PDF/ocr_dataset_1037.jsonl

{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 50, "total_pages": 66, "image_filename": "19930082914_p50.jpg", "text": "NACA TN No. 1857\n49\n\n[Figure: A black and white image showing a pattern of wavy, horizontal lines. The lines are more distorted and irregular in the upper portion of the image and become progressively straighter and more uniform towards the bottom.]\n\n(b) $1\\frac{1}{2}$ to 3 inches from nozzle.\nFigure 11.- Continued.\nNACA", "timestamp": "2026-07-22T05:26:12.164943+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 32, "total_pages": 33, "image_filename": "19930085544_p32.jpg", "text": "NACA RM No. L8K26\n31\n\n<!-- Image (188, 110, 860, 934) -->\n\nFigure 14.- Variation of bending-moment coefficient with blade position. Propeller 4-(3.9)(07)-0345-B.\n$\\beta_{.75}C_o^o$; $J_1=1.2$; $C_{n_1}=.4$.", "timestamp": "2026-07-22T05:26:12.866632+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 27, "total_pages": 36, "image_filename": "19930085869_p27.jpg", "text": "NACA RM L59D15\n\nCONFIDENTIAL\n\nTrim, deg\n16\n14\n12\n10\n8\n6\n4\n2\n0\n\nMaximum trim\nUpper limit\nIncreasing trim\nDecreasing trim\nUnstable\nStable\nUnswept hull\nSwept hull\nUnswept hull\n\nSpeed coefficient, $C_V$\n0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0\n\nFigure 8.- Trim limits of stability.\nCONFIDENTIAL\n\n25", "timestamp": "2026-07-22T05:26:13.688333+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 19, "total_pages": 37, "image_filename": "19930085889_p19.jpg", "text": "13\nNACA RM L9F14\n\nCONFIDENTIAL\n\n<!-- Image (139, 143, 744, 480) -->\n\n<!-- Image (141, 563, 733, 826) -->\n\nCONFIDENTIAL\n\nFigure 1.- System of axes used. Arrows indicate positive direction\nof angles, forces, and moments.", "timestamp": "2026-07-22T05:26:18.741598+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 23, "total_pages": 96, "image_filename": "19930085880_p23.jpg", "text": "NACA RM No. L9C03\n21\n\nResistance, lb\n.30\n.25\n.20\n.15\n.10\n.05\n0\n\nSpeed, fps\n0 10 15 20 25 30 35\n\nWetted area\n(sq ft)\n0\n.05\n.10\n.15\n.20\n.25\n.30\n\n[Figure: Two vertical rectangular shapes labeled 250A and 250B]\n\nNACA\n\nFigure 10.- Aerodynamic drag of models 250A and 250B.", "timestamp": "2026-07-22T05:26:30.875244+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 26, "total_pages": 46, "image_filename": "19930085542_p26.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:26:31.413189+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 5, "total_pages": 42, "image_filename": "19930085938_p5.jpg", "text": "```markdown\n4\nNACA RM No. L9B04\n\nPROCEDURE\n\nTake-off Stability\n\nThe center-of-gravity limits of stability were determined by making accelerated runs at a constant acceleration of one foot per second per second to take-off, with fixed elevators and full power. A sufficient number of center-of-gravity locations and elevator settings were tested to define the stability limits. A center-of-gravity limit of stability is defined as that condition at which the amplitude of trim oscillation reaches a value of $2^\\circ$ or the trims at high speeds become less than $2^\\circ$. Trims of less than $2^\\circ$ at high speeds were considered to be unsafe for practical operation. The variation of trim with speed was also observed during these runs. To find the trim limits of stability, the towing carriage was held at constant speeds, while the model trim was slowly increased or decreased until the porpoising limit was crossed.\n\nLanding Stability\n\nPrior to landing, the model was trimmed in the air to the desired contact trim with the carriage held at a constant speed slightly greater than the model flying speed. The carriage was then decelerated at a constant rate of three feet per second per second allowing the model to glide onto the water with fixed elevators in simulation of an actual landing. The descent to the water from flight was made from a height of 0.3b above the water. This procedure was used to hold the sinking speeds to reasonable values (approx. 300 ft/min full size). After the first contact the rise restriction was removed. Landings were made with the center of gravity located at 0.20$\\bar{c}$, 0.30$\\bar{c}$, and 0.40$\\bar{c}$, using one-quarter static thrust.\n\nSpray\n\nThe range of speeds over which spray was in the propellers was defined for a series of gross loads. (See reference 5.) The model was free to trim about the 0.30$\\bar{c}$ location of the center of gravity with the elevators fixed at $0^\\circ$. Constant-speed runs were made at full power starting with a light load on the water and increasing the load until spray entered the propellers.\n\nResistance\n\nThe resistance characteristics were obtained with the wing and tail surfaces removed. The tail booms of both configurations were supported by auxiliary means. A lift curve was determined from the variation in take-off speed with trim observed in the take-off stability tests. The load on\n```", "timestamp": "2026-07-22T05:26:33.465733+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 18, "total_pages": 31, "image_filename": "19930085881_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:26:33.669163+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 60, "total_pages": 99, "image_filename": "19930082511_p60.jpg", "text": "58\nNACA TN No. 1826\n\nFor $\\xi < a$,\n$$\n\\begin{aligned}\n(-1)^n \\sin \\frac{|b - \\xi|}{b - a} n\\pi &= -(-1)^n \\sin \\left[ \\frac{b - a + (a - \\xi)}{b - a} n\\pi \\right] \\\\\n&= \\sin n\\pi \\frac{a - \\xi}{b - a} \\\\\n&= \\sin n\\pi \\frac{|\\xi - a|}{b - a}\n\\end{aligned}\n$$\n\nFor $\\xi > b$,\n$$\n\\begin{aligned}\n(-1)^n \\sin \\frac{|b - \\xi|}{b - a} n\\pi &= (-1)^n \\sin \\left[ \\frac{\\xi - a - (b - a)}{b - a} n\\pi \\right] \\\\\n&= \\sin n\\pi \\frac{\\xi - a}{b - a} \\\\\n&= \\sin n\\pi \\frac{|\\xi - a|}{b - a}\n\\end{aligned}\n$$\n\nThe first term on the right-hand side of equation (A3) is thus equal to zero for $\\xi < a$ or $\\xi > b$. The desired expressions for $P_{mn}(\\xi, \\rho)$, $n \\neq 0$, are therefore, for $a \\le \\xi \\le b$,\n$$\nP_{mn}(\\xi, \\rho) = \\frac{J_m \\left( 4\\zeta \\frac{n\\pi}{b - a} \\right)}{4 \\frac{n\\pi}{b - a} J_m' \\left( 4\\zeta \\frac{n\\pi}{b - a} \\right)} \\sin n\\pi \\frac{\\xi - a}{b - a}\n$$\n$$\n- \\frac{n\\pi}{b - a} \\sum_{s=0}^{\\infty} Q_{mn}^{(s)}(\\rho) \\left[ e^{-(\\xi - a)y_{sm}} - (-1)^n e^{-(b - \\xi)y_{sm}} \\right]\n$$\n\nand, for $\\xi \\le a$ or $\\xi \\ge b$,\n$$\nP_{mn}(\\xi, \\rho) = - \\frac{n\\pi}{b - a} \\sum_{s=0}^{\\infty} Q_{mn}^{(s)}(\\rho) \\left[ e^{-|\\xi - a|y_{sm}} - (-1)^n e^{-|b - \\xi|y_{sm}} \\right]\n$$", "timestamp": "2026-07-22T05:26:34.344636+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 3, "total_pages": 18, "image_filename": "19930085964_p3.jpg", "text": "2\nNACA RM E9C25\n\nIn general, little similarity existed between the nodal patterns of the hollow blades and of the solid blade.\n\nINTRODUCTION\n\nThe turbine blades of jet-propulsion engines fabricated in this country and in England have been almost exclusively of the solid type. The use of such blades was prompted primarily by the expediency of the rapid development of aircraft turbines. The rapidly increasing production of high-temperature-alloy turbine blades has, however, created a serious problem of raw material supply, which may make a reduction in the quantity of strategic materials used in the fabrication of turbine blades essential. The use of hollow turbine blades provides a possibility of considerable savings in strategic materials. This savings may be effected not only by use of a smaller quantity of material but by use of other less strategic material that can be adequately cooled because the blades are hollow (reference 1). In addition, hollow turbine blades permit a considerable reduction in the weight of the turbine wheel because of the smaller stresses imposed on the wheel by the lighter blades.\n\nDuring the development of hollow turbine blades by several aircraft-engine manufacturers, failures were experienced that were traced to excessive vibrational stresses. A research program has therefore been undertaken at the NACA Lewis laboratory to investigate the stresses and the thermal conditions that limit the service lives of hollow blades. As part of the general program, the preliminary experimental investigation reported herein was made to determine the vibrational characteristics of several hollow turbine blades and to compare them with the vibrational characteristics determined for a solid blade of approximately the same external dimensions and airfoil contour. The effects of flutter are not considered in this report.\n\nAPPARATUS AND PROCEDURE\n\nDescription of turbine blades. - Turbine blades of the three types for which vibrational modes were determined are shown in figure 1. All blades were approximately 4 inches long and had approximately the same dimensions and airfoil contours. The trailing edges of the hollow blades were thicker than the trailing edge of the solid blade due to the minimum thickness being equal to twice the wall thickness.", "timestamp": "2026-07-22T05:26:37.513138+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 9, "total_pages": 42, "image_filename": "19930085914_p9.jpg", "text": "8\nNACA RM A9D25\n\nThe reduction of lift-curve slope and static longitudinal stability which occurred near a lift coefficient of 0.2 is more apparent than has been observed in other investigations. It is felt that this deviation was due to separation and consequent loss of lift at the wing tips, which, being well back of the moment reference, would have caused a reduction in static longitudinal stability. It is further believed that the rearward movement of the aerodynamic center (subsequent to the forward movement near a lift coefficient of 0.2) resulted from a chord-wise redistribution of load, due to separation, wherein the section centers of pressure moved aft. This phenomenon was noted in reference 7. The lift-curve slope increased throughout the range of lift coefficients in which this rearward aerodynamic-center movement occurred, as evidenced by figures 4(b) and 5(b). The abrupt forward movement of the aerodynamic center, beginning at a lift coefficient of about 0.5 to 0.6, probably resulted from wing stall beginning at the tips and progressing inward. Increasing the Mach number reduced the severity of this abrupt forward movement. The pitching-moment coefficient at zero lift was approximately -0.006 and changed very little throughout the Mach number range investigated.\n\nMinimum drag coefficient.— The effect of Mach number on minimum drag coefficient is shown in figure 6 for a Reynolds number of 2.0 million. The minimum drag coefficient increased from about 0.007 to 0.008 for a range of Mach numbers from 0.20 to 0.93. At zero angle of attack (the angle of attack for minimum drag) the outboard sections of the wing were at negative angles, which probably resulted in a greater increase of drag with Mach number than would be the case if all sections were at zero angle of attack.\n\nLift-curve slope.— The effect of Mach number on lift-curve slope at Reynolds numbers of 0.8 million and 2.0 million is presented in figure 7. Lift-curve slope increased from approximately 0.052 to 0.058 for the test Mach number range at a Reynolds number of 0.8 million and from approximately 0.049 to 0.055 at a Reynolds number of 2.0 million. In all cases, lift-curve slope was measured between lift coefficients of -0.1 and 0.1.\n\nLift-drag ratio.— The effect of Mach number on lift-drag ratio is presented in figures 8 and 9, which show the variation of lift-drag ratio with lift coefficient for various Mach numbers. The separation at the wing tips, the effects of which have been noted in the lift and moment data at a lift coefficient of about 0.20, is seen to manifest itself as an abrupt termination of the rise of lift-drag ratio with lift coefficient, which occurred at this same lift coefficient (about 0.2). The sharp reduction in lift-drag ratio is a result of the rapid increase of drag which occurred as the lift coefficient was increased above 0.20 or 0.25. A general decrease in maximum lift-drag ratio with increasing Mach number is seen in figure 10. The lift-drag ratios presented are of limited quantitative value however, because of the low degree of accuracy of the drag data at small angles of attack.", "timestamp": "2026-07-22T05:26:39.180616+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 43, "total_pages": 46, "image_filename": "19930085519_p43.jpg", "text": "42\nNACA RM No. L8K19\n\n$\\delta_F$\n(percent local\nwing chord)\nActuating\narms\n$\\nabla$ -3 filled-in\n$\\nabla$ -5 filled-in\n$\\nabla$ -7 filled-in\n$\\diamond$ -3 open\n$\\triangle$ -5 open\n$\\triangledown$ -7 open\n\nRolling-moment coefficient, $C_l$\nYawing-moment coefficient, $C_n$\n\n<!-- Image (129, 286, 821, 806) -->\n\n(b) Full-span slotted flap at $\\delta_F = 50^\\circ$.\nFigure 20.- Concluded.", "timestamp": "2026-07-22T05:26:40.434905+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 26, "total_pages": 29, "image_filename": "19930085879_p26.jpg", "text": "24\n\n[Figure: Graph plotting separation velocity against time. The y-axis is labeled \"Separation velocity, ft/sec\" ranging from 0 to 70. The x-axis is labeled \"Time from launching, sec\" ranging from 100 to 150. A curve rises from approximately (100, 18) to (145, 68). An annotation points to the initial part of the curve: \"Obtained from longitudinal accelerometers\". The NACA logo is in the bottom right corner of the graph area.]\n\nFigure 12.— Separation velocity plotted against time.\n\nNACA RM L9D11", "timestamp": "2026-07-22T05:26:41.143056+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 39, "total_pages": 53, "image_filename": "19930082542_p39.jpg", "text": "NACA TN No. 1867\n41\n\n[Figure: Micrograph showing grain structure at 100X magnification]\n[Figure: Micrograph showing grain structure at 1000X magnification]\n\n100X\n(e) $2100^\\circ$ F 1 hour.\n1000X\n\n[Figure: Micrograph showing grain structure at 100X magnification]\n[Figure: Micrograph showing grain structure at 1000X magnification]\n\n100X\n(f) $2150^\\circ$ F 1 hour\n1000X\n\nFigure 6.- Continued.\nNACA", "timestamp": "2026-07-22T05:26:42.637932+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 3, "total_pages": 51, "image_filename": "19930085962_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM A9E05\n\nthe subject of an investigation in the Ames 12-foot pressure wind\ntunnel. The aim of the investigation was to determine the aero-\ndynamic characteristics of such a wing plan form throughout the range\nof subsonic Mach numbers up to 0.94. Various phases of the investi-\ngation have been reported in references 1, 2, 3, and 4. Application\nof the data of these references to a wing-body-tail combination\nindicated the possibility of obtaining adequate longitudinal control\nthroughout the speed range by using either an all-movable stabilizer\nor an adjustable stabilizer with an elevator. To provide data on the\neffectiveness of an elevator applied to such a plan form, the model\npreviously tested as a wing has been tested as a horizontal tail with\na full-span elevator. The tests were conducted at a constant Reynolds\nnumber of 2,000,000 at Mach numbers from 0.20 to 0.94.\n\nCOEFFICIENTS AND SYMBOLS\n\nThe following coefficients are used in this report:\n\n| | | |\n| :--- | :--- | :--- |\n| $C_L$ | lift coefficient | $\\left(\\frac{\\text{lift}}{qS}\\right)$ |\n| $C_D$ | drag coefficient | $\\left(\\frac{\\text{drag}}{qS}\\right)$ |\n| $C_m$ | pitching-moment coefficient about quarter-chord point | |\n| | of the mean aerodynamic chord | $\\left(\\frac{\\text{pitching moment}}{qS\\bar{c}}\\right)$ |\n\nThe following symbols are used in this report:\n\n| | | |\n| :--- | :--- | :--- |\n| a | speed of sound, feet per second | |\n| b | twice model semispan, feet | |\n| c | local chord, feet | |\n| $\\bar{c}$ | mean aerodynamic chord, chord through centroid of | |\n| | semispan plan-form area | $\\left(\\frac{\\int_{0}^{b/2} c^2 dy}{\\int_{0}^{b/2} c dy}\\right)$, feet |\n| M | Mach number | $\\left(\\frac{V}{a}\\right)$ |\n| q | free-stream dynamic pressure | $\\left(\\frac{\\rho V^2}{2}\\right)$, pounds per square foot |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:26:44.841461+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 42, "total_pages": 58, "image_filename": "19930082617_p42.jpg", "text": "NACA TN 1962\n\n41\n\n[Figure: Side view of a buckled cylindrical structure with riveted panels and handwritten markings including “A”, “B”, “C”, “5”, “6”, “7”, etc.]\n\nFigure 30. - Side view of cylinder 72 after buckling. Near side or stringer 8.\n\nNACA", "timestamp": "2026-07-22T05:26:45.635869+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 12, "total_pages": 23, "image_filename": "19930085906_p12.jpg", "text": "NACA RM E9F20 CONFIDENTIAL 11\n\n[Figure: A circular mechanical assembly with concentric rings, radial supports, and multiple perforations. A scale marked \"INCHES\" is visible at the top. The outer ring has stamped text including \"FLAME HOLDER\" and \"NACA\". A NACA logo with identifier \"C-19910 10-24-47\" is in the lower right corner of the image.]\n\nFigure 3. - View of flame holder looking upstream from combustion chamber.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:26:50.679685+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 51, "total_pages": 66, "image_filename": "19930082914_p51.jpg", "text": "50\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:26:53.523104+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 10, "total_pages": 36, "image_filename": "19930085912_p10.jpg", "text": "8\nNACA RM No. E9C16\n\nwith a free-stream total temperature of $0^\\circ$ F and a liquid-water content of 1.0 gram per cubic meter. A procedure for obtaining a satisfactory orifice configuration is presented in the appendix.\n\nModel-air temperature distribution. - When the penetration equation of reference 2\n\n$$\n\\left(\\frac{l}{D_j}\\right)^{1.65} = 2.91 \\frac{\\rho_j V_j}{\\rho_1 V_1} \\sqrt{\\frac{s}{D_j}}\n$$\n\nis utilized, it is apparent that the penetration is a function of the product of the density and the velocity ratios because the geometric parameters are fixed. The total jet area $A_j$ and the inlet area $A_1$ of the model are fixed; hence the penetration becomes a function of the bleedback because\n\n$$\nl = f \\left( \\frac{\\rho_j V_j A_j}{\\rho_1 V_1 A_1} \\right) = f \\left( \\frac{\\dot{W}_g}{\\dot{W}_a} \\right)\n$$\n\nAfter a single value of bleedback (4.4 percent) corresponding to the most uniform temperature distribution is determined, it should be possible to maintain this optimum temperature distribution for a range of tunnel velocities if the bleedback is held constant. Because the air flow through the model varies nearly linearly with tunnel velocity, the variation of gas flow must also be nearly linear in order to maintain constant bleedback. For a fixed gas temperature, the gas flow varies linearly with gas pressure for a choked jet and nearly linearly with gas pressure for a high subsonic jet. The gas pressure and the tunnel velocity should therefore be linearly related for a fixed temperature distribution. The plenum-chamber gas pressure corresponding with optimum temperature distribution was determined for a value of gas temperature of $1000^\\circ$ F as a function of tunnel velocity, and the variation of gas pressure with tunnel velocity was found to be linear (fig. 7). For each tunnel velocity, the experimental value of bleedback proved to be the same (4.4 percent) and resulted in identical average air-temperature rises of $46^\\circ$ F with a maximum deviation of $6^\\circ$ F.\n\nThe effect of employing gas pressures other than the optimum (3850 lb/sq ft) is illustrated in figure 8 for a tunnel velocity of 290 feet per second and a plenum-chamber gas temperature of $1000^\\circ$ F. Lines of constant total-temperature ratio $T_x/T_{av}$ are", "timestamp": "2026-07-22T05:26:53.914044+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 28, "total_pages": 36, "image_filename": "19930085869_p28.jpg", "text": "```markdown\n26\n\nCONFIDENTIAL\n\nUpper limit\nIncreasing trim\nDecreasing trim\n\nElevator\ndeflection,\ndeg\n-15 -30\n\nTrim, deg\n16\n14\n12\n10\n8\n6\n4\n2\n0\n\nLower limit\n\n0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0 11.0 12.0 13.0 14.0\nSpeed coefficient, $C_V$\n\n(a) Center of gravity, $0.20 \\bar{c}$.\n\nFigure 9.— Variation of trim with speed coefficient.\nCONFIDENTIAL\n\nNACA RM L9D15\n\nNACA\n```", "timestamp": "2026-07-22T05:26:58.860284+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 33, "total_pages": 33, "image_filename": "19930085544_p33.jpg", "text": "32\nNACA RM No. L8K26\n\n$$ \\left(\\frac{d m_c}{d x}\\right)_{\\text{steady}} - .06 \\left(\\frac{d m_c}{d x}\\right)_{\\text{osc}} $$\n\n| | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| .006 | | | | | | | | | | |\n| .005 | | | | | | | | | | |\n| .004 | | | | | | | | | | |\n| .003 | | | | | | | | | | |\n| .002 | | | | | | | | | | |\n| .001 | | | | | | | | | | |\n| 0 | | | | | | | | | | |\n| | .2 | .3 | .4 | .5 | .6 | .7 | .8 | .9 | 1.0 | |\n| | | | | | | | | | | $m_c$ |\n| | | | | | | | | | | .00308 |\n| | | | | | | | | | | .00259 |\n| | | | | | | | | | | .00210 |\n\n---\nSteady state, compressible airfoil characteristics\nSteady state corrected for oscillations\nOscillating-airfoil theory; incompressible airfoil characteristics\n\nFigure 15.- Distribution of maximum bending-moment coefficient for propeller 4-(3.9)(07)-0345-B.\n$\\beta, 26^\\circ, J.1.2; C_{n_T}, 46, B, 2.$", "timestamp": "2026-07-22T05:27:01.105180+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 20, "total_pages": 37, "image_filename": "19930085889_p20.jpg", "text": "NACA RM L9F14\n19\n\nCONFIDENTIAL\n\n[Figure: Sketch of wing planforms with annotations:\nTop view: 50 percent c, 3.6°, 25 percent c, 7°, 0.278ft, L.E.\nMiddle view: 30°, 0.674ft, 25 percent c, 34.6°, 35°\nBottom view: 0.964ft, 45°, 25 percent c, 46.7°, 48.4°, Plane of section, Mounting point 1/4 mean aerodynamic chord, NACA logo]\n\nWing:\nArea, sq ft . . . . . . . . 2.25\nAspect ratio . . . . . . . . 4.0\nAirfoil section . . NACA 65A006\nSpan, ft . . . . . . . . . . 3.0\nMean aerodynamic\nchord, ft . . . . . . . . . 0.765\nTaper ratio . . . . . . . . . 0.60\nRoot chord, ft . . . . . . . 0.938\nTip chord, ft . . . . . . . . 0.563\n\nAileron:\nType . .True contour, sealed gap\nChord, percent c . . . . . . . 20\nSpan, percent b/2 . . . . . . 40\nInboard station,\npercent b/2 . . . . . . . . . 55\nOutboard station,\npercent b/2 . . . . . . . . . 95\n\nCONFIDENTIAL\n\nFigure 2.— Sketch and dimensions of wings tested.", "timestamp": "2026-07-22T05:27:05.732403+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 24, "total_pages": 96, "image_filename": "19930085880_p24.jpg", "text": "22\nNACA RM No. L9C03\n\n[Figure: Graph showing Resistance (lb) vs. Speed (fps) for various Wetted areas (sq ft). The graph includes a small inset diagram of a triangular shape. The NACA logo is present at the bottom right of the graph.]\n\n(a) $\\tau = 4^\\circ$.\n\nFigure 11.- Aerodynamic drag of model 250D.", "timestamp": "2026-07-22T05:27:05.787810+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 44, "total_pages": 78, "image_filename": "19930082618_p44.jpg", "text": "```markdown\n42\n\n<!-- Image (89, 70, 887, 760) -->\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\nFigure 8.- Aerodynamic characteristics of the NACA 63$_{2}$-415 airfoil section, 24-inch chord.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:27:14.303084+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 44, "total_pages": 46, "image_filename": "19930085519_p44.jpg", "text": "NACA RM No. L8K19\n43\n\nGap between plug-\naileron lower edge\nand wing upper\nsurface\n\n$\\delta_p$\n(percent\nideal wing\nchord)\n\n$\\Delta$ -5 filled-in\n$\\Delta$ -7 filled-in\n$\\nabla$ -5 open\n$\\nabla$ -7 open\n\n[Figure: Graph plotting Rolling-moment coefficient, $C_l$ and Yawing-moment coefficient, $C_n$ against Angle of attack, $\\alpha$, deg. The graph contains multiple data series represented by triangles and inverted triangles. A NACA logo is visible within the graph area.]\n\nFigure 21.- The effect of a gap between the lower edge of the plug\naileron and the wing upper surface at large plug-aileron projections\non the rolling-moment and yawing-moment coefficients of the plug\naileron on the 42° sweptback wing. Faired plug-slot lower lip,\nflap retracted, plug-aileron actuating arms filled in.", "timestamp": "2026-07-22T05:27:24.223578+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 40, "total_pages": 53, "image_filename": "19930082542_p40.jpg", "text": "NACA TN No. 1867\n43\n\n[Figure: Micrographs showing material structure at different magnifications]\n\n100X\n1000X\n(g) $2200^\\circ$ F 1 hour.\n\n100X\n1000X\n(h) $2250^\\circ$ F $\\frac{1}{2}$ hour.\nFigure 6.- Continued.", "timestamp": "2026-07-22T05:27:24.441139+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 19, "total_pages": 31, "image_filename": "19930085881_p19.jpg", "text": "CONFIDENTIAL\n\n$10 \\times 10^6$\n\nReynolds number\n\n8\n\n6\n\n4\n\n2\n\n0\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\n\nMach number\n\nNACA\n\nFigure 3.- Variation of Reynolds number with Mach number for range of climatic conditions encountered during tests.\n\nCONFIDENTIAL\n\nNACA RM L9D12\n\n17", "timestamp": "2026-07-22T05:27:26.049684+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 27, "total_pages": 29, "image_filename": "19930085879_p27.jpg", "text": "NACA RM L9D11\n25\n\n.0019\n.0020\n.0021 1108\n.0022 1104\n.0023 1096\n.0024 1088\n.0025 1076\n\nAir density, $\\rho$, slugs/ft$^3$\nVelocity of sound, $V_c$, ft/sec\nModel flight time from launching, $t$, sec\n\n30\n25\n20\n15\n10\n5\n0\n\n$t_{nose}$\n$t_{rearbody}$\n$\\rho$\nSeparation\n$V_c$\n\n0 2000 4000 6000\nAltitude, $h$, ft\n\nNACA\n\nFigure 13.— Flight conditions of RM-11B.", "timestamp": "2026-07-22T05:27:28.456181+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 10, "total_pages": 42, "image_filename": "19930085914_p10.jpg", "text": "NACA RM A9D25\n\nThe curves presented for 0.4 Mach number at Reynolds numbers of 0.8 million and 2.0 million (figs. 8 and 9) do not correspond to the data of figures 4 and 5. The erroneously low drags obtained at these test conditions, attributed to malfunction of the balance, resulted in corresponding values of lift-drag ratio which were unreasonably high; consequently, the data were retaken. It was later discovered that one flap was deflected slightly during the reruns and the data indicated zero lift at zero angle of attack. The lift-drag ratios presented are from results of the reruns. The erroneous flap angle was very small and it is reasoned that this deflection would not affect the general variation of lift-drag ratio with lift coefficient, although the angle of attack for a given lift-drag ratio would be affected.\n\nAerodynamic center.— The variation of aerodynamic-center position with Mach number is presented in figure 11. Aerodynamic-center locations were obtained from the linear portions of the moment curves through zero lift; consequently, they are significant only for that limited range. A small rearward movement was noted from approximately 41 to 45 percent of the mean aerodynamic chord as the Mach number increased from 0.20 to 0.93 at Reynolds numbers of 0.8 and 2.0 million.\n\nEffects of Reynolds Number\n\nGeneral aerodynamic characteristics.— General aerodynamic characteristics of the wing-fuselage combination are presented in figure 12 for several Reynolds numbers from 0.8 million to 9.0 million for a Mach number of 0.20. Increasing the Reynolds number reduced the drag at positive lift coefficients above about 0.2 (fig.12(a)). The lift data (fig. 12(b)) indicate that at the higher Reynolds numbers the slight reduction of lift-curve slope due to separation at the tips was reduced in magnitude and delayed to a higher lift coefficient. At a Reynolds number of 9.0 million this reduction began at a lift coefficient of about 0.35. However, the pitching moments at a Reynolds number of 9.0 million (fig. 12(c)) show very little movement of the aerodynamic center from a lift coefficient of -0.1 to a lift coefficient of 0.55, the highest value obtained at this Reynolds number. These data indicate that certain important effects of boundary-layer separation which are evident from tests of highly swept-back wings at low Reynolds may not be present under full-scale conditions.\n\nMinimum drag coefficient.— The variation of minimum drag coefficient with Reynolds number for a Mach number of 0.2 may be seen in figure 6. A gradual increase is noted from approximately 0.007 at a Reynolds number of 2.0 million to 0.010 at 9.0 million.\n\nLift-curve slope.— Variation of lift-curve slope with Reynolds number is shown in figure 7 for a Mach number of 0.2. The lift-curve slope decreased gradually from 0.051 at a Reynolds number of 0.8 million to 0.046 at 9.0 million.", "timestamp": "2026-07-22T05:27:28.620460+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 52, "total_pages": 66, "image_filename": "19930082914_p52.jpg", "text": "NACA TN No. 1857\n51\n\n[Figure: A black and white image showing a pattern of wavy, horizontal lines. The lines are more distorted and irregular in the upper portion of the image and become more regular and parallel in the lower portion. There are small, dark, triangular shapes at the bottom left and bottom right corners of the image.]\n\n(c) 3 to $4\\frac{1}{2}$ inches from nozzle.\nFigure 11.— Continued.", "timestamp": "2026-07-22T05:27:31.591418+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 61, "total_pages": 99, "image_filename": "19930082511_p61.jpg", "text": "NACA TN No. 1826\n59\n\nwhere\n\n$$Q_{mn}^{(s)} = \\frac{J_m(\\rho y_{sm}) y_{sm}}{\\left[ y_{sm}^2 + \\left( \\frac{n\\pi}{b-a} \\right)^2 \\right] (m^2 - y_{sm}^2) J_m(y_{sm})} \\quad (n \\neq 0)$$\n\nThe calculations for $n = 0$, which follow similar lines, are not given here. The final formulas are, for $a \\le \\xi < b$,\n\n$$P_{m0}(\\xi, \\rho) = \\frac{J_m\\left( \\frac{1}{2} \\rho \\frac{\\pi}{b-a} \\right)}{\\frac{1}{2(b-a)} J_m'\\left( \\frac{1}{2} \\frac{\\pi}{b-a} \\right)} \\sin \\frac{\\pi}{2} \\frac{\\xi-a}{b-a}$$\n\n$$+ \\sum_{s=0}^{\\infty} Q_{m0}^{(s)}(\\rho) \\left[ y_{sm} e^{-(b-\\xi)y_{sm}} - \\frac{\\pi}{2(b-a)} e^{-(\\xi-a)y_{sm}} \\right]$$\n\nfor $\\xi \\le a$,\n\n$$P_{m0}(\\xi, \\rho) = \\sum_{s=0}^{\\infty} Q_{m0}^{(s)}(\\rho) \\left[ y_{sm} e^{-(b-\\xi)y_{sm}} - \\frac{\\pi}{2(b-a)} e^{-(a-\\xi)y_{sm}} \\right]$$\n\nand, for $\\xi > b$,\n\n$$P_{m0}(\\xi, \\rho) = - \\sum_{s=0}^{\\infty} Q_{m0}^{(s)}(\\rho) \\left[ y_{sm} e^{-(\\xi-b)y_{sm}} + \\frac{\\pi}{2(b-a)} e^{-(\\xi-a)y_{sm}} \\right]$$\n\nwhere\n\n$$Q_{m0}^{(s)}(\\rho) = \\frac{J_m(\\rho y_{sm}) y_{sm}}{\\left\\{ y_{sm}^2 + \\left[ \\frac{\\pi}{2(b-a)} \\right]^2 \\right\\} (m^2 - y_{sm}^2) J_m(y_{sm})}$$\n\nEvaluation for $m = 0$\n\nThe evaluation of $P_{mn}(\\xi, \\rho)$ for $m = 0$ proceeds essentially as before with the difference that the contour of integration must avoid the origin. For $n = 0$, the final formulas are for $a \\le \\xi < b$,", "timestamp": "2026-07-22T05:27:33.208110+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 4, "total_pages": 18, "image_filename": "19930085964_p4.jpg", "text": "NACA RM E9G25\n\nThe two hollow blades of elementary design for which data are presented were chosen principally because of the ease with which they could be procured. Hollow-blade type A was plain with neither damping nor stiffening features and represented the most simple form of hollow blade of the aerodynamic design investigated having wall-thickness taper. The blade-wall thickness varied from 0.074 inch at the base to 0.024 inch at the tip. The material used in the fabrication of the blade was N-155 alloy sheet, which was formed into a cylinder and welded at the butting edges. Wall-thickness taper was then obtained by mounting the cylinder on a mandrel and machining it. The tube was filled with a hard wax and pressed between dies to the finished airfoil contour. The vibrational modes of one blade of type A were obtained.\n\nBlades of type B were similar to type A except that two stiffening features were incorporated (fig. 1). Three rivets were installed along the trailing edge of each blade to prevent relative motion of the sides of the blade and to supply additional damping. Also, the thickness of the blade wall was increased to 0.055 inch over a portion of the blade extending from the tip to a peripheral line approximately 1/4 inch below the tip. This thickening made the tip section more rigid and, consequently, acted to reduce relative motion of the sides of the blade. The wall thickness was 0.062 inch at the base and 0.024 inch at the tip below the thickened band. Vibrational modes of four blades of this type, designated B₁, B₂, B₃, and B₄, were obtained.\n\nBlade type C (fig. 1) was a solid Vitallium blade fabricated by precision casting. The vibrational modes of one blade of this type were obtained.\n\nExcitation. - Resonant vibration of the turbine blades was produced by the apparatus shown in figure 2. A partly split turbine wheel with fir-tree serrations was used to hold the turbine blades for individual investigation. A turbine blade was inserted in the fir-tree serration through which the slit in the wheel passed and then the two portions of the turbine wheel were tightly clamped together by bolts. The wheel was suspended near a speaker-type exciter, which was used to produce the various modes of vibration of the blades, and was connected to the moving coil of the exciter with a metal rod. The frequency of excitation was varied by a variable-beat-frequency oscillator. The probe of a crystal-type pickup was drawn back and forth across the blade at each of the resonant conditions. The output of the crystal pickup was connected to the Y-axis of a cathode-ray oscilloscope and a small proportion of the output of the beat-frequency oscillator connected to the X-axis. The nodal patterns of each mode were accurately plotted by", "timestamp": "2026-07-22T05:27:35.725367+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 29, "total_pages": 36, "image_filename": "19930085869_p29.jpg", "text": "```markdown\nUpper limit\nIncreasing trim\nDecreasing trim\nCONFIDENTIAL\n\nElevator\ndeflection,\ndeg\n-30\n-15\n\nTrim, deg\n16\n14\n12\n10\n8\n6\n4\n2\n0\n\nLower limit\n\n0\n+5\n\nNACA\n\nSpeed coefficient, $C_V$\n0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0 11.0 12.0 13.0 14.0\n\n(b) Center of gravity, 0.30 $\\bar{c}$.\n\nFigure 9.— Continued.\nCONFIDENTIAL\n\nNACA RM L9D15\n27\n```", "timestamp": "2026-07-22T05:27:38.514790+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 4, "total_pages": 51, "image_filename": "19930085962_p4.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 3\n\nR Reynolds number $\\left(\\frac{\\rho V \\bar{c}}{\\mu}\\right)$\n\nS area of semispan tail, square feet\n\nV airspeed, feet per second\n\ny distance from plane of symmetry to any spanwise station, feet\n\n$\\alpha$ angle of attack of tail-chord plane, degrees\n\n$\\delta_e$ elevator deflection, positive to increase lift, degrees\n\n$\\mu$ viscosity of air, slugs per foot-second\n\n$\\rho$ mass density of air, slugs per cubic foot\n\n$C_{L_\\alpha}$ $\\left(\\frac{\\partial C_L}{\\partial \\alpha}\\right)_{\\delta=0^\\circ}$, measured at $\\alpha=0^\\circ$\n\n$C_{L_\\delta}^*$ $\\left(\\frac{C_{L_{\\delta=4^\\circ}} - C_{L_{\\delta=0^\\circ}}}{4}\\right)_{\\alpha=0^\\circ}$\n\n$a_\\delta$ $-\\left(\\frac{C_{L_\\delta}^*}{C_{L_\\alpha}}\\right)_{C_L=0}$\n\n(The subscripts outside the parenthesis indicate the factor held constant during the measurement of the parameters.)\n\nMODEL AND APPARATUS\n\nThe tests were conducted in the Ames 12-foot pressure wind tunnel. The horizontal tail with full-span elevator used in this investigation was the same model as that used in the tests reported in reference 1. The semispan model represented a tail of aspect ratio 4 and taper ratio 0.5. The 50-percent-chord line of the tail was normal to the plane of symmetry and the airfoil profile was a sharp-edged, symmetrical, faired double wedge with a thickness-chord ratio of 0.042. The constant-chord elevator had an area equal to 20 percent of the total semispan tail area. The unsealed gap between the elevator and the tail was 0.015 inch. Dimensions of the model are given in figure 1. The semispan model was mounted vertically in the tunnel as shown in figure 2. The leading-edge flap with which the model was equipped remained undeflected throughout the present series of tests and all surface roughness\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:27:39.309946+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 11, "total_pages": 36, "image_filename": "19930085912_p11.jpg", "text": "NACA RM No. E9C16\n\nindicated for plenum-chamber gas pressures of 3010, 3850, 5100, and 5840 pounds per square foot absolute. The use of a pressure lower than the optimum results in low temperature ratios at the center with increasing temperature ratios near the duct walls (fig. 8(a)). For the plenum-chamber gas pressures higher than the optimum, the region of highest temperature ratio occurred at the center of the model and decreasing temperature ratios were encountered as the distance from the center increased (figs. 8(c) and 8(d)). A comparison of figures 8(a) to 8(d) shows that increasing temperature ratios are obtained near the center of the model with increasing gas pressures. In addition, the temperature-ratio gradient becomes more severe with increasing pressure.\n\nThe effect of angle of attack on the temperature distribution was negligible and almost no change in distribution was obtained when the other factors were maintained constant.\n\nIncreasing the plenum-chamber gas temperature at a fixed angle of attack also had very little effect on the temperature distribution aside from increasing the temperature rise; nearly identical total-temperature-ratio contours were obtained at constant tunnel speed and plenum-chamber gas pressure.\n\nMass-flow loss. - A reduction in mass flow through the model occurred with increasing model-air total temperature at a fixed tunnel velocity, free-stream total temperature, and angle of attack. The decrease in mass flow with increasing model-air temperature for several values of tunnel velocity and an angle of attack of $0^\\circ$ is shown in figure 9. The decrease in mass flow is linearly related to the model-air temperature for a fixed tunnel velocity. The reduction in flow is due to the decrease in air density associated with increasing temperature, which indicates that the effect of compressibility and the changes in static pressure inside the model with bleedback have a negligible effect on mass flow. At a fixed tunnel velocity, the model therefore operates as a constant-volume machine.\n\nRam-pressure recovery. - The ram-pressure loss associated with the addition of heat by jets directed perpendicularly to a moving air stream consists of two components. The first component arises from the momentum pressure loss associated with changing the direction of the jets. The other component arises from the change in density of the air due to the addition of heat. The effect of the first component is illustrated in figure 5. Because the model is a constant-volume machine and the velocities through the model are sufficiently low that compressibility effects may be neglected, the ram-pressure loss due to", "timestamp": "2026-07-22T05:27:40.715320+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 27, "total_pages": 46, "image_filename": "19930085542_p27.jpg", "text": "4E\nNACA RM No. L5L29\n\n[Figure: A model aircraft mounted in a wind tunnel. The model is viewed from the front, showing its wings and fuselage. A ruler marked in inches is visible below the model. A label on the image reads \"NACA L-58010\".]\n\nFigure 7.- Model 10 mounted in tunnel. $A = 1.0$; $\\Lambda_{c/4} = 36.9^\\circ$; $\\lambda = 0.58$. Profile, NACA 0012.\n\n25", "timestamp": "2026-07-22T05:27:41.554136+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 21, "total_pages": 37, "image_filename": "19930085889_p21.jpg", "text": "20\nNACA RM L9F14\n\nCONFIDENTIAL\n\nMounting point\n$\\frac{1}{4}$ mean aerodynamic chord\n\n35 in.\n15 in.\n32.6°\n25 percent c\n\nA\nA\n5 in.\n\nElliptical nose\nSemi-major axis = 5 in.\nSemi-minor axis = $2\\frac{1}{4}$ in.\n\n$4\\frac{1}{2}$ in. Diam.\n\nSection A-A\nCONFIDENTIAL\n\n[Figure: NACA logo]\n\nFigure 3.- Sketch of the fuselage and 32.6° sweptback wing giving the principal dimensions of the fuselage.", "timestamp": "2026-07-22T05:27:46.085507+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 25, "total_pages": 96, "image_filename": "19930085880_p25.jpg", "text": "NACA RM No. L9C03\n23\n\n[Figure: A graph plotting Resistance (lb) against Speed (fps). The y-axis ranges from 0 to .30. The x-axis ranges from 0 to 35. There are multiple lines originating from (0,0) representing different Wetted areas (sq ft) from 0 to .30. An inset diagram shows a triangle shape.]\n\nResistance, lb\nSpeed , fps\nWetted area\n(sq ft)\n0\n.05\n.10\n.15\n.20\n.25\n.30\n\nNACA\n\n(b) $\\tau = 8^\\circ$.\nFigure 11.- Continued.", "timestamp": "2026-07-22T05:27:47.071716+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 1, "total_pages": 46, "image_filename": "19930085972_p1.jpg", "text": "NACA RM L9B18\n\n170\nCopy\nRM L9B18\n\nPaul G. Fournier\n\nNACA\n\nRESEARCH MEMORANDUM\n\nLOW-SPEED WIND-TUNNEL INVESTIGATION OF THE LONGITUDINAL\nSTABILITY CHARACTERISTICS OF A MODEL EQUIPPED\nWITH A VARIABLE-SWEEP WING\n\nBy\nCharles J. Donlan and William C. Sleeman, Jr.\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Laws,\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\nCLASSIFICATION CANCELLED\nBY AUTHORITY J. W. CROWLEY\nCHANGE #1637 DATE 12-7-53 T.C.F.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nMay 23, 1949", "timestamp": "2026-07-22T05:27:48.910920+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 45, "total_pages": 78, "image_filename": "19930082618_p45.jpg", "text": "```markdown\nNACA TN 1945\n\nSection lift coefficient, $c_l$\n2.8\n2.4\n2.0\n1.6\n1.2\n.8\n.4\n0\n\nMoment coefficient, $c_{m_{c/4}}$\n-.1\n-.2\n-.3\n-.4\n\nR\n$\\square$ 0.7 $\\times$ $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\triangleright$ 6.0\nFlagged symbols denote\nstandard roughness\n\n[NACA logo]\n\n-16 -8 0 8 16 0 0 0 0\nSection angle of attack, $\\alpha_{0L}$, deg\n\n(b) Section lift and pitching-moment characteristics of the NACA 63$_2$-415 airfoil section with a\n0.20c simulated split flap deflected 60$^\\circ$.\n\nFigure 8.- Continued.\n\n43\n```", "timestamp": "2026-07-22T05:27:53.589851+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 6, "total_pages": 42, "image_filename": "19930085938_p6.jpg", "text": "NACA RM No. L9B04\n\nthe water corresponding to this curve was applied by dead weights. The range of trim tested at any speed was selected from the stability tests as being the range of stable trim obtainable at that speed by the use of the elevators. The resistance selected at each speed was the lowest resistance obtained at that speed. The trims at high speed were arbitrarily limited to $12^\\circ$.\n\n### Static Transverse Stability\n\nThe static transverse stability was determined by inclining the model with the wing removed. The model was balanced at its normal gross load with weights located on the center line at the 0.30c location of the center of gravity and then moved outboard to apply an upsetting moment. The resultant angle of heel was measured as the angle between the plane of symmetry and the vertical.\n\n### RESULTS AND DISCUSSION\n\nThe exploratory tests made with a tail float (see fig. 1(c)) indicated that such a configuration operated in a range of trim which was lower than that obtained with the boom alone. Near the take-off speed the model trimmed up suddenly resulting in premature take-offs. Because the tail float was apparently clear at the start of the motion, this trimming up was thought to be the result of negative air pressures acting on the float bottom as it operated in the trough of water formed in the forebody wake. This hydrodynamic feature, coupled with the increase in air drag due to the float, caused interest to be centered on the hull with the boom alone.\n\n### Take-Off Stability and Trim\n\nThe center-of-gravity limits of stability for the two models are given in figure 8 as a plot of elevator deflections against center-of-gravity locations. The range of fixed elevator deflection for stable take-offs was large for both configurations. For the single-boom configuration, this range increased from $15^\\circ$ at 0.20c to $30^\\circ$ at 0.40c. The range of fixed elevator deflection for the twin-boom configuration was about $25^\\circ$ at 0.20c and $40^\\circ$ at 0.40c. At the maximum fixed elevator deflection of $-30^\\circ$ take-offs of both configurations were stable.\n\nThe region of lower-limit porpoising encountered with the lower elevator deflections is shown in figure 9 where the trim limits of stability for the two models are plotted against speed coefficient. The lower trim limits were the same for both configurations with the exception of a slight difference in the minimum speed at which lower-limit porpoising was first encountered. The maximum trim at which lower-limit porpoising appeared was high for both configurations. No upper-limit porpoising was encountered", "timestamp": "2026-07-22T05:28:00.879334+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 41, "total_pages": 53, "image_filename": "19930082542_p41.jpg", "text": "NACA TN No. 1867\n45\n\n[Figure: Micrograph showing a dense, textured structure with vertical striations and irregular shapes.]\n[Figure: Micrograph showing a lighter background with scattered dark spots and voids.]\n\n100X\n(i) $2300^\\circ$ F $\\frac{1}{2}$ hour.\n1000X\n[Logo: NACA]\n\nFigure 6.- Concluded.", "timestamp": "2026-07-22T05:28:05.690624+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 43, "total_pages": 58, "image_filename": "19930082617_p43.jpg", "text": "42\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:28:06.642705+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 20, "total_pages": 31, "image_filename": "19930085881_p20.jpg", "text": "CONFIDENTIAL\n\n$\\frac{pb}{2V}$\n\nMeasured\nCorrected to steady-roll conditions\n\nM\n\nNACA\n\nFigure 4.- Effect of moment of inertia about roll axis on measured variation of $\\frac{pb}{2V}$ with Mach number.\n\nCONFIDENTIAL\n\n18\n\nNACA RM L9D12", "timestamp": "2026-07-22T05:28:07.442342+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 53, "total_pages": 66, "image_filename": "19930082914_p53.jpg", "text": "52\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:28:13.739880+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 28, "total_pages": 29, "image_filename": "19930085879_p28.jpg", "text": "26\n\nDrag coefficient, $C_D$\n\n.26\nFlaps start opening\nActual drag coefficient from\naccelerometer records\n.24\n.22\nDrag coefficient expected\nwithout flaps\n.20\n.18\n900 800 700 600 500\nVelocity, ft/sec\n\nNACA\n\nFigure 14.- RM-11B drag coefficient of nose section plotted against velocity.\n\nNACA RM 19D11", "timestamp": "2026-07-22T05:28:17.944933+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 11, "total_pages": 42, "image_filename": "19930085914_p11.jpg", "text": "```markdown\n10\nNACA RM A9D25\n\nLift-drag ratio.- The effect of Reynolds number on lift-drag ratio is presented in figure 13 which shows the variation of lift-drag ratio with lift coefficient at a Mach number of 0.20. It should be noted that at the higher Reynolds numbers (5.0, 7.0, and 9.0 million) the steep drop in lift-drag ratio was delayed to slightly higher lift coefficients, and that, as a result, the lift-drag ratios were near their maximum values over a greater range of lift coefficients.\n\nAttempts were made to obtain an insight on the tip separation at a lift coefficient of 0.2 by employing surface roughness. Full-span roughness strips of 2-percent-chord width were alternately placed at the leading edge of the wing and centered on the 5-percent-chord line. The roughness was achieved by sprinkling carborundum particles (grit No. 180) on an adhesive agent brushed over the desired areas of the wing. The particles covered approximately 80 percent of the area of the strips. The effects of these strips are shown in figure 14. The maximum lift-drag ratio was reduced, probably largely as a result of the increased friction drag due to the increase in the extent of the turbulent boundary layer as a result of fixing transition. However, the alleviation of the premature arrest of the rise of lift-drag ratio with lift coefficient by use of the roughness at the leading edge would seem to indicate that the boundary layer separating at the tip was laminar.\n\nVariation of maximum lift-drag ratio with Reynolds number is presented in figure 10. A decrease occurred from approximately 15.8 at 2.0 million Reynolds number to approximately 14.5 at 6.5 million with a subsequent increase to 15.2 at 9.0 million.\n\nAerodynamic center.- The variation of aerodynamic-center location with Reynolds number is presented in figure 11. A slight and nearly linear forward movement of the aerodynamic center is noted from 41 percent of the mean aerodynamic chord at 0.8 million Reynolds number to 39 percent at 9.0 million.\n\nAerodynamic Characteristics of the Fuselage\n\nAerodynamic characteristics of the fuselage are presented in figure 15 for several Mach numbers for a Reynolds number of 2.0 million. Evidence of the effect of friction in the balance is noted in the discontinuous character of the drag data near zero angle of attack for the lower values of Mach number. Friction, which acted in opposite directions for positive and negative angles of attack, accounts for the asymmetry of the curves of drag-coefficient variation with angle of attack.\n\nEffects of Camber and Twist\n\nGeneral aerodynamic characteristics.- Characteristics of the wing alone are compared in figure 16 with those of a wing of identical plan\n```", "timestamp": "2026-07-22T05:28:20.014721+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 30, "total_pages": 36, "image_filename": "19930085869_p30.jpg", "text": "```markdown\nCONFIDENTIAL\n\nUpper limit\nIncreasing trim\nDecreasing trim\n\nElevator\ndeflection\ndeg\n-30\n-15\n\nTrim, deg\n\n16\n14\n12\n10\n8\n6\n4\n2\n0\n\nLower limit\n\n0\n1.0\n2.0\n3.0\n4.0\n5.0\n6.0\n7.0\n8.0\n9.0\n10.0\n11.0\n12.0\n13.0\n14.0\n\nSpeed coefficient, $C_V$\n\n(c) Center of gravity, $0.40 \\bar{c}$.\n\nFigure 9.— Concluded.\nCONFIDENTIAL\n\nNACA\nNACA RM L9D15\n28\n```", "timestamp": "2026-07-22T05:28:21.701695+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 5, "total_pages": 18, "image_filename": "19930085964_p5.jpg", "text": "4\nNACA RM E9G25\n\nobservation of the amplitudes and phase relations of the traces on the oscilloscope. When sufficient points had been plotted on the blade to determine the nodal pattern, lines were drawn to connect the points. This procedure was followed for each side of each blade.\n\nRESULTS AND DISCUSSION\n\nSignificance of nodal patterns. - The vibrational modes of the hollow blades and of the solid blade are shown in figures 3 to 8. The exciting frequency in cycles per second is given below each nodal pattern. The solid lines in these figures represent the node lines on the concave side of the blade, and the dashed lines represent the node lines on the convex side of the blade. The node lines represent the locus of points of minimum amplitude, and by study of these lines the manner in which the specimen is vibrating may be determined. In some of the simple modes the locations of the maximum stress may be inferred from these patterns, although the node lines do not necessarily represent the locations of either the minimum or the maximum stresses.\n\nVibrational modes of blade A. - The nodal patterns determined for the plain hollow blade A are shown in figure 3. The lowest mode, excited at a frequency of 845 cycles per second (fig. 3(a)), has a node line running from the tip toward the base of the blade. Although this mode has the lowest frequency, which appears to be a beam vibration, it cannot be designated fundamental bending. In addition, no other detectable mode could be called fundamental bending.\n\nThe first five vibrational modes, figures 3(a) to 3(e), are representative of beam-type vibrations having similar nodal patterns on both sides of the blade. The modes observed at frequencies above 2590 cycles per second have, in general, dissimilar nodal patterns on the two sides of the blade. In these modes, the patterns become more complex at the higher frequencies and are representative of plate-type vibrations in which the blade vibrates as two plates fastened together at two sides, fixed at a third side, and free at the fourth side. Because the leading and trailing edges do not represent discontinuities, the node lines running to these edges do not stop but continue to the other side of the blade or travel along the edge to the base or the tip. The base and the tip, however, are discontinuities and node lines appearing on one side at the base or the tip do not necessarily appear on the opposite side. Although the node lines are generally discontinuous, a few are continuous. An example of a continuous node line is shown in figure 3(k) where a", "timestamp": "2026-07-22T05:28:22.939991+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 5, "total_pages": 51, "image_filename": "19930085962_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM No. A9E05\n\nassociated with the leading-edge flap and its angle brackets was minimized by sealing the gap and smoothing the surface. The elevator was attached to the tail by hinges and rigidly held in position by steel angle plates. Angular distortion of the elevator due to aero-dynamic loads was negligible.\n\nCORRECTIONS TO DATA\n\nThe data of this investigation have been corrected for tunnel-wall interference, constriction due to the tunnel walls, and model-support tare forces. The method of reference 5 was used in correcting the data for tunnel-wall interference. The following corrections were added:\n\n$$\n\\Delta\\alpha = 0.363 \\text{ C}_L\n$$\n\n$$\n\\Delta\\text{C}_D = 0.0056 \\text{ C}_L^2\n$$\n\n$$\n\\Delta\\text{C}_m = 0\n$$\n\nCorrections to the data for constriction effects of the tunnel walls have been evaluated by the method of reference 6. The magnitude of these corrections as applied to Mach number and to dynamic pressure (measured with the tunnel empty) is illustrated by the following table:\n\n| Corrected Mach number | Uncorrected Mach number | $\\frac{\\text{q}_{\\text{corrected}}}{\\text{q}_{\\text{uncorrected}}}^1$ |\n| :---: | :---: | :---: |\n| 0.94 | 0.931 | 1.004 |\n| .92 | .915 | 1.003 |\n| .90 | .897 | 1.002 |\n| .87 | .868 | 1.002 |\n| .85 | .848 | 1.002 |\n| .80 | .799 | 1.001 |\n| .70 | .700 | 1.001 |\n| .50 | .500 | 1.001 |\n| .20 | .200 | 1.001 |\n\n$^1$The values of $\\text{q}_{\\text{corrected}}/\\text{q}_{\\text{uncorrected}}$ which were presented in references 1, 3, and 4, were erroneously tabulated and were not the values used in the reduction of the data. The correct values are presented herein and are the values which were actually applied to all test data on this model.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:28:28.518113+00:00"}