\\partial C_l}{\\partial \\frac{pb}{2V}}\\right)$\nM free-stream Mach number (V/a)\np rate of roll, radians per second\nq dynamic pressure ($\\rho V^2/2$), pounds per square foot\nR Reynolds number ($\\rho V \\bar{c}/\\mu$)\nS wing area (3.17 sq ft on model)\nV free-stream velocity, feet per second\n$\\rho$ mass density of air, slugs per cubic foot\n$\\mu$ absolute viscosity, pound-seconds per square foot\n$\\alpha$ angle of attack, degrees\n$\\delta$ control-surface deflection with reference to wing-chord line\nparallel to plane of symmetry; positive deflection\ndown, degrees\npb/2V wing-tip helix angle, radians\n\n$$C_{l_\\delta} = \\frac{\\partial C_l}{\\partial \\delta}$$\n\n$$(pb/2V)_\\delta = \\frac{\\partial \\frac{pb}{2V}}{\\partial \\delta}$$", "timestamp": "2026-07-22T06:34:43.817398+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 17, "total_pages": 33, "image_filename": "19930085900_p17.jpg", "text": "16\nNACA RM L9D20\n\nCONFIDENTIAL\n\nStation 10\n42.2\n\n34 jets\n66 jets\n130 jets\n258 jets\n\nSide views of port chine jets\n\nTrim, deg\n10\n8\n6\n4\n2\n0\n\n2-inch spacing (34 jets)\n1-inch spacing (66 jets)\n1/2-inch spacing (130 jets)\n1/4-inch spacing (258 jets)\n\nEffective hydrodynamic lift and load on water, lb\n8\n7\n6\n5\n4\n3\n2\n1\n0\n-1\n-2\n\nLoad on water\n\nResistance, lb\n6\n5\n4\n3\n2\n1\n0\n\nSpeed, fps\n0 4 8 12 16 20 24 28 32 36 40 44 48 52 56 60\n\nNACA\n\nFigure 6.— Effect of jet spacing; chines 32 inches long; jets normal to center line.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:34:44.695375+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 56, "total_pages": 78, "image_filename": "19930082483_p56.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:34:48.117865+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 38, "total_pages": 62, "image_filename": "19930082918_p38.jpg", "text": "NACA TN 1940\n37\n\n[Figure: Microstructure image showing a fine, speckled grain pattern]\n\n(a) As-rolled.\n\n[Figure: Microstructure image showing a coarse, polygonal grain pattern]\n\n(b) Solution-treated 10 hours at $2200^\\circ$ F, water-quenched, and aged\n1 hour at $1400^\\circ$ F.\n\nNACA\n\nFigure 1.- Typical microstructures of heat A-1726 of low-carbon N-155 alloy.\nSmall-grained area. Cross sections of bar X100. Electrolytically\netched in 10 percent chromic acid.", "timestamp": "2026-07-22T06:34:49.563882+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 6, "total_pages": 52, "image_filename": "19930085911_p6.jpg", "text": "NACA RM E5F22 CONFIDENTIAL 5\n\nrates of fuel flow directly from the measured pressure drop. The telemetering antenna of the ram-jet unit was also a pitot-static tube from which the free-stream total and static pressures were obtained. A single flush wall orifice measured the static pressure in the diffuser inlet at station 2. This pressure was used to determine the transitions between subsonic and supersonic flow at station 2 when a shock passed over the orifice. Dynamic pressure in the diffuser at station 3 was the measured difference in pressure between two manifolds, one connected to eight total-pressure tubes and the other connected to two static-pressure tubes. The air flow was calculated at this station. The total pressure at the diffuser outlet was measured farther downstream (station 4) by eight total-pressure tubes manifolded together. A single flush wall orifice was placed just inside the exhaust-nozzle outlet to determine the outlet static pressure used for temperature and thrust calculations.\n\nPROCEDURE\n\nTheoretical calculations were made to determine the probable performance of the ram-jet units. The flight path of the ram-jet unit was calculated for combustion at a fuel-air ratio of 0.067 and a combustion efficiency of 60 percent. It was determined that the ram jet would reach its design conditions shortly before impact if it adhered to the calculated flight path. Air flow was calculated throughout the flight. The rates of fuel flow necessary to maintain a fuel-air ratio of 0.067 were then determined and plotted as a function of free-stream total pressure. Before each flight the fuel regulator, which was actuated by the free-stream total pressure, and the spring tensions in the fuel valves were adjusted to maintain the desired fuel-air ratio. Deviations from the scheduled free-stream Mach number had little effect on the required fuel flow because the regulator was actuated by the free-stream total pressure, which in turn largely determined the air flow. Deviation from the assumed combustion efficiency, however, would affect the air flow and alter the fuel-air ratio. The altitude for releasing each ram-jet unit was determined by the predicted fuel consumption in order to exhaust the fuel before impact and obtain drag data without combustion at high flight Mach numbers.\n\nImmediately after each drop, an atmospheric survey was made by the descending airplane to determine static temperature and pressure as a function of true altitude. True altitude was determined by radar-tracking the airplane. Weather balloons were released and radar-tracked to determine the wind-velocity corrections for the different altitudes.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:34:49.783348+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 38, "total_pages": 47, "image_filename": "19930083221_p38.jpg", "text": "```markdown\n36\nNACA TN No. 1824\n\nIf the boundary-value problem is one of the second kind, the vertical induced velocity is given on the wing and the load distribution is to be found. In this case the use of equation (59) leads to the consideration of an integral equation. Since, however, this integral equation is a common one in aerodynamic theory, certain established methods may be applied to it.\n\nAfter noting that $\\Delta\\varphi_o(x,y) = 0$ at the leading edge, integration by parts and introduction of the relation\n\n$$\n\\Delta v_o = \\frac{\\partial \\Delta \\varphi_o}{\\partial y}\n$$\n\nyields for perturbation potential the expression\n\n$$\n\\varphi = \\frac{1}{2\\pi} \\int_{b_1}^{b_2} \\Delta v_o(x,y_1) \\text{arc tan} \\frac{y-y_1}{z} dy_1 \\quad (61)\n$$\n\nIn the limit as $z$ approaches zero the derivative of $\\varphi$ with respect to $z$ reduces to the form\n\n$$\nw_o = -\\frac{1}{2\\pi} \\int_{-b}^{b} \\frac{\\Delta v_o(x,y_1) dy_1}{y-y_1} \\quad (62)\n$$\n\nFor a given distribution of $w_o$ over the plan form of the wing, equation (62) represents an integral equation to be solved for $\\Delta v_o(x,y)$ subject only to the condition that the Kutta-Joukowski condition is satisfied at all subsonic trailing edges. Once $\\Delta v_o(x,y)$ is determined it follows that\n\n$$\n\\Delta \\varphi_o = \\int_{-b}^{y} \\Delta v_o(x,y_1) dy_1 \\quad (63)\n$$\n\nand\n\n$$\n\\frac{\\Delta p}{q} = \\frac{2}{V_o} \\frac{\\partial \\Delta \\varphi_o}{\\partial x} \\quad (64)\n$$\n\nIn the present report the solution to the wing plan form shown in figure 15 will be presented. The value of $\\Delta \\varphi_o$ which satisfies\n```", "timestamp": "2026-07-22T06:34:52.454858+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 31, "total_pages": 31, "image_filename": "19930085859_p31.jpg", "text": "Wing alone\nWing fuselage\n\nNACA RM No. 19B25\n\n$C_{D_{L=0}}$\n.04\n0\n\n$\\left(\\frac{\\partial C_m}{\\partial C_L}\\right)_M$\n.2\n0\n-.2\n$C_L = 0$\n$C_L = 0.4$\n\n$\\left(\\frac{L}{D}\\right)_{Max}$\n20\n10\n0\n\n$\\left(\\frac{\\partial \\epsilon}{\\partial \\alpha}\\right)_M$\n.8\n.4\n0\n$h_t$\n-30\n0\n$C_L = 0$\n-30\n\n$\\left(\\frac{\\partial C_L}{\\partial \\alpha}\\right)_M$\n.12\n.08\n.04\n$C_L = 0$\n$C_L = 0.4$\n\n$\\left(\\frac{\\partial \\epsilon}{\\partial \\alpha}\\right)_M$\n.8\n.4\n0\n$h_t$\n0\n-30\n$C_L = 0$\n-30\n\n$\\gamma_{c.p.}$\n60\n40\n$C_L = 0.4$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n$h_t$\n$\\pm 30$\n0\n$C_L = 0$\n\nMach number, M\n.6 .7 .8 .9 1.0 1.1 1.2\n\nMach number, M\n.6 .7 .8 .9 1.0 1.1 1.2\n\nNACA\n\nFigure 13.- Summary of aerodynamic characteristics for a model with 35° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\n29", "timestamp": "2026-07-22T06:34:54.873643+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 36, "total_pages": 46, "image_filename": "19930085548_p36.jpg", "text": "NACA RM No. E8L30\n35\n\nMaximum cylinder pressure\nlb/sq in.\n1600\n1500\n1400\n\nimep, lb/sq in.\n400\n300\n200\n100\n\n[Figure: Two graphs plotting pressure against exhaust-gas temperature. The top graph shows \"Maximum cylinder pressure\" ranging from 1400 to 1600 lb/sq in. The bottom graph shows \"imep\" ranging from 100 to 400 lb/sq in. Both graphs share an x-axis for \"Exhaust-gas temperature, °F\" from 1000 to 2000. The bottom graph includes a legend distinguishing \"Experimental\" (solid line) and \"Calculated\" (dashed line) data, and a NACA logo.]\n\nExhaust-gas temperature, °F\n\nFigure 13. - Correlation of experimental and calculated values of indicated mean effective pressure for experimental cylinder. Inlet-manifold temperature, 400° F; inlet-manifold pressure, 100 pounds per square inch absolute.", "timestamp": "2026-07-22T06:34:54.990803+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 54, "total_pages": 65, "image_filename": "19930082546_p54.jpg", "text": "NACA TN No. 1870\n53\n\n(a) Experiment.\n(b) Theory. $R_e = 0.8R$.\n\nFigure 14.- Effect of tip Mach number on pressure amplitude of the first three harmonics for NACA 4-(5)(08)-03 propeller. B = 2; $\\beta_{0.75} = 10^\\circ$; $\\frac{M}{D} = -\\frac{1}{8}$; $\\frac{d}{D} = 0.083$.\n\nFigure 15.- Effect of tip Mach number and effective radius on pressure amplitude of the first harmonic for NACA 4-(5)(08)-03 propeller. B = 2; $\\beta_{0.75} = 10^\\circ$; $\\frac{M}{D} = -\\frac{1}{8}$; $\\frac{d}{D} = 0.083$.", "timestamp": "2026-07-22T06:34:58.405256+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 3, "total_pages": 47, "image_filename": "19930085919_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM No. A9C21\n\na wing with the leading edge swept back $63^\\circ$ and having an aspect ratio of 3.5. According to the theoretical considerations of reference 1, a wing of this plan form should be capable of efficient flight at supersonic Mach numbers up to 1.5. Experimental results from tests of wings of this plan form at high Mach or Reynolds numbers are presented in references 2, 3, and 4.\n\nA wing-fuselage combination having a wing of the plan form just described was investigated in one of the Ames 7- by 10-foot wind tunnels to evaluate the effectiveness of various flaps and particularly their capacity for eliminating the large changes in the longitudinal stability which have been found to occur above a lift coefficient of 0.4 (reference 4). In this connection, a drooped-nose flap and an extended-nose flap were tested in conjunction with trailing-edge flaps. Furthermore, an investigation was made to determine the optimum chordwise position of split flaps and the effectiveness of elevons of two different plan forms.\n\nNOTATION\n\nAll forces and moments are referred to the wind axes with the origin on an extension of the wing root chord at the same longitudinal position as a point at 25 percent of the wing mean aerodynamic chord.\n\n$C_D$ drag coefficient $\\left(\\frac{\\text{drag}}{qS}\\right)$\n\n$C_L$ lift coefficient $\\left(\\frac{\\text{lift}}{qS}\\right)$\n\n$C_l$ rolling-moment coefficient $\\left(\\frac{\\text{rolling moment}}{4qSb}\\right)$\n\n$C_m$ pitching-moment coefficient $\\left(\\frac{\\text{pitching moment}}{qS\\bar{c}}\\right)$\n\nA aspect ratio $\\left(\\frac{2b^2}{S}\\right)$\n\nb span of semispan wing perpendicular to the plane of symmetry, feet\n\nc wing chord parallel to plane of symmetry, feet\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:35:00.384272+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 32, "total_pages": 32, "image_filename": "19930085847_p32.jpg", "text": "30\nCONFIDENTIAL\nNACA RM A9D04\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (244, 109, 716, 858) -->\n\n(e) M = 0.83 ($C_L = 0.13$).\n\nFigure 8.- Concluded.\n\nCONFIDENTIAL\nNACA-Langley - 9-20-49 - 250", "timestamp": "2026-07-22T06:35:05.139038+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 68, "total_pages": 114, "image_filename": "19930086061_p68.jpg", "text": "64\nNACA RM L9J07\n\nLeft semispan\nRight semispan\nUpper\nLower\n\n(a) $\\psi = 0^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\n\n(b) $\\psi = 10^\\circ$\n\nFigure 21.- Pressure distribution about wing 2 at various angles of yaw; $\\alpha = 44.1^\\circ$.", "timestamp": "2026-07-22T06:35:05.177535+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 39, "total_pages": 72, "image_filename": "19930085491_p39.jpg", "text": "TABLE II.- SUMMARY OF RESULTS\n\n| Con- fig- ura- tion | R (mil- lion) | Lift | | Drag | | | | Lift-drag ratio | |\n|---|---|---|---|---|---|---|---|---|---|\n| | | $\\left(\\frac{dC_L}{d\\alpha}\\right)_{L=0}$ (per deg) | $\\left(\\frac{dC_L}{d\\alpha}\\right)_{opt}$ (per deg) | $C_{Dmin}$ | $\\left[\\frac{\\Delta C_D}{(\\Delta C_L)^2}\\right]_1$ | $\\left[\\frac{\\Delta C_D}{(\\Delta C_L)^2}\\right]_{opt}$ | $(k_a)_{L=0}$ | $(k_a)_{opt}$ | $(L/D)_{max}$ | $C_{Lopt}$ |\n| WF-57 | 0.62 | 0.041 (.063) | 0.048 (.063) | 0.0220 (.0233) | 0.296 (.189) | 0.305 (.189) | 0.70 (.68) | 0.80 (.68) | 6.1 (7.5) | 0.25 (.32) |\n| WF-60 | .62 | .040 (.057) | .047 (.057) | .0190 (.0161) | .303 (.184) | .311 (.184) | .69 (.60) | .82 (.60) | 6.5 (9.2) | .23 (.30) |\n| WF-63 | .31 | .038 (.051) | .045 (.051) | .0210 (*) | .336 (.185) | .354 (.185) | .77 (.54) | .90 (.54) | 5.8 (*) | .21 (*) |\n| | .62 | .038 (.051) | .045 (.051) | .0175 (.0133) | .310 (.185) | .318 (.185) | .67 (.54) | .80 (.54) | 6.7 (10.1) | .21 (.27) |\n| | .84 | .038 (.051) | .045 (.051) | .0160 (*) | .288 (.185) | .300 (.185) | .66 (.54) | .74 (.54) | 7.2 (*) | .21 (*) |\n| WF-67 | .62 | .034 (.043) | .040 (.043) | .0140 (.0112) | .347 (.196) | .354 (.196) | .67 (.48) | .77 (.48) | 7.1 (10.7) | .17 (.24) |\n| | .95 | .034 (.043) | .040 (.043) | .0135 (*) | .328 (.196) | .338 (.196) | .66 (.48) | .74 (.48) | 7.4 (*) | .17 (*) |\n| WF-70 | .62 | .033 (*) | .034 (*) | .0125 (*) | .410 (*) | .420 (*) | .78 (*) | .94 (*) | 6.9 (*) | .15 (*) |\n\nNote: For each wing the experimental value is given first and the corresponding theoretical value indicated in parentheses directly below. Where an asterisk is used, the theoretical value has not been computed.\n\n[Figure: NACA logo]\n\nCONFIDENTIAL\nCONFIDENTIAL\nNACA RM No. A8J04\n38", "timestamp": "2026-07-22T06:35:07.092094+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 5, "total_pages": 34, "image_filename": "19930085913_p5.jpg", "text": "4\nNACA RM L9F24\n\nAPPARATUS\n\nThe Langley 4.5-foot flutter research tunnel was used for this series of tests. The testing medium was air under approximate atmospheric conditions.\n\nThe models used in the investigation were flat-plate cantilever wings made from 24S-T aluminum with their leading edges rounded off. The unswept model had a length of 3 feet, a chord of 0.667 foot, and a thickness of 0.0900 inch. Two sweptback models were used, one having a sweep angle of $45^\\circ$ and the other, $60^\\circ$. The midchord lines of the swept models had a length of 3 feet, and the other properties were the same as for the unswept wing.\n\nEach model was mounted as a vertical rigid cantilever wing with its root at the top of the test section parallel to the air stream. This type of mounting resulted in flutter involving no bending or torsional displacements of the root. The sweptback models were effectively obtained by rotating the unswept wing about the intersection of the midchord line and the root. A sketch of each model is shown, with its data in table I. The wing properties based on the unswept wing are as follows:\n\n| Property | Value |\n| :--- | :--- |\n| Chord, 2b, feet | 0.667 |\n| Length, l, feet | 3 |\n| Aspect ratio (geometric) | 4.5 |\n| Taper ratio | 1 |\n| Airfoil section | Flat plate |\n| $e_h$, nondimensional | 0.01 (approx.) |\n| $e_\\alpha$, nondimensional | 0.005 (approx.) |\n| Thickness, t, inches | 0.0900 |\n| W, pounds | 2.735 |\n| $I_{CG}$, inch-pound-second$^2$ per inch | 0.00995 |\n| $I_{EA}$, inch-pound-second$^2$ per inch | 0.00995 |\n| EI, pound-inch$^2$ | $0.00506 \\times 10^6$ |\n| GJ, pound-inch$^2$ | $0.0080 \\times 10^6$ |\n| $x_\\alpha$, nondimensional | 0.0 |\n| $r_\\alpha^2$, nondimensional | 0.334 |\n| 1/k (standard air, no weight) | 34.1 |\n| $\\Lambda$, degrees | 0 |\n\nThe elastic axis and center of gravity of the wing sections were located at midchord. Because of the type of mount used in the investigation, it was necessary to use three wing models, each having the same properties, except as changed by the sweepback angle.", "timestamp": "2026-07-22T06:35:07.331899+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 18, "total_pages": 33, "image_filename": "19930085900_p18.jpg", "text": "NACA RM L9D20 CONFIDENTIAL 17\n\n[Figure: Two photographs of a model in a wind tunnel with water jets spraying from its sides. The top photo shows a model with two distinct spray patterns, and the bottom photo shows a model with a more diffuse spray pattern. A label on the bottom photo reads \"NACA L-59858\".]\n\n(a) 2-inch-spaced jets; trim, $8.0^\\circ$.\n\n(b) $\\frac{1}{4}$-inch-spaced jets; trim, $6.3^\\circ$.\n\nFigure 7.- Effect of jet spacing at 35 feet per second; 32-inch chines; station 10 to 42.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:35:11.847149+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 57, "total_pages": 78, "image_filename": "19930082483_p57.jpg", "text": "NACA TN NO. 1807\n\n[Figure: Turbine rotor mounted on a shaft, supported by two stands, with a scale bar labeled \"INCHES\" showing 0 to 2 inches. The rotor has a large disc with radial blades around its perimeter.]\n\nNACA  \nC-14257  \n1-30-46  \n\nFigure 2. - Turbine rotor used in investigation of effects of partial admission on performance of gas turbine.\n\n55", "timestamp": "2026-07-22T06:35:12.106915+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 32, "total_pages": 60, "image_filename": "19930085862_p32.jpg", "text": "30\nNACA RM No. L9A07\n\n$C_{Na}$\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n\n$\\delta_a$\n(deg)\n-25\n-20\n-15\n-10\n-5\n0\n5\n6\n10\n15\n20\n25\n\n$C_n$\n.01\n0\n-.01\n\n$C_l$\n.02\n.01\n0\n-.01\n-.02\n-.03\n\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\nNACA\n\n(a) $C_l$, $C_n$, and $C_{Na}$ against $\\alpha$.\nFigure 8.— Aileron characteristics of wing with extensible leading-edge\nand split flaps.", "timestamp": "2026-07-22T06:35:16.160740+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 35, "total_pages": 149, "image_filename": "19930083192_p35.jpg", "text": "NACA TN 1976\n31\n\ndifference of 0.26g shows that roughening an airfoil to produce a turbulent boundary layer has no effect on the slope of the lift curve that is applicable in the unsteady flow conditions of a sharp-edge gust. If these results are extended by the assumption that the flow conditions for a conventional airfoil in the steady state are simulated by those for the roughened low-drag wing, airfoil section characteristics are concluded to have no significant effect on the acceleration increment obtained in unsteady flow conditions of a sharp-edge gust.\n\nSince the most probable gust is one with a gradient distance of 10 chords, the averages of the maximum acceleration increments for the two test conditions were calculated for gusts with gradient distance up to 12 chords by applying the principle of superposition to the test results for the sharp-edge gust. The results of these calculations showed only $2\\frac{1}{2}$-percent difference between the average values. On the basis of this simple analysis, airfoil section characteristics have no significant effect on the acceleration increments for gusts with gradient distances from 0 to 12 chords. From the present data, however, it is not possible to specify the gradient distance at which the difference would become noticeable as the result of attaining quasi-steady conditions.\n\nFor the laminar-flow type of airfoil, tests in low-turbulence wind tunnels together with the aspect-ratio corrections of reference 27 should yield an adequate prediction of the acceleration increment due to a sharp-edge gust. The use of the corrections and the determination of section characteristics for an arbitrary airfoil are, however, still open to question since tests in low-turbulence tunnels on the arbitrary airfoil sections tend to show low slopes of the lift curve. In view of the many interrelated factors, some of which are discussed in subsequent paragraphs, it appears that the selection of an arbitrary section lift-curve slope of about 6 per radian will be adequate for gust-load calculations.\n\n**Aspect-ratio corrections.**- As previously noted, the data in figure 28 were obtained for conventional airplanes, but the data in table VIII apply to a skeleton airplane with little or no fuselage. The combined effects of fuselage interference, aspect-ratio corrections, unsteady-lift functions, and section characteristics are all involved and at the present time cannot be entirely segregated. For the time being, simple aspect-ratio correction factors appear to be adequate. The factor $\\frac{6A}{A + 2}$ has been used with satisfactory results in connection with gust-load studies.", "timestamp": "2026-07-22T06:35:18.560842+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 18, "total_pages": 26, "image_filename": "19930085890_p18.jpg", "text": "NACA RM No. E9C11\n\nOxygen\nnozzles\n\nDiborane\nnozzles\n\nInjection\nplate A\n\nCombustion chamber A\n$15\\frac{1}{4}''$\n\nExhaust\nnozzle\n\n$1.043''$\n\n$0.549''$\n\n$3''$\n\nOxygen\n\nDiborane\n\nInjection\nplate B\n\nCombustion chamber B\n$6\\frac{3}{8}''$\n\n[Figure: Interchangeable rocket-engine assemblies.]\n\nFigure 3. - Interchangeable rocket-engine assemblies.\n\nNACA\n\n17", "timestamp": "2026-07-22T06:35:18.564526+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 39, "total_pages": 62, "image_filename": "19930082918_p39.jpg", "text": "NACA TN 1940\n39\n\n(a) Ground on medium coarse grinder and finished with No. 1 emery paper.\n(b) After removal of 0.0045 inch of metal.\n(c) After removal of 0.010 inch of metal.\n(d) After removal of 0.017 inch of metal.\n\nFigure 2.- Diffraction patterns showing removal by electrolytic polishing of cold-worked surface on low-carbon N-155 alloy solution-treated 10 hours at 2200° F and water-quenched. Unfiltered molybdenum radiation (55 KV).", "timestamp": "2026-07-22T06:35:19.491933+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 4, "total_pages": 22, "image_filename": "19930085922_p4.jpg", "text": "NACA RM No. L9C23\n\nSubscripts:\n\n$a_{l}$ left aileron\n\n$a_{r}$ right aileron\n\nMODEL AND APPARATUS\n\nThe pertinent dimensions of the steel wing, ailerons, and vertical fins are given in figure 1. The ailerons were of true contour and constant chord and had sealed-nose gaps. The ordinates of the symmetrical airfoil section, which is not a standard NACA section, are given in table I.\n\nThe model was supported by a sting extending forward into the test section from a vertical strut located behind the model. The vertical strut was part of the wind-tunnel balance system and both the strut and a portion of the sting were shielded from the air stream by a fairing. A schematic drawing of the support system and rolling apparatus is shown in figure 2. The angle of attack of the model was changed by varying the angle of incidence of the wing relative to the sting. This was accomplished by utilizing various incidence blocks fitted into the sting. A photograph of the installation is shown in figure 3. The rolling-moment data were obtained from wind-tunnel balance measurements with the sting restrained in roll. When the model was permitted to roll freely under the moment created by the deflected aileron, the rate of roll was recorded electrically.\n\nTESTS AND PROCEDURE\n\nScope\n\nThe model was tested in two configurations, fins off and fins on, through a Mach number range of 0.40 to 0.91. The fins-off configuration was tested through an angle-of-attack range from $0.30^{\\circ}$ to $6.50^{\\circ}$; whereas, the fins-on configuration was tested only at an angle of attack of $0.30^{\\circ}$. However, both configurations were tested through an aileron-deflection range of $0^{\\circ}$ to $19.4^{\\circ}$ and only the left aileron was deflected.\n\nThe size of the model used in the present investigation resulted in an estimated choking Mach number of 0.94 and the data are believed to be reliable to a corrected Mach number of about 0.91. The variation of test Reynolds number with Mach number for average test conditions is presented in figure 4.", "timestamp": "2026-07-22T06:35:24.944806+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 21, "total_pages": 92, "image_filename": "19930085870_p21.jpg", "text": "20\nCONFIDENTIAL\nNACA RM No. L9D07\n\nAPPENDIX B\n\nCALCULATION OF PRESSURE DISTRIBUTIONS\n\nThe method and equations for computation of the pressure distributions over triangular wings are as follows:\n\nThe wing is broken down into two infinite wedge wings, and by superposition of the conical-flow solutions, the pressure distribution is obtained for each wedge. Combining the solutions yields the pressure distribution for the composite wing. The flow solutions for the given conditions are presented as follows:\n\n(1) Leading edge within Mach cone\n\n$$\n\\frac{\\Delta p}{q} = \\frac{4\\delta w_1}{\\pi\\beta\\sqrt{1 - w_1^2}} \\tan h^{-1} \\sqrt{\\frac{1 - w_1^2}{1 - \\left(\\frac{\\beta y}{x}\\right)^2}} \\quad 0 \\le w \\le w_1\n$$\n\n$$\n\\frac{\\Delta p}{q} = \\frac{4\\delta w_1}{\\pi\\beta\\sqrt{1 - w_1^2}} \\tan h^{-1} \\sqrt{\\frac{1 - \\left(\\frac{\\beta y}{x}\\right)^2}{1 - w_1^2}} \\quad w_1 \\le w \\le 1\n$$\n\n(2) Leading edge outside Mach cone\n\n$$\n\\frac{\\Delta p}{q} = \\frac{4\\delta w_1}{\\pi\\beta\\sqrt{w_1^2 - 1}} \\tan^{-1} \\sqrt{\\frac{w_1^2 - 1}{1 - \\left(\\frac{\\beta y}{x}\\right)^2}} \\quad 0 \\le w \\le 1\n$$\n\n$$\n\\frac{\\Delta p}{q} = \\frac{2\\delta w_1}{\\beta\\sqrt{w_1^2 - 1}} \\quad 1 \\le w \\le w_1\n$$\n\nwhere $w_1 = \\frac{\\tan \\epsilon}{\\tan m}$, $w$, in like manner, represents the position of a radial line through the apex of the wedge being analyzed; $\\delta$, the deflection or wedge half-angle with the proper sign attached; $y$, the span ordinate of the given chordwise station; and $x$, the chordwise ordinate at the same station with reference to the apex of the particular wedge.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:35:28.208470+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 39, "total_pages": 47, "image_filename": "19930083221_p39.jpg", "text": "NACA TN No. 1824\n37\n\nequation (62) is, in region 1\n$$ \\Delta\\Phi = -2w_o \\sqrt{x^2 \\tan^2 \\theta - y^2} \\quad (65) $$\n\nand in region 2\n$$ \\Delta\\Phi = -2w_o x \\tan \\theta \\left[ E(\\Psi_o, k_o) - k_o'^2 F(\\Psi_o, k_o) \\right] \\quad (66) $$\n\nwhere E and F are defined in the appendix and where\n$$ \\Psi_o = \\text{arc sin} \\sqrt{\\frac{x^2 \\tan^2 \\theta - y^2}{x^2 \\tan^2 \\theta - a_1^2}} \\quad (67) $$\n$$ k_o' = \\frac{a_1}{x \\tan \\theta} = \\sqrt{1 - k_o^2} \\quad (68) $$\n\nThe equation $y = a_1(x)$ of the trailing edge for which equation (66) is valid is given by the formula\n$$ a_1 = \\frac{k_o'}{E_o - k_o'^2 K_o} \\quad (69) $$\n\nwhich expresses $a_1$ explicitly as a function of $\\left( \\frac{a_1}{x \\tan \\theta} \\right)$. This particular choice of trailing-edge shape was used to simplify the analysis. The resulting plan form approaches a constant-chord wing as the span increases. The variation of $a_1$ with $x$ is given in figure 16; and figure 17 shows the relation between aspect ratio and span.\n\n[Figure: Regions diagram showing Regions 1 and 2, with labels $2\\theta$, $C_o$, $y=x \\tan \\theta$, $y=a_1(x)$, $t_o$, $S_o$, $x$, and $y$]\n\nFigure 15.— Dimensions and regions used in discussion of swept-back wings.\n\nThe loading coefficient is given in the two regions (defined in fig. 15) as follows:", "timestamp": "2026-07-22T06:35:29.774723+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 69, "total_pages": 114, "image_filename": "19930086061_p69.jpg", "text": "NACA RM L9J07\n65\n\nLeft semispan\nRight semispan\nUpper\nLower\n(c) $\\psi = 20^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\n(d) $\\psi = 35^\\circ$\n\nFigure 21.- Concluded.", "timestamp": "2026-07-22T06:35:29.978976+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 4, "total_pages": 47, "image_filename": "19930085919_p4.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 3\n\n$\\overline{c}$ mean aerodynamic chord $\\left(\\frac{\\int_{o}^{b} c^{2} d y}{\\int_{o}^{b} c d y}\\right)$, feet\n\n$l_{w}$ wing loading, pounds per square foot\n\n$q$ free-stream dynamic pressure $\\left(\\frac{1}{2} \\rho V^{2}\\right)$, pounds per square foot\n\n$R$ Reynolds number $\\left(\\frac{V \\bar{c}}{\\nu}\\right)$\n\n$r$ fuselage radius, feet\n\n$S$ area of semispan wing, square feet\n\n$V$ free-stream velocity, feet per second\n\n$V_{s}$ sinking speed, feet per second\n\n$x$ longitudinal distance, feet\n\n$y$ lateral distance, feet\n\n$\\alpha$ angle of attack of the wing chord plane, degrees\n\n$\\delta$ control-surface deflection measured in a plane normal to the hinge line (For positive deflections, the flap is below the wing-chord plane.), degrees\n\n$\\nu$ kinematic viscosity of air, feet squared per second\n\n$\\rho$ mass density of air, slugs per cubic foot\n\nSubscripts\n\n$d$ drooped-nose flap\n\n$e$ elevon\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:35:31.469679+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 37, "total_pages": 46, "image_filename": "19930085548_p37.jpg", "text": "36\nNACA RM No. E8L30\n\nCompression pressure, lb/sq in.\nCompression ratio, r\n\n[Figure: Graph plotting Compression pressure vs Compression ratio. Two curves are shown: a solid line with circles labeled \"Experimental\" and a dashed line labeled \"Calculated\". The experimental curve is annotated with $p_m r^{1.35}$ and the calculated curve with $p_e r^{1.35}$. The NACA logo is present in the bottom left corner of the graph area.]\n\nFigure 14. - Comparison of experimental and calculated values of compression pressures. Inlet-manifold temperature, 400° F; inlet-manifold pressure $p_m$, 100 pounds per square inch absolute; exhaust-manifold pressure $p_e$, 92.5 pounds per square inch absolute; exponent for compression, 1.35.", "timestamp": "2026-07-22T06:35:33.424993+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 19, "total_pages": 33, "image_filename": "19930085900_p19.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:35:35.723446+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 1, "total_pages": 92, "image_filename": "19930085951_p1.jpg", "text": "NACA RM L9D29\nNACA\nRM L9D29\nCONFIDENTIAL\nUNCLASSIFIED\nClassification Changed to\nUNCLASSIFIED\nNACA Research Author #21\ndated April 24, 1952\nDate\n6/23/52\nBy\nJ.E. Newlan\nNACA\n\nNACA\nRESEARCH MEMORANDUM\n\nTHE EFFECT OF BLADE-SECTION THICKNESS RATIOS\nON THE AERODYNAMIC CHARACTERISTICS\nOF RELATED FULL-SCALE PROPELLERS\nAT MACH NUMBERS UP TO 0.65\n\nBy\nJulian D. Maynard and Seymour Steinberg\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCASE FILE\nCOPY\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nJPL LIBRARY\nCALIFORNIA INSTITUTE OF TECHNOLOGY\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nJune 8, 1949\n\nUNCLASSIFIED\nCONFIDENTIAL\nJUN 13 1949", "timestamp": "2026-07-22T06:35:37.065687+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 58, "total_pages": 78, "image_filename": "19930082483_p58.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:35:40.640272+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 7, "total_pages": 52, "image_filename": "19930085911_p7.jpg", "text": "6\nCONFIDENTIAL\nNACA RM E9F22\n\nGENERAL METHOD OF CALCULATION\n\nFree-Stream Conditions\n\nFree-stream static temperatures and pressures encountered by\nthe ram-jet unit during a flight were calculated from atmospheric-\nsurvey and radar-tracking data. These pressures were used in the\nperformance calculations in preference to the telemetered static\npressures, which included possible shock-wave and interference\neffects from the ram-jet unit. The free-stream velocity of the\nram-jet unit was determined by plotting both the radar velocity,\nobtained by differentiating the flight path, and the velocity\nobtained by integrating the total acceleration of the ram jet. An\naverage velocity curve was faired and corrected for the variation\nin wind velocity encountered at different altitudes in order to\nobtain the relative air velocity. From the relative air velocity\nand the free-stream static temperature and pressure, it was pos-\nsible to calculate the remaining free-stream conditions of free-\nstream Mach number, total temperature, and total pressure. Cal-\nculated values of total pressure were believed to be less subject\nto possible error than the measured telemetered values and were\ntherefore used in the performance calculations wherever possible.\n\nPerformance Calculations\n\nAir flow through the unit was calculated from the data\nobtained at station 3, where the dynamic pressure in the diffuser\nwas measured. In order to calculate the air flow, the total pres-\nsure was assumed equal to the measured total pressure at the dif-\nfuser outlet and the total temperature was assumed equal to the\nfree-stream total temperature. The flow conditions at the diffuser\noutlet were then calculated by assuming isentropic compression, in\naccordance with the area change, between station 3 and the diffuser\noutlet. Fuel flow was determined from the telemetered static-\npressure drop across a calibrated orifice in the fuel line. Cal-\nculation of the fuel-air ratio completed the flow calculations at\nthe combustion-chamber inlet.\n\nOn the basis of wind-tunnel investigations of similar flame\nholders, the total-pressure drop across the flame holder was\nassumed to be twice the dynamic pressure at the upstream side of\nthe flame holder. For this total-pressure loss, the flow condi-\ntions at the downstream side of the flame holder were calculated\nassuming no change in total temperature from the free-stream value.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:35:42.102235+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 40, "total_pages": 72, "image_filename": "19930085491_p40.jpg", "text": "NACA RM NO. A8J04\n\nCONFIDENTIAL\n\n[Figure: Cutaway view of strain-gage balance system with labeled components]\n\n1 Balance cap  \n2 Front lift springs  \n3 Rear lift springs  \n4 Sting shroud  \n5 Sting moment gage  \n6 Balance beam  \n7 Drag spring  \n8 Pivot  \n\nNACA  \nA-11877\n\nFigure 1.— Cutaway view of strain-gage balance system.\n\nCONFIDENTIAL\n\n39", "timestamp": "2026-07-22T06:35:43.637654+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 55, "total_pages": 65, "image_filename": "19930082546_p55.jpg", "text": "54\nNACA TN No. 1870\n\n$$ \\frac{d\\bar{p}}{d\\bar{a}} \\times 10^4 $$\n\n[Figure: Graph showing $\\frac{d\\bar{p}}{d\\bar{a}} \\times 10^4$ vs $M_t$. The legend indicates symbols for $B=2$, $B=3$, and $B=4$. The x-axis ranges from 0.5 to 1.0. The y-axis ranges from 0.20 to 1.60. The NACA logo is present.]\n\nFigure 16.- Free-space oscillating pressure divided by power per unit disk area as a function of tip Mach number. NACA 4-(5)(08)-03 propeller. $B = 2$; $\\beta_{0.75} = 10^\\circ$; $\\frac{c}{D} = -\\frac{1}{8}$; $\\frac{d}{D} = 0.083$.\n\nOutin' # = amplitude/semiamplitude(2)\n\n[Figure: Graph showing Outin' # vs $M_t$. The legend indicates \"Calculated for $M_t = 0.75$\" (solid line with circles) and \"Calculated for $M_t = 1.00$\" (dashed line with squares). The x-axis ranges from 0 to 1.0. The y-axis ranges from 0.20 to 1.00. The NACA logo is present.]\n\nFigure 17.- Comparison of Gotin's simplified solution with equation (2) for the fundamental frequency of a two-blade propeller in the plane of rotation.", "timestamp": "2026-07-22T06:35:44.464116+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 6, "total_pages": 34, "image_filename": "19930085913_p6.jpg", "text": "NACA RM L9F24\n5\n\nTwo weights which were essentially the same were used; one was moved along the leading edge and the other, along the midchord line of each model. The properties of the weights were as follows:\n\n| | Leading-edge weight | Midchord weight |\n| :--- | :---: | :---: |\n| $W_w$, pounds . . | 3.12 | 3.12 |\n| $e_w$ . . . . . . | -1.0 | 0 |\n| $I_w$ . . . . . . | 0.0100 | 0.0098 |\n\nThe two weights were each about 14 percent heavier than the wing.\n\nVibration records of the bending and torsional oscillations of the wing during flutter were obtained electrically by the use of strain gages cemented on the wing. The strain gages were connected through a system of bridges and amplifiers to a recording oscillograph. Two sets of gages were used on each model. One set of gages was mounted on the midchord line approximately 4 inches from the root and the other, on the same line about 4 inches from the tip. The approximate location of the strain gages is illustrated as follows:\n\n[Figure: Diagram of a wing planform with a hatched root on the left. A vertical dashed line labeled 'A' at the top and bottom indicates a cross-section. To the left of the line, there are two circles labeled '1' stacked vertically, and a square labeled '2' to their right. To the right of the line, there are two circles labeled '3' stacked vertically, and a square labeled '4' to their right.]\n\nThe squares represent the bending gages and the circles, the torsion gages. The numbers 1, 2, 3, and 4 represent the root-torsion, root-bending, tip-torsion, and tip-bending gages, respectively.", "timestamp": "2026-07-22T06:35:45.154036+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 33, "total_pages": 60, "image_filename": "19930085862_p33.jpg", "text": "NACA RM No. L9A07\n31\n\n$P_R$\n1.2\n.8\n.4\n0\n-.4\n-.8\n\n$\\delta_a$ (deg)\n-25\n-20\n-15\n-10\n-6\n-3\n0\n3\n6\n10\n15\n20\n25\n\n$C_{h_a}$\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\n(b) $C_{h_a}$ and $P_R$ against $\\alpha$.\nFigure 8.— Continued.", "timestamp": "2026-07-22T06:35:45.774330+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 70, "total_pages": 98, "image_filename": "19930086073_p70.jpg", "text": "68\n\nLift coefficient, $C_L$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:35:48.569586+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 19, "total_pages": 26, "image_filename": "19930085890_p19.jpg", "text": "```markdown\n18\nNACA RM No. E9C11\n\n<!-- Image (109, 117, 297, 407) -->\nCombustion-chamber pressure\n\n<!-- Image (325, 117, 562, 400) -->\nThrust\n\nValues from records\nThrust, (lb) 104\nDiborane flow, (lb/sec) 0.164\nOxygen flow, (lb/sec) 0.258\nFuel, (percent by weight) 39.8\nSpecific impulse\n(83.3 percent theo-\nretical), (lb-sec/lb) 247.3\nCombustion pressure,\n(lb/sq in. gage) 285-300\n\n<!-- Image (117, 503, 357, 806) -->\nDiborane flow\n\n<!-- Image (373, 503, 604, 806) -->\nOxygen flow\n\n<!-- Image (608, 387, 817, 806) -->\nCombustion-chamber pressure\n\nFigure 4. - Typical data obtained during experiments.\n\nNACA\n```", "timestamp": "2026-07-22T06:35:49.169717+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 40, "total_pages": 62, "image_filename": "19930082918_p40.jpg", "text": "NACA TN 1940\n41\n\nX-ray beam\nCollimator\nX-ray tube\n$2\\frac{1}{2}$\"\nFilm\n$40^\\circ$\n$15^\\circ$\nDiffracted X-ray\nbeams\nSpecimen\nNACA\n\nFigure 3.- Sectional schematic diagram of experimental setup for\nstudying surfaces of polished samples.\n\nNACA\n\nFigure 4.- Electrolytic cell for gross metal removal X$\\frac{1}{2}$.", "timestamp": "2026-07-22T06:35:52.100440+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 36, "total_pages": 149, "image_filename": "19930083192_p36.jpg", "text": "32\nNACA TN 1976\n\nSwept wings.- Table IX (from reference 30) shows the results of gust-tunnel tests on a wing swept back 45°. The tests were made in a sharp-edge gust and the acceleration increments for a gradient distance of 9 chords were obtained by superposition. All results are corrected to zero pitch of the model. Calculated values are also included. One set of values is based on the results of wind-tunnel tests of the airplane model, tail off, and the other set differs from the first in that the slope of the lift curve for the swept wing was taken as that for the straight wing multiplied by the cosine of the sweep angle. In both calculations, the unsteady-lift function for the wing $C_{L_g}$ was obtained by means of strip theory.\n\nInspection of the results given in table IX shows that the use of the cosine law for predicting the slope of the lift curve for the swept wing gives excellent results, the differences between experiment and calculation being within the experimental error. The use of lift-curve slopes from steady-flow tests to calculate the acceleration increments for the swept wing yields values some 20 percent below experimental results. Similar tests on a 45° sweptforward wing verify this result.\n\nThe low value of acceleration increments calculated from the results of wind-tunnel tests has been of concern. At present, no definite evidence is available to explain this discrepancy, but it is believed that the difference can be ascribed to the behavior of the boundary layer in the unsteady flow condition that exists during traverse of a gust. The discrepancy emphasizes the importance in the prediction of gust load factors of using the proper lift-curve slope. On the basis of limited data better results seem to be obtained if the cosine law of variation is used rather than wind-tunnel results.\n\nThe results obtained to date apply to a definite method of sweeping the wing (rotating the wing panel) to obtain the equivalent straight wing. Other systems lead to other combinations of aspect ratio and sweep and may yield different results. In such calculations, the same procedure as that described in the present paper should be utilized.\n\nScale effects.- Tests in the old and new gust tunnels of models of the Boeing B-247 airplane (table X) were made to obtain a measure of any scale effects. A 3-foot-span model was tested in the old gust tunnel as a reference for design requirements. When the new gust tunnel was built this model and one dynamically similar but twice as large were tested. The results of all three tests are shown in figure 32 as a plot of the acceleration ratio as a function of gradient distance in chords.\n\nInspection of figure 32 shows excellent agreement between the results for the two models; this fact indicates that for the conditions tested no scale effect was present. The largest discrepancy is for a sharp-edge gust and is no greater than differences in results for tests", "timestamp": "2026-07-22T06:35:52.800534+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 38, "total_pages": 46, "image_filename": "19930085548_p38.jpg", "text": "NACA RM No. E8L30\n37\n\nMaximum cylinder pressure, lb/sq in.\nFuel-air ratio\n\n| Inlet-manifold pressure (lb/sq in. abs.) | |\n| :--- | :--- |\n| ○ | 80 |\n| □ | 100 |\n| ◇ | 120 |\n| △ | 135 |\n\n[Figure: A line graph showing the relationship between Fuel-air ratio (x-axis, 0 to .05) and Maximum cylinder pressure (y-axis, 600 to 2200 lb/sq in.). Four curves represent different inlet-manifold pressures (80, 100, 120, 135 lb/sq in. abs.), indicated by symbols ○, □, ◇, and △ respectively. All curves show an increasing trend in pressure as the fuel-air ratio increases. The NACA logo is present in the top right corner of the graph area.]\n\nFigure 15. - Effect of inlet-manifold pressure and fuel-air ratio on maximum cylinder pressure in experimental cylinder. Compression ratio, 5.25; inlet-manifold temperature, 400° F.", "timestamp": "2026-07-22T06:35:58.345542+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 5, "total_pages": 22, "image_filename": "19930085922_p5.jpg", "text": "```markdown\n4\nNACA RM No. L9C23\n\nCorrections\n\nA small tare correction in the form of bearing friction was applied to the results as an increment of damping-in-roll coefficient equal to -0.005. The rolling-moment coefficients and Mach numbers have been corrected for blocking by the model and its wake, by the method of reference 2.\n\nReduction of Data\n\nThe coefficient of damping in roll $C_{l_p}$ is defined as follows:\n\n$$C_{l_p} = \\frac{\\partial C_l}{\\partial \\frac{pb}{2V}} = - \\frac{C_{l_\\delta}}{\\left(\\frac{pb}{2V}\\right)_\\delta}$$\n\nwhere the expressions $C_{l_\\delta}$ and $(pb/2V)_\\delta$ were evaluated graphically as the slopes of the curves of the rolling-moment coefficient $C_l$ plotted against aileron deflection $\\delta$ and of the nondimensional steady rate of rolling $(pb/2V)$ plotted against aileron deflection $\\delta$, respectively.\n\nRESULTS AND DISCUSSION\n\nThe results are presented in the following figures:\n\n| | | Figure |\n| :--- | :--- | :--- |\n| Basic force data $C_l$ plotted against M | | |\n| Fins off | | 5 |\n| Fins on | | 6 |\n| Basic rolling data $\\frac{pb}{2V}$ plotted against M | | |\n| Fins off | | 7 |\n| Fins on | | 8 |\n| Summary data | | |\n| Aileron effectiveness $C_{l_\\delta}$ and $\\left(\\frac{pb}{2V}\\right)_\\delta$ plotted against M | | |\n| Fins off | | 9 |\n| Fins on | | 10 |\n| Damping-in-roll coefficient $C_{l_p}$ plotted against M | | |\n| Fins off | | 11 |\n| Fins on | | 12 |\n```", "timestamp": "2026-07-22T06:36:03.011497+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 22, "total_pages": 92, "image_filename": "19930085870_p22.jpg", "text": "NACA RM No. L9D07 CONFIDENTIAL 21\n\nREFERENCES\n\n1. Jones, Robert T.: Wing Plan Forms for High-Speed Flight. NACA Rep. No. 863, 1947.\n\n2. Busemann, A.: Aerodynamischer Auftrieb bei Überschallgeschwindigkeit. Luftfahrtforschung, Bd. 12, Nr. 6, Oct. 3, 1935, pp. 210-220.\n\n3. Jones, Robert T.: Properties of Low-Aspect-Ratio Pointed Wings at Speeds below and above the Speed of Sound. NACA Rep. No. 835, 1946.\n\n4. Jones, Robert T.: Estimated Lift-Drag Ratios at Supersonic Speed. NACA TN No. 1350, 1947.\n\n5. Brown, Clinton E.: Theoretical Lift and Drag of Thin Triangular Wings at Supersonic Speeds. NACA Rep. No. 839, 1946.\n\n6. Stewart, H. J.: The Lift of a Delta Wing at Supersonic Speeds. Quarterly Appl. Math., vol. IV, no. 3, Oct. 1946, pp. 246-254.\n\n7. Puckett, Allen E.: Supersonic Wave Drag of Thin Airfoils. Jour. Aero. Sci., vol. 13, no. 9, Sept. 1946, pp. 475-484.\n\n8. Puckett, A. E., and Stewart, H. J.: Aerodynamic Performance of Delta Wings at Supersonic Speeds. Jour. Aero. Sci., vol. 14, no. 10, Oct. 1947, pp. 567-578.\n\n9. Von Kármán, Theodore: Supersonic Aerodynamics - Principles and Applications. Jour. Aero. Sci., vol. 14, no. 7, July 1947, pp. 373-409.\n\n10. Ellis, Macon C., Jr., and Hasel, Lowell E.: Preliminary Tests at Supersonic Speeds of Triangular and Swept-Back Wings. NACA RM No. L6L17, 1947.\n\n11. Vincenti, Walter G., Nielsen, Jack N., and Matteson, Frederick H.: Investigation of Wing Characteristics at a Mach Number of 1.53. I - Triangular Wings of Aspect Ratio 2. NACA RM No. A7I10, 1947.\n\n12. Gray, W. E.: A Simple Visual Method of Recording Boundary Layer Transition (Liquid Film). TN No. Aero 1816, British R.A.E., Aug. 1946.\n\n13. Delameter, H. D., and Miller, W.: Model XAAM-2. Preliminary Analysis of Wind Tunnel Test Data. First Daingerfield Wind Tunnel Test Period. Mach Number = 1.73. Rep. No. SM-13256, Douglas Aircraft Co., Inc., April 19, 1948.\n\n14. Hayes, W. D., Browne, S. H., and Lew, R. J.: Linearized Theory of Conical Supersonic Flow with Application to Triangular Wings. Rep. No. NA-46-818, North American Aviation, Inc., Sept. 30, 1946. (Revised June 26, 1947.)\n\nCONFIDENTIAL\n\n15. [illegible] (NACA Conf. July 8-10 1953)", "timestamp": "2026-07-22T06:36:05.768902+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 40, "total_pages": 47, "image_filename": "19930083221_p40.jpg", "text": "38\nNACA TN No. 1824\n\n$$ \\frac{a_1}{c_o \\tan \\theta} $$\n\n[Figure: Graph showing trailing-edge position of the swept-back wings studied. The vertical axis is labeled $x/c_o$ ranging from 1.0 to 1.8. The horizontal axis is labeled $a_1 / (c_o \\tan \\theta)$ ranging from 0 to 1.0. A curve starts at (0, 1.0) and decreases to approximately (1.0, 1.7).]\n\nFigure 16.— Graph showing trailing-edge position of the swept-back wings studied.\n\n$$ \\frac{A}{\\tan \\theta} $$\n\n[Figure: Relation between aspect ratio and wing semi-span. The vertical axis is labeled $A / \\tan \\theta$ ranging from 4 to 8. The horizontal axis is labeled $s_o / (c_o \\tan \\theta)$ ranging from 1.0 to 2.6. A curve starts at (1.0, 4) and increases to approximately (2.6, 8).]\n\nFigure 17.— Relation between aspect ratio and wing semi-span.\n\nRegion 1\n\n$$ \\frac{\\Delta p}{q \\alpha} = \\frac{4x \\tan^2 \\theta}{\\sqrt{x^2 \\tan^2 \\theta - y^2}} \\quad (70) $$\n\nRegion 2\n\n$$ \\frac{\\Delta p}{q \\alpha} = 4 \\tan \\theta \\left[ E(\\psi_o, k_o) + \\frac{y}{x \\tan \\theta} \\sqrt{\\frac{y^2 - a_1^2}{x^2 \\tan^2 \\theta - y^2}} - \\frac{E_o}{K_o} F(\\psi_o, k_o) \\right] \\quad (71) $$\n\nThis load distribution is shown in figure 18 for a triangular and a swept-back wing. It is seen that the loading at sonic speed bears a close resemblance to those found at higher Mach numbers. Two similarities of note are the discontinuity in the pressure gradient at the Mach wave originating from the trailing edge of the root chord and the satisfying of the Kutta condition only where the trailing edge is subsonic. The lift and induced drag coefficients are given, respectively, by", "timestamp": "2026-07-22T06:36:08.378584+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 7, "total_pages": 34, "image_filename": "19930085913_p7.jpg", "text": "6\nNACA RM L9F24\n\nThe system used in obtaining the proper phase-angle relationship between the bending and torsional stresses of the wing as recorded in table I is as follows:\n\n[Figure: Diagram showing a wing section with labels: Leading edge (-1), Midchord (0), Trailing edge (1), Zero displacement, h, $\\alpha$, V cos $\\Lambda$, c.g. of wing section. Caption: Section A-A]\n\nThe preceding sketch shows the relative directions of positive bending (h) and torsional ($\\alpha$) displacements of the wing section. A couple, which twisted the wing in the positive $\\alpha$-direction, was applied at the tip. This action induced positive twist at each section of the wing; therefore, the direction in which the torsion-gage traces moved on the oscillograph record for positive twist at the gage stations was obtained. A force was then applied, which deflected the tip in the positive h-direction, thus producing positive bending curvature at each section of the wing; therefore, the direction in which the bending-gage traces moved on the oscillograph record for positive bending curvature at the gage stations was determined. Thus, the phase-angle relationships between the strain-gage traces on the oscillograph record for positive torsional and bending stress at the gage stations was established. Each model was treated in the same manner. The root-torsion-gage trace was used as a reference. If the traces of the other gages were displaced in the same direction as the reference-gage trace for positive twist or bending curvature, they were in phase ($0^\\circ$); if not, they were out of phase ($180^\\circ$). The following table gives the results of the calibration:", "timestamp": "2026-07-22T06:36:11.865986+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 71, "total_pages": 98, "image_filename": "19930086073_p71.jpg", "text": "```markdown\nNACA RM A9H04\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\n+10.8 0 -10.8\nRight aileron deflection, $\\delta_{a_R}$, deg\n\n(b) $C_L$ vs $C_D$.\n\nFigure 15.- Continued.\n\n[Figure: Graph plotting Lift coefficient vs Drag coefficient with three curves corresponding to right aileron deflections of +10.8, 0, and -10.8 degrees. The NACA logo is visible in the bottom right corner of the plot area.]\n\n69\n```", "timestamp": "2026-07-22T06:36:14.528280+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 8, "total_pages": 52, "image_filename": "19930085911_p8.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 7\n\nThe flow conditions after the combustion process were determined by two independent methods. When no choking existed at the engine outlet, the calculations were based on a trial-and-error solution involving the relations for heat addition in a constant-area pipe. For the choking condition, the calculations involved the continuity expression for mass flow. The heat-addition expression could also be used for the choking condition. The good agreement obtained by both methods tended to justify the necessary assumptions made for the constant-area heat-addition expression. When choking occurred in the exhaust nozzle, the static pressure at the outlet was obtained from the telemetered data. The total temperature at the outlet necessary for this condition was then calculated by using the continuity expression for mass flow. When subsonic velocity existed in the exhaust nozzle, a trial-and-error solution was required. In this case, it was assumed that all the heat was released in the constant-area section of the combustion chamber and that no friction losses occurred in the combustion chamber and the exhaust nozzle. The trial-and-error solution involved tentative selection of a total temperature after combustion at station 6, calculation of total pressure and Mach number after combustion, and calculation of flow conditions at the engine outlet by isentropic expansion for the area change. If the static pressure at the outlet did not equal the measured static pressure, the calculations were repeated for different values of total temperature after combustion. In some cases, the telemetered static pressure at the engine outlet was unreliable due to vibratory burning. The free-stream static pressure was used as a substitute value under these conditions. Both methods of calculation involved the assumption that all the fuel was vaporized in the ram jet. Included in the calculations were the variations in specific-heat ratio of the gases with combustion.\n\nNet thrust was calculated as the difference in the momentum of the exhaust gases at the engine outlet and the momentum of the free-stream air. External drag was determined as the difference between the net thrust and the product of the net acceleration times the weight of the ram-jet unit. Thrust and external-drag coefficients were calculated for the maximum cross-sectional area.\n\nCombustion efficiency was determined as the increase in enthalpy of the fuel-air mixture across the combustion chamber divided by the available chemical energy of the fuel and the flares. The gas total-temperature ratio was obtained by dividing the calculated total temperature of the exhaust gases at the engine outlet by the total temperature of the air at the combustion-chamber inlet.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:36:18.144320+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 41, "total_pages": 62, "image_filename": "19930082918_p41.jpg", "text": "NACA TN 1940\n43\n\n<!-- Image (96, 110, 470, 450) -->\n(a) Unaged.\n\n<!-- Image (526, 113, 900, 452) -->\n(b) Aged 0.5 hour.\n\n<!-- Image (96, 504, 469, 846) -->\n(c) Aged 1 hour.\n\n<!-- Image (524, 506, 898, 848) -->\n(d) Aged 3 hours.\n\nNACA\nFigure 5.- Effect of aging at $1200^\\circ$ F on microstructure of low-carbon\nN-155 alloy solution-treated 10 hours at $2200^\\circ$ F and water-\nquenched. Cross sections of bar X1000. Electrolytically etched\nin 10 percent chromic acid.", "timestamp": "2026-07-22T06:36:21.412993+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 20, "total_pages": 26, "image_filename": "19930085890_p20.jpg", "text": "```markdown\nNACA RM No. E9C11\n19\n\n| Combustion- chamber pressure $P_c$ (lb/ sq in. abs.) | Character- istic length $L^*$ (in.) | Injection system |\n| :--- | :--- | :--- |\n| $\\diamond$ 300$\\pm$15 | 325 | Eight hole |\n| $\\diamond$ 343 | 325 | Eight hole |\n| $\\square$ 300$\\pm$15 | 159 | Eight hole |\n| $\\Delta$ 300$\\pm$15 | 325 | Four hole |\n\nSpecific impulse, I, lb-sec/lb\n\nStoichiometric\n\nTheoretical; equilibrium expansion (reference 1)\nTheoretical; equilibrium expansion; 95-percent diborane; total nozzle divergent angle, 30$^\\circ$\nExperimental, corrected\nExperimental\n\nRatio of fuel weight to total propellant weight\n\nFigure 5. - Theoretical and experimental specific impulse of 100- pound-thrust rocket engine using liquid diborane and liquid oxygen.\n\nNACA\n```", "timestamp": "2026-07-22T06:36:21.848754+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 41, "total_pages": 72, "image_filename": "19930085491_p41.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:36:22.844663+00:00"}