PDF/ocr_dataset_1069.jsonl

{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 3, "total_pages": 34, "image_filename": "19930085913_p3.jpg", "text": "2\nNACA RM L9F24\n\npositions on the leading edge and midchord line. The investigation covered sweepback angles of $0^\\circ$, $45^\\circ$, and $60^\\circ$.\n\nSeveral analytical methods have been devised for calculating the flutter speed of an unswept wing carrying an arbitrarily placed weight. The methods of reference 1, which treats a uniform unswept wing by a differential-equation method, and of reference 2, which treats a general unswept wing by using chosen modes, were appraised with the aid of the experimental data presented in reference 3. Reference 4, which presents a general analytical method for swept wings, does not explicitly develop the procedures for including a concentrated weight and comparison is made with experiment for uniform wings only. The present paper furnishes experimental data that can be used to examine methods for predicting the flutter speed of a weighted sweptback wing.\n\nBecause of the importance of the vibration characteristics in a flutter analysis, the nodal-line patterns associated with the second and third natural frequencies of the models at zero airspeed are presented. This information may serve to check the method used in a flutter analysis for analytically obtaining the coupled modes of vibration of a wing at zero airspeed.\n\nThe models were made from uniform thin sheet metal with their leading edges rounded off and, if destroyed by flutter, they could easily be reproduced. The models tested were practically the same in characteristics except as changed by sweepback.\n\nEssentially one weight was used throughout the series of tests. This weight approximately simulated the mass characteristics of an engine. The results of tests as presented in this paper may be regarded qualitatively for the effects investigated and furthermore used quantitatively for comparison with subsequent analyses.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| W | weight of wing model, pounds |\n| $W_w$ | weight of concentrated weight, pounds |\n| $l$ | length of midchord line, feet |\n| b | half-chord of wing model measured perpendicular to midchord line, feet |\n| t | thickness of wing section, inches |", "timestamp": "2026-07-22T06:33:05.549357+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 50, "total_pages": 65, "image_filename": "19930082546_p50.jpg", "text": "NACA TN No. 1870\n49\n\n<!-- Image (129, 128, 842, 713) -->\n\n(a) $M_t = 0.75$.\nFigure 12.- Relative amplitudes of four harmonics of NACA 4-(5)(08)-03 propeller. $B = 2$; $\\beta_{0.75} = 10^\\circ$; $\\frac{a}{D} = 0.083$.", "timestamp": "2026-07-22T06:33:06.611230+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 46, "total_pages": 46, "image_filename": "19930082613_p46.jpg", "text": "```markdown\nNACA TN 1938\n45\n\nCool zone\n\nMetal\nexpands\n\nMetal\nexpands\n\nAir-intake hole\n\nLouver\n\nMetal\nexpands\n\nCool zone\n\nCool zone\n\nCool zone\n\nFigure 16. - Hot and cool zones about louver. Dashed lines indicate approximate location\nof buckles.\n\nRigidly supported\nend of flap\n\nTensile stresses in upper\nportion of louver flap (cool portion)\n\nAir\n\nHot gases\n\nCompressive stresses in lower portion\nof louver flap (hot portion)\n\n(a) Normal position\nof louver flap.\n\n(b) Conditions during\noperation.\n\nFigure 17. - Conditions in louver flaps.\n\nNACA-Langley - 10-4-49 - 850\n```", "timestamp": "2026-07-22T06:33:12.687468+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 27, "total_pages": 32, "image_filename": "19930085847_p27.jpg", "text": "NACA RM A9D04\nCONFIDENTIAL\n25\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (273, 108, 761, 854) -->\n\n(h) M = 0.83 ($C_L = 0.16$).\n\nFigure 7.- Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:33:13.106291+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 35, "total_pages": 47, "image_filename": "19930083221_p35.jpg", "text": "```markdown\nNACA TN No. 1824\n33\n\n$$\n\\left( \\frac{\\partial \\varphi}{\\partial x} \\right)_o = \\frac{V_o \\lambda \\cos \\psi}{\\pi \\tan \\psi} \\ln \\left[ - \\frac{(y \\tan \\psi - x - \\frac{1}{2}c_o \\sec \\psi)(y \\tan \\psi - x + \\frac{1}{2}c_o \\sec \\psi)}{(y \\tan \\psi - x)^2} \\right]\n$$\n\nand, using the definition for pressure coefficient, $C_p = - \\frac{2u}{V_o}$, this becomes\n\n$$\nC_p = - \\frac{2\\lambda \\cos \\psi}{\\pi \\tan \\psi} \\ln \\left[ \\left( \\frac{\\frac{1}{2}c_o}{\\sqrt{y \\sin \\psi - x \\cos \\psi}} \\right)^2 - 1 \\right] \\quad (55)\n$$\n\nEquation (55) can be derived by entirely different methods. Perhaps the most direct of these alternative derivations is the one introduced by R. T. Jones in reference 18. The general statement used in that report is that the component of translational velocity of a cylindrical body in the direction of its long axis has no effect on the motion of a frictionless fluid. Hence, the pressures over the wing shown in figure 13 are the same as those over a wing moving normal to a free stream with a velocity $V_o \\cos \\psi$. Using the Prandtl-Glauert correction to the thin airfoil solution of a two-dimensional, diamond-shaped, nonlifting section exposed to a free stream with velocity $V_o \\cos \\psi$, one obtains, for $M_o \\cos \\psi < 1$,\n\n$$\n\\varphi = - \\frac{1}{2\\pi} \\int_{-c_o/2}^{c_o/2} \\frac{w_o}{\\sqrt{1 - M_o^2 \\cos^2 \\psi}} \\ln[(x' - x_1')^2 + z^2] dx_1' \\quad (56)\n$$\n\nwhere $w_o$ is the vertical induced velocity on the upper side of the $z = 0$ plane and $x'$ is measured normal to the leading edge. If this solution is referred to the axial system of figure 13 by the transformation\n\n$$\nx' = x \\cos \\psi - y \\sin \\psi\n$$\n\nand the integration is performed after taking the partial derivative with respect to $x$, the resultant expression for pressure coefficient is\n\n$$\nC_p = - \\frac{2\\lambda \\cos^2 \\psi}{\\pi \\sqrt{1 - M_o^2 \\cos^2 \\psi}} \\ln \\left[ \\left( \\frac{\\frac{1}{2}c_o}{x \\cos \\psi - y \\sin \\psi} \\right)^2 - 1 \\right] \\quad (57)\n$$\n```", "timestamp": "2026-07-22T06:33:14.836026+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 28, "total_pages": 60, "image_filename": "19930085862_p28.jpg", "text": "26\nNACA RM No. L9A07\n\n<!-- Image (213, 195, 692, 410) -->\n\n$\\delta_e$\n(deg)\n$\\nabla$ -15\n$\\circ$ 0\n$\\triangle$ 15\n\n<!-- Image (213, 535, 692, 750) -->\n\n(c) $C_L$ and $C_m$ against $\\alpha$.\nFigure 6.- Concluded.", "timestamp": "2026-07-22T06:33:18.164857+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 5, "total_pages": 52, "image_filename": "19930085911_p5.jpg", "text": "4 CONFIDENTIAL NACA RM E9F22\n\nseven magnesium flares (fig. 6). A similar flame holder was investigated in a wind tunnel (reference 2) over approximately the same range of combustion-chamber-inlet conditions encountered in the flight investigation of ram-jet unit A-5.\n\nINSTRUMENTATION\n\nA portable radar-tracking unit, type SCR-584, with both optical and autotracking facilities was used to obtain a time history of the position of the ram-jet unit relative to the ground during flight. The telemetering-receiver antenna was directionally controlled by the radar tracker in order to maintain the strongest possible telemetering signal. A plotter was synchronized with the radar tracker and automatically charted the course of the ram-jet units.\n\nThe eight-channel telemetering equipment in the ram-jet unit transmitted the following data to two ground receiving recorders:\n\n1. Axial net acceleration\n2. Pressure drop across fuel-flow orifice\n3. Free-stream total pressure\n4. Free-stream static pressure\n5. Static pressure in diffuser, $4\\frac{5}{8}$ inches downstream of diffuser inlet, station 2 (fig. 3)\n6. Dynamic pressure in diffuser, 65 inches downstream of diffuser inlet, station 3 (fig. 3)\n7. Total pressure at diffuser outlet, station 4 (fig. 3)\n8. Static pressure at engine outlet, exhaust-nozzle outlet, station 7 (fig. 3)\n\nThe axial net acceleration (total acceleration minus the component due to gravity) was measured by a cantilever-beam-type accelerometer, with the beam fixed at one end, weighted at the other, and free to move in the direction of the axis of the ram jet. The force of gravity did not affect the deflection of the beam because it acted equally upon both the ram jet and the accelerometer. The fuel-flow orifice was calibrated to determine the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:33:19.625198+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 37, "total_pages": 72, "image_filename": "19930085491_p37.jpg", "text": "36 CONFIDENTIAL NACA RM No. A8J04\n\n18. Nitzberg, Gerald E., and Crandall, Stewart: Some Fundamental Similarities Between Boundary-Layer Flow at Transonic and Low Speeds. NACA TN No. 1623, 1948.\n\n19. Graham, Donald J., Nitzberg, Gerald E., and Olson, Robert N.: A Systematic Investigation of Pressure Distributions at High Speeds Over Five Representative NACA Low-Drag and Conventional Airfoil Sections. NACA RM No. A7B04, 1947.\n\n20. Frick, Charles W., and Boyd, John W.: Investigation at Supersonic Speed (M=1.53) of the Pressure Distribution Over a $63^\\circ$ Swept Airfoil of Biconvex Section at Zero Lift. NACA RM No. A8C22, 1948.\n\n21. von Doenhoff, Albert E., and Tetervin, Neal: Investigation of the Variation of Lift Coefficient with Reynolds Number at a Moderate Angle of Attack on a Low-Drag Airfoil. NACA CB, Nov. 1942.\n\n22. Theodorsen, Theodore and Regier, Arthur: Experiments on Drag of Revolving Disks, Cylinders and Streamline Rods at High Speeds. NACA Rep. No. 793, 1944.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:33:21.422215+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 67, "total_pages": 98, "image_filename": "19930086073_p67.jpg", "text": "NACA RM A57E04\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\nLeft aileron deflection, $\\delta_{a_2}$, deg\n\n+10.8 0 -10.8\n\n(b) $C_L$ vs $C_D$.\n\nFigure 14.—Continued.\n\n65", "timestamp": "2026-07-22T06:33:23.208533+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 28, "total_pages": 32, "image_filename": "19930085847_p28.jpg", "text": "26\nCONFIDENTIAL\nNACA RM A9D04\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (218, 101, 650, 803) -->\n\n(a) M = 0.76 ($C_L = 0.16$).\n\nFigure 8- Curves of measured chordwise and thickness-wise pressure distributions over the test panel at selected Mach numbers in the test range. With suction, full span slot.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:33:32.130057+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 1, "total_pages": 22, "image_filename": "19930085922_p1.jpg", "text": "NACA RM No. L9C23\n\nRM No. L9C23\n\n[Figure: NACA logo with wings]\n\nRESEARCH MEMORANDUM\n\nHIGH-SUBSONIC DAMPING-IN-ROLL CHARACTERISTICS OF A WING  \nWITH THE QUARTER-CHORD LINE SWEPT BACK $35^\\circ$ AND  \nWITH ASPECT RATIO 3 AND TAPER RATIO 0.6\n\nBy  \nBoyd C. Myers, II and Richard E. Kuhn\n\nLangley Aeronautical Laboratory  \nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE  \nFOR AERONAUTICS  \nWASHINGTON  \nMay 10, 1949", "timestamp": "2026-07-22T06:33:33.336512+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 29, "total_pages": 60, "image_filename": "19930085862_p29.jpg", "text": "```markdown\nNACA RM No. L9A07\n27\n\n$C_{Na}$\n$\\delta_a$ (deg)\n-25\n-20\n-15\n-10\n-5\n0\n5\n10\n15\n20\n25\n\n$C_n$\n\n$C_l$\nNACA\n\n$\\alpha$, deg\n\n(a) $C_l$, $C_n$, and $C_{Na}$ against $\\alpha$.\n\nFigure 7.— Aileron characteristics of wing with drooped-nose and split flaps and fences.\n```", "timestamp": "2026-07-22T06:33:38.887830+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 35, "total_pages": 62, "image_filename": "19930082918_p35.jpg", "text": "34\nNACA TN 1940\n\nTABLE 6\nROOM-TEMPERATURE PHYSICAL CONSTANTS FOR LOW-CARBON\nM-155 ALLOY AND CONSTITUENTS\n\n| Metal | Crystallographic system | Unit cell size ($\\AA$) | Closest approach of atoms ($\\AA$) | Atomic fraction in low-carbon M-155 (a) |\n| :--- | :--- | :--- | :--- | :--- |\n| Iron$^b$ | Body-centered cubic | 2.8606 | 2.476 | 0.322 |\n| Chromium$^b$ | Body-centered cubic | 2.8787 | 2.493 | .233 |\n| Nickel$^b$ | Face-centered cubic | 2.5167 | 2.486 | .185 |\n| Cobalt$^d$ | Close-packed hexagonal | 2.502 to 3.066 | 2.494 | .192 |\n| | Face-centered cubic | 2.540 | 2.502 | |\n| Manganese$^b$ | Cubic (complex) | 8.894 | 2.24 | .0171 |\n| Silicon$^b$ | Diamond cubic | 5.4173 | 2.346 | .0085 |\n| Carbon$^b$ | Hexagonal | 2.4564 to 6.6906 | 1.42 | .00618 |\n| Nitrogen | | | | .00532 |\n| Tungsten$^b$ | Body-centered cubic | 3.1585 | 2.734 | .00813 |\n| Columbium$^b$ | Body-centered cubic | 3.2941 | 2.853 | .00521 |\n| Molybdenum | Body-centered cubic | 3.140 | 2.720 | .0174 |\n| Low-carbon M-155 | Face-centered cubic | 3.560 | 2.54 | 1.000 |\n\n$^a$Converted from data reported by manufacturer.\n$^b$Reference 13.\n$^c$By difference.\n$^d$Reference 14.\n$^e$Solution-treated.\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T06:33:40.158004+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 33, "total_pages": 46, "image_filename": "19930085548_p33.jpg", "text": "```markdown\n32\nNACA RM No. E8L30\n\n<!-- Image (229, 109, 801, 835) -->\n\nFigure 10. - Effect of over-all fuel-air ratio on\nperformance of experimental cylinder with two-stroke\nand four-stroke cycle operation. Inlet-manifold\ntemperature, 400° F; two-stroke cycle; compression\nratio, 5.35; inlet-manifold pressure, 80 pounds per\nsquare inch absolute; four-stroke cycle; compression\nratio, 4.6; inlet-manifold pressure, 100 pounds per\nsquare inch absolute.\n\n1077\n```", "timestamp": "2026-07-22T06:33:40.787322+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 1, "total_pages": 47, "image_filename": "19930085919_p1.jpg", "text": "Copy No. 178\nRM No. A9C21\n\nNACA RM No. A9C21\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION\n\nEMPLOYING A WING SWEPT BACK 63°. - EFFECTS\n\nOF SPLIT FLAPS, ELEVONS, AND LEADING-\n\nEDGE DEVICES AT LOW SPEED\n\nBy Edward J. Hopkins\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: NACA RESEARCH ABSTRACT NO. 113\nEFFECTIVE DATE: MARCH 19, 1957\nWHL\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nMay 19, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:33:41.037706+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 36, "total_pages": 47, "image_filename": "19930083221_p36.jpg", "text": "34\nNACA TN No. 1824\n\nAt sonic speed this equation reduces immediately to\n\n$$C_p = - \\frac{2\\lambda \\cos \\psi}{\\pi \\tan \\psi} \\ln \\left[ \\left( \\frac{\\frac{1}{2} c_0}{x \\cos \\psi - y \\sin \\psi} \\right)^2 - 1 \\right] \\quad (58)$$\n\nwhich is identical to equation (55). The result expressed by the two equations is, of course, not new. The significant point is that the same variation in pressure coefficient was obtained by two widely different avenues of approach and that the result obtained from the particular methods applicable to sonic speed theory is in agreement with that derived from more conventional analysis.\n\nLifting-surface solutions at $M_0 = 1$.— It should be mentioned at this point that Robinson and Young (reference 19) have shown by means of linearized theory that supersonic triangular wings and subsonic elliptical wings of the same aspect ratio have values of lift-curve slope which approach a common and finite limit as $M_0 = 1$. The present section of this report is concerned only with the study of lifting surfaces at a fixed sonic velocity but the results to be obtained are in agreement with the limiting values of reference 19.\n\nA further application of the results in this section can be made to the case of very low aspect ratio wings at arbitrary Mach numbers. This viewpoint of the theory was first presented by R. T. Jones in reference 20 and applied to triangular wings while in reference 21 extension was made to include pointed wings on slender bodies of revolution. This duality of interpretation, that is, to all aspect ratios at sonic speed or low aspect ratios at all Mach numbers, applies to all solutions of three-dimensional problems obtained from equation (13). In the subsequent analysis, attention will be confined to swept-back plan forms of lifting surfaces with pointed vertices and thus doublets will be used exclusively.\n\nIn application, the two types of boundary conditions to be considered are as follows:\n\n1. Boundary-value problem of the first kind, loading specified.— It is given that $\\Delta u_0 = u_u - u_l = 0$ over the xy plane except for the region occupied by the wing where $2u_u = -2u_l = \\Delta u_0 = f(x,y)$, the function being determined by the specified loading. Over all of the xy plane, the imposed conditions are $\\Delta w_0 = 0$.", "timestamp": "2026-07-22T06:33:41.822799+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 65, "total_pages": 114, "image_filename": "19930086061_p65.jpg", "text": "NACA RM L9J07\n61\n\nLeft semispan\nRight semispan\nUpper\nLower\n\n(c) $\\psi = 20^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\n\n(d) $\\psi = 35^\\circ$\n\nFigure 19.- Concluded.", "timestamp": "2026-07-22T06:33:42.399908+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 51, "total_pages": 65, "image_filename": "19930082546_p51.jpg", "text": "50\nNACA TN No. 1870\n\nPressure, p, dynes/cm$^2$\n\n| mB | |\n| :--- | :--- |\n| $\\circ$ | 2 |\n| $\\square$ | 4 |\n| $\\diamond$ | 6 |\n| $\\triangle$ | 8 |\n\nx/D\n\n(b) $M_t = 0.90$.\nFigure 12.— Continued.", "timestamp": "2026-07-22T06:33:43.349675+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 68, "total_pages": 98, "image_filename": "19930086073_p68.jpg", "text": "66\nNACA RM A9H04\n\n<!-- Image (176, 116, 804, 865) -->\n\n$$\n\\begin{array}{ccc}\n\\square & \\bigcirc & \\diamond \\\\\n+10.8 & 0 & -10.8\n\\end{array}\n$$\n\nLeft aileron deflection, $\\delta_{a_L}$, deg\n\n(c) $C_L$ vs $C_m$.\n\nFigure 14.— Continued.", "timestamp": "2026-07-22T06:33:44.073989+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 29, "total_pages": 32, "image_filename": "19930085847_p29.jpg", "text": "NACA RM A9D04\nCONFIDENTIAL\n27\n\nDrag-producing area\nThrust-producing area\n\n[Figure: Graph showing pressure distribution. Y-axis: y/c (upper and lower). X-axis: $\\Delta p/q$. Shaded regions indicate drag and thrust producing areas. A line indicates $P_{crit}$.]\n\n[Figure: Graph showing pressure distribution. Y-axis: $\\Delta p/q$. X-axis: $x/c$. Data points for upper and lower surfaces. Lines indicate $P_{crit}$ and $t_{max}$. NACA logo present.]\n\n(b) $M = 0.80$ ($C_L = 0.18$).\n\nFigure 8.- Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:33:45.735234+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 30, "total_pages": 60, "image_filename": "19930085862_p30.jpg", "text": "```markdown\n28\nNACA RM No. L9A07\n\n<!-- Image (125, 109, 733, 326) -->\n\n<!-- Image (125, 387, 733, 850) -->\n\n(b) $C_{h_a}$ and $P_R$ against $\\alpha$.\nFigure 7.— Continued.\n```", "timestamp": "2026-07-22T06:33:49.586288+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 38, "total_pages": 72, "image_filename": "19930085491_p38.jpg", "text": "```markdown\nTABLE I.- SUMMARY OF GEOMETRIC PROPERTIES OF WINGS\n\n| Config- | $\\Lambda_{L.E.}$ | A | S | $\\frac{c_t}{c_r}$ | $\\bar{c}$ | $\\bar{c}_g$ | t/c | h/c | $M_n$ | m |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| uration | (deg) | | (sq in.) | | (in.) | (in.) | (percent) | (percent) | ($M_o=1.53$) | |\n| WF-57 | 57.0 | 4.49 | 7.791 | 0.27 | 1.452 | 1.318 | 6.70 | 38.5 | 0.83 | 0.75 |\n| WF-60 | 60.4 | 4.03 | 7.809 | .24 | 1.543 | 1.391 | 6.40 | 39.2 | .76 | .66 |\n| WF-63 | 63.0 | 3.42 | 7.223 | .25 | 1.615 | 1.455 | 6.00 | 40.0 | .69 | .59 |\n| WF-67 | 67.0 | 2.71 | 7.344 | .26 | 1.868 | 1.646 | 5.30 | 41.5 | .60 | .49 |\n| WF-70 | 69.9 | 2.23 | 7.600 | .25 | 2.078 | 1.845 | 4.65 | 42.6 | .53 | .42 |\n\nNote: The aspect ratios and mean geometric chords are based on the wing area including that blanketed by the fuselage. The taper ratios and mean geometric chords neglect the slight rounding of the wing tips by assuming them to be straight lines parallel to the stream direction and tangent to the outermost true tip contour.\n\nCONFIDENTIAL\n\nNACA RM No. A8J04\n\nCONFIDENTIAL\n\nNACA\n\n37\n```", "timestamp": "2026-07-22T06:33:50.824629+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 2, "total_pages": 22, "image_filename": "19930085922_p2.jpg", "text": "1C\n\nNACA RM No. L9C23\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nHIGH-SUBSONIC DAMPING-IN-ROLL CHARACTERISTICS OF A WING\n\nWITH THE QUARTER-CHORD LINE SWEPT BACK $35^\\circ$ AND\n\nWITH ASPECT RATIO 3 AND TAPER RATIO 0.6\n\nBy Boyd C. Myers, II and Richard E. Kuhn\n\nSUMMARY\n\nAn investigation of the damping-in-roll characteristics of a $35^\\circ$ swept-back wing, with and without vertical fins, has been made through a Mach number range of 0.40 to 0.91 in the Langley high-speed 7- by 10-foot tunnel utilizing a free-to-roll technique.\n\nThe damping-in-roll coefficient increased in magnitude with Mach number in the manner indicated by theory and generally was found to increase in magnitude with angle of attack over the range tested. Vertical fins located at about the midspan station of each wing panel had little effect on the damping-in-roll characteristics of the wing but increased the aileron effectiveness in producing rolling moment.\n\nINTRODUCTION\n\nAn extensive investigation of the effects of compressibility on the damping-in-roll characteristics of various wing plan forms is being conducted in the Langley high-speed 7- by 10-foot tunnel. The $35^\\circ$ swept-back wing of aspect ratio 3 and taper ratio 0.6 employed in the present investigation was that used in the high-speed wind-tunnel investigation of the tailless model reported in reference 1. The present investigation was carried out on a free-rotating apparatus with the purpose of determining the effects of compressibility and angle of attack on the damping-in-roll characteristics of the wing, with and without the vertical fins that were used on the tailless model reported in reference 1.\n\nCOEFFICIENTS AND SYMBOLS\n\nA wing aspect ratio\n\na speed of sound, feet per second", "timestamp": "2026-07-22T06:33:51.029645+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 15, "total_pages": 33, "image_filename": "19930085900_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:33:52.466334+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 36, "total_pages": 62, "image_filename": "19930082918_p36.jpg", "text": "NACA TN 1940\n35\n\nTABLE 7\nELEMENTS OF ATOMIC RADIUS OVER 2.8A FOR POSSIBLE\nUSE IN LOW-CARBON N-155 ALLOY\n\n| Element | Crystallographic system | Unit cell size (A) | Closest approach of atoms (A) |\n| :--- | :--- | :--- | :--- |\n| Cerium | Face-centered cubic | 5.143 | 3.64 |\n| Zirconium | Close-packed hexagonal | 3.223 to 5.123 | 3.16 |\n| Titanium | Close-packed hexagonal | 2.953 to 4.729 | 2.91 |\n\n[Figure: NACA logo]\n\nTABLE 8\nELEMENTS WITH AN ATOMIC RADIUS OF APPROXIMATELY 2.8A\nFOR SUBSTITUTION OF TUNGSTEN, MOLYBDENUM,\nOR COLUMBIUM IN LOW-CARBON N-155 ALLOY\n\n| Element | Crystallographic system | Unit cell size (A) | Closest approach of atoms (A) |\n| :--- | :--- | :--- | :--- |\n| Aluminum | Face-centered cubic | 4.0408 | 2.856 |\n| Silver | Face-centered cubic | 4.2774 | 2.88 |\n| Tantalum | Body-centered cubic | 3.2959 | 2.85 |\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T06:33:55.493906+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 15, "total_pages": 26, "image_filename": "19930085890_p15.jpg", "text": "```markdown\n14\n\nHelium pressurizing system\n\nCounterbalanced\nweighing beam\n\nLiquid oxygen\nVacuum jacket\n\nWater\n\nPropellant valves\npacked in dry ice\n\nHelium\nFilters\n\nHelium\n\nCounterbalanced\nweighing beam\n\nHelium\nbleed\n\nChamber\npressure\n\nEngine\n\nPivoted\nthrust stand\n\nLiquid\nnitrogen\n\nWater\n\nDiborane\n\nAir-operated valve\n\nFirst-stage pressure\nregulator\n\nDome-loaded pressure\nregulator\n\nSolenoid valve\n\nCheck valve\n\nHand valve\n\nPrincipal and auxiliary liquid lines\n\nPrincipal and auxiliary gas lines\n\nGage\n\nCantilever beam fitted with strain gages\n\nNACA\n\nFigure 1. - Diagrammatic sketch of 100-pound-thrust rocket apparatus for liquid diborane and liquid oxygen.\n\nNACA RM No. E9C11\n```", "timestamp": "2026-07-22T06:33:55.734914+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 29, "total_pages": 31, "image_filename": "19930085859_p29.jpg", "text": "NACA RM No. L9B25\n\n$\\frac{q_{\\text{wake}}}{q}$\n\nM = 0.90 $\\alpha = 10^\\circ$ M = 0.95 $\\alpha = 10^\\circ$ M = 1.00\n\nWing alone\nWing fuselage\n\n$\\frac{q_{\\text{wake}}}{q}$\n\n$\\alpha = 4^\\circ$ $\\alpha = 4^\\circ$\n\n$\\frac{q_{\\text{wake}}}{q}$\n\n$\\alpha = 0^\\circ$ $\\alpha = 0^\\circ$\n\n-80 -40 0 40 80 -80 -40 0 40 80 -80 -40 0 40 80\n\nTail height, $h_t$, percent semispan\n\nNACA\n\nFigure 12.- Continued.\n\n27", "timestamp": "2026-07-22T06:33:57.975391+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 55, "total_pages": 78, "image_filename": "19930082483_p55.jpg", "text": "NACA TN No. 1807\n53\n\n<!-- Image (154, 197, 875, 421) -->\n\n(a) Full admission (360°).\n\n<!-- Image (154, 522, 867, 742) -->\n\n(b) Partial admission.\n\nFigure 1. - Schematic relation of specific turbine-power concepts\nshowing specific power-loss dissipations.", "timestamp": "2026-07-22T06:33:58.635483+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 34, "total_pages": 46, "image_filename": "19930085548_p34.jpg", "text": "NACA RM No. EBL30\n33\n\n1077\n\nInlet-manifold\npressure\n(lb/sq in. abs.)\n○ 80\n□ 100\n◇ 120\n△ 135\n\nimep, lb/sq in.\nFuel-air ratio\n\n[Figure: Graph showing imep vs. fuel-air ratio for different inlet-manifold pressures]\n\nNACA\n\nFigure 11. - Effect of inlet-manifold pressure and\nfuel-air ratio on power output of experimental\ncylinder. Compression ratio, 5.25; inlet-manifold\ntemperature, 400° F.", "timestamp": "2026-07-22T06:34:00.671805+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 16, "total_pages": 29, "image_filename": "19930085899_p16.jpg", "text": "```markdown\nM = .78\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| 6 | | | |\n| | $M_1$ | | $M_2$ |\n| | .72 | | .74 |\n| | .73 | | .75 |\n| 4 | | | .76 |\n| | | | .77 |\n| | | | .78 |\n| 2 | | | .79 |\n| | | | |\n| 0 | | | .80 |\n| | 8 | 10 | 12 | 14 | 16 | 18 |\n\nM = .91\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| 6 | | | |\n| | $M_1$ | | $M_2$ |\n| | .84 | | .86 |\n| | .85 | | .87 |\n| 4 | | | .88 |\n| | | | .89 |\n| | | | .90 |\n| 2 | | | .91 |\n| | | | .92 |\n| | | | .93 |\n| 0 | | | .92 |\n| | 8 | 10 | 12 | 14 | 16 | 18 |\n\nVertical distance above bump, in.\n\nNominal boundary-layer thickness\n\nNACA\n\nM = 1.00\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| 6 | | | |\n| | $M_1$ | | $M_2$ |\n| | .92 | | .94 |\n| | .93 | | .95 |\n| 4 | | | .96 |\n| | | | .97 |\n| | | | .98 |\n| 2 | | | .99 |\n| | | | 1.00 |\n| | | | 1.01 |\n| | | | 1.02 |\n| 0 | | | 1.03 |\n| | | | 1.04 |\n| | 8 | 10 | 12 | 14 | 16 | 18 |\n\nM = 1.17\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| 6 | | | |\n| | $M_1$ | | $M_2$ |\n| | 1.07 | | 1.13 |\n| | 1.08 | | 1.14 |\n| | 1.09 | | 1.15 |\n| 4 | 1.10 | | 1.16 |\n| | 1.11 | | 1.17 |\n| | 1.12 | | 1.18 |\n| 2 | | | 1.19 |\n| | | | 1.20 |\n| | | | |\n| 0 | | | 1.21 |\n| | 8 | 10 | 12 | 14 | 16 | 18 |\n\nVertical distance above bump, in.\n\nStation on bump, in.\n\nStation on bump, in.\n\nFigure 5.— Typical Mach number contours over transonic bump in region of model location.\n\nNACA RM No. L9A21\n\n15\n```", "timestamp": "2026-07-22T06:34:03.982162+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 19, "total_pages": 92, "image_filename": "19930085870_p19.jpg", "text": "18\nCONFIDENTIAL\nNACA RM No. L9D07\n\nAPPENDIX A\n\nCALCULATION OF PRESSURE DRAG\n\nThe equations for computation of the pressure drag of triangular wings are as follows:\n\n(1) Mach line behind both the leading edge and the ridge line\n\n$$C_D = \\frac{2r^2}{\\beta\\pi(1-r^2)} \\left\\{ \\frac{1}{\\sqrt{1-n^2}} \\cos^{-1}n + \\frac{1}{r\\sqrt{1-r^2n^2}} \\left[ \\frac{\\pi}{2} + \\sin^{-1}(rn) \\right] \\right\\}$$\n(A1)\n\n(2) Mach line ahead of the leading edge but behind the ridge line\n\n$$C_D = \\frac{2r^2}{\\beta\\pi} \\left[ \\frac{G_2(n,r)}{r(1-r)^2} + \\frac{1}{r(1-r)} \\left( \\frac{\\pi}{2} - \\frac{\\log n}{\\sqrt{n^2-1}} - \\sin^{-1}\\frac{1}{n} \\right) \\right]$$\n(A2)\n\nwhere\n\n$$G_2(n,r) = \\frac{1-r}{1+r} \\left[ \\frac{\\log n}{\\sqrt{n^2-1}} + \\frac{r \\cosh^{-1}n}{\\sqrt{n^2-1}} + \\frac{2}{\\sqrt{1-r^2n^2}} \\tan^{-1} \\left( \\frac{\\sqrt{1-r^2n^2}}{\\sqrt{n^2-1}(1-rn)} \\right) \\right]$$\n(A3)\n\n(3) Mach line ahead of both leading edge and ridge line\n\n$$C_D = \\frac{2r^2}{\\beta\\pi} \\left[ \\frac{G_2'}{r(1-r)^2} - \\frac{F'}{(1-r)^2} + \\frac{1}{r(1-r)} \\left( \\frac{\\log nr}{\\sqrt{r^2n^2-1}} - \\frac{\\log n}{\\sqrt{n^2-1}} + \\sin^{-1}\\frac{1}{rn} - \\sin^{-1}\\frac{1}{n} \\right) \\right]$$\n(A4)\n\n[Handwritten annotation pointing to equation (A3): see ref. 7]\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:34:04.714373+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 33, "total_pages": 149, "image_filename": "19930083192_p33.jpg", "text": "NACA TN 1976\n29\n\ncalculations should yield essentially the same values for the larger distances since the conditions tend to approach the steady state. The discrepancy between experiment and strip theory appears to be due in part to positive pitch of the model during the traverse of the gust. Therefore, the effect of sweep on the unsteady-lift functions cannot be neglected if the shape of the curve is considered important.\n\nThe unsteady-lift theory as developed by Wagner, Jones, and others has been evolved for monoplanes only but the unsteady-lift functions for biplanes may be needed in connection with gust-load calculations. No theory is available on this problem, but consideration of the vortex sheets associated with the biplane indicates that, for the first few chords after a sudden change in angle of attack, the tip vortices of the wings are short and have a negligible effect on the mutually induced angle of attack of one wing upon the other. Since the equivalent monoplane theory assumes completely developed vortex systems with the shed vortex at infinity, it might be expected that for a biplane in a sharp gust the two wings would act independently, and, therefore, the unsteady-lift functions and other associated aerodynamic parameters for the biplane should be based on the characteristics for each individual wing. Figure 30, which has been reproduced from reference 8, shows the results of tests of a biplane. The results in the figure are the accelerations obtained as a result of experiment in the Langley gust tunnel, the curve predicted by assuming that the two wings act independently, and the curve predicted by assuming that the equivalent-monoplane theory holds. Infinite-aspect-ratio unsteady-lift functions were used. For the shorter gust-gradient distances where the tip vortices might be expected to have little effect, theory and experiment are in excellent agreement if the two wings are assumed to act independently. The computed curve for acceleration increment, based on reference 4, is considerably below the experimental data; thus the tip vortices have little effect, at least up to 8 or 10 wing chords. As the gradient distance is increased to 20 chords, the experimental data fall below both calculated curves because of the pitching action of the airplane.\n\nThe effect of compressibility on the unsteady-lift functions is of interest because it may change the shape of the curves, but little or no information on this problem is available. Jones has indicated that for subsonic speeds the effect of compressibility would be to reduce the aspect ratio of the wing according to the relation $\\sqrt{1 - M^2}$. The E correction of reference 27 is changed at the same time to correspond. For the transonic range where mixed flows occur the variation in the unsteady-lift functions is unknown. At supersonic speeds the effect of unsteady lift would be expected to disappear since the effect requires the transmission of pressures forward. It would be expected that the unsteady-lift functions would change only slightly up to high subsonic speeds and that at supersonic speeds, the functions would disappear.", "timestamp": "2026-07-22T06:34:04.887365+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 2, "total_pages": 47, "image_filename": "19930085919_p2.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION\n\nEMPLOYING A WING SWEPT BACK $63^\\circ$. - EFFECTS\n\nOF SPLIT FLAPS, ELEVONS, AND LEADING-\n\nEDGE DEVICES AT LOW SPEED\n\nBy Edward J. Hopkins\n\nSUMMARY\n\nAn investigation was conducted to evaluate the effects of split flaps, elevons, sharp leading edges, drooped-nose flaps, and extended-nose flaps on the lift, drag, and pitching-moment characteristics at low speed of a wing-fuselage combination having a wing with the leading edge swept back $63^\\circ$ and having an aspect ratio of 3.5. Measurements were also made of the rolling moments produced by the elevons. In addition, a study was made to evaluate the effects of the fuselage and possible Reynolds number effects on the characteristics of the wing.\n\nThe optimum chordwise position of the split flap for increasing the lift coefficient attained before the occurrence of longitudinal instability and for reducing the drag at high lift coefficients was the position with the split flap hinge line coincident with the trailing edge of the wing. The effectiveness of the elevons for producing rolling moments was nearly constant up to an angle of attack of $9^\\circ$, but decreased at greater angles of attack. The full-span leading-edge flaps increased the lift coefficient attained before the occurrence of longitudinal instability considerably more than did the 50-percent span leading-edge flaps. The extended-nose flap was about twice as effective as the drooped-nose flap in reducing the drag of the model at the higher lift coefficients.\n\nINTRODUCTION\n\nA coordinated program is being conducted at Ames Aeronautical Laboratory to provide information throughout an extensive range of Mach and Reynolds numbers on a wing-fuselage combination employing\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:34:09.146591+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 52, "total_pages": 65, "image_filename": "19930082546_p52.jpg", "text": "NACA TN No. 1870\n51\n\nPressure, P, dynes/cm²\n1800\n1600\n1400\n1200\n1000\n800\n600\n400\n200\n0\n\nx/D\n+1/4\n+1/8\n0\n-1/8\n-1/4\n\nmB\n○ 2\n□ 4\n◇ 6\n△ 8\n\n(c) M_t = 1.00.\nFigure 12.— Concluded.", "timestamp": "2026-07-22T06:34:14.092985+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 66, "total_pages": 114, "image_filename": "19930086061_p66.jpg", "text": "62\nNACA RM L9J07\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n-3\n-2\n-1 P 1\n0\n1\n\n-3\n-2\n-1\n0\n1\n\n(a) $\\psi = 0^\\circ$\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n-4\n-3\n-2\n-1 P 1\n0\n1\n\n-4\n-3\n-2\n-1\n0\n1\n\n$10^\\circ$\n\n(b) $\\psi = 10^\\circ$\n\nNACA\n\nFigure 20.- Pressure distribution about wing 2 at various angles of yaw; $\\alpha = 34.1^\\circ$.", "timestamp": "2026-07-22T06:34:14.410722+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 30, "total_pages": 32, "image_filename": "19930085847_p30.jpg", "text": "28\nCONFIDENTIAL\nNACA RM A9D04\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (266, 152, 732, 499) -->\n\n<!-- Image (281, 508, 626, 839) -->\n\n(c) M = 0.81 ($C_L = 0.21$).\n\nFigure 8.- Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:34:16.814523+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 69, "total_pages": 98, "image_filename": "19930086073_p69.jpg", "text": "NACA RM A9H04\n\nLift coefficient, $C_L$\n\n$\\delta_{aL}, deg$\n□ +10.8\n○ 0\n◇ -10.8\n\nRolling-moment coefficient, $C_l$\n\nYawing-moment coefficient, $C_n$\n\nSide-force coefficient, $C_Y$\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 14.— Concluded.\n\n67", "timestamp": "2026-07-22T06:34:16.925087+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 31, "total_pages": 60, "image_filename": "19930085862_p31.jpg", "text": "NACA RM No. L9A07\n29\n\n$C_L$\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\n$\\delta_e$\n(deg)\n$\\diamond$ -25\n$\\circ$ 0\n$\\square$ 25\n\n$C_m$\n.04\n0\n-.04\n-.08\n-.12\n-.16\n\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\n(c) $C_L$ and $C_m$ against $\\alpha$.\nFigure 7.- Concluded.", "timestamp": "2026-07-22T06:34:17.586112+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 4, "total_pages": 34, "image_filename": "19930085913_p4.jpg", "text": "NACA RM L9F24\n3\n\n$\\Lambda$\nsweep angle, positive for sweepback, degrees\n\n$x_\\alpha$\ndistance between elastic axis and center of gravity of wing section, referred to half-chord\n\n$e_w$\ndistance between elastic axis of wing section and center of gravity of weight, referred to half-chord, negative for forward weight location\n\n$I_{CG}$\nmass moment of inertia of wing section about its center of gravity, inch-pound-second$^2$ per inch\n\n$I_{EA}$\nmass moment of inertia of wing section about its elastic axis, inch-pound-second$^2$ per inch\n\n$I_w$\nmass moment of inertia of weight about an axis parallel to leading edge through its center of gravity, inch-pound-second$^2$\n\n$EI$\nbending rigidity of wing section, pound-inch$^2$\n\n$GJ$\ntorsional rigidity of wing section, pound-inch$^2$\n\n$m$\nmass of wing per unit length, slugs per foot\n\n$r_\\alpha$\nnondimensional radius of gyration of wing section about its elastic axis $\\left(\\sqrt{\\frac{I_{EA}}{mb^2}}\\right)$\n\n$q_F$\ndynamic pressure at flutter, pounds per square foot\n\n$\\rho$\nair density, slugs per cubic foot\n\n$v_F$\ntrue-stream velocity at flutter, feet per second\n\n$\\kappa$\nmass ratio $\\left(\\frac{\\pi\\rho b^2}{m}\\right)$\n\n$g_h, g_\\alpha$\nstructural damping coefficient in degree of freedom indicated by subscript\n\n$\\alpha$\nangle of attack of wing section, positive leading edge up\n\n$h$\nbending deflection of wing section at elastic axis, positive downward", "timestamp": "2026-07-22T06:34:18.515841+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 16, "total_pages": 26, "image_filename": "19930085890_p16.jpg", "text": "NACA RM No. E9C11\n\nCounterbalanced\nweighing beam\n\nStrain gage\n\nHelium\n\nLiquid-oxygen\ntank\n\nDiborane\ntank\n\nLiquid nitrogen\n\nEngine\n\nPivoted\nthrust\nstand\n\nNACA\nC-21248\n4-21-48\n\nFigure 2. - Rocket engine and auxiliary equipment.\n\n15", "timestamp": "2026-07-22T06:34:19.075820+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 16, "total_pages": 33, "image_filename": "19930085900_p16.jpg", "text": "NACA RM L9D20\n15\n\nCONFIDENTIAL\n\nMaximum trim permitted by apparatus\n\nTrim, deg\nResistance, effective hydrodynamic lift, and load on water, lb\n\nTrim\nResistance\nLoad on water\nEffective hydrodynamic lift\n\nSpeed, fps\n\nFigure 5.- Hydrodynamic characteristics of basic or unmodified model.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:34:20.044933+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 37, "total_pages": 62, "image_filename": "19930082918_p37.jpg", "text": "**Page intentionally left blank**\n\nFrom pge 37 — 59 (every other page is blank)\n\n**Page intentionally left blank**", "timestamp": "2026-07-22T06:34:20.811731+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 37, "total_pages": 47, "image_filename": "19930083221_p37.jpg", "text": "NACA TN No. 1824\n35\n\n2. Boundary-value problem of the second kind, surface specified.-\n\nOver the xy plane, the imposed conditions are $\\Delta w_o = 0$ everywhere and, except for the region occupied by the wing, $\\Delta u_o = 0$. Over the region occupied by the wing $w_o = w_u = w_l = f(x,y)$ where $f(x,y)$ is determined by known camber, twist, and angle of incidence. (The delta notation again indicates the jump in the value of the variable at the $z = 0$ plane. Subscripts u and l indicate conditions on the upper and lower surface, respectively, of this plane.)\n\nThe nature of the differential equation shows that the value of $\\varphi$ is a consequence of boundary conditions along lateral strips. If, as in figure 14, the two leading edges are given by the expressions $y = b_1(x)$ and $y = b_2(x)$, the velocity potential is expressible in the form\n\n[Figure: Diagram of a swept-back plan form with curved trailing edge. The y-axis is horizontal, the x-axis is vertical pointing downwards. The leading edges are labeled $y=b_1(x)$ and $y=b_2(x)$.]\n\nFigure 14.- Swept-back plan form with curved trailing edge.\n\n$$ \\varphi(x,y,z) = \\frac{z}{2\\pi} \\int_{b_1}^{b_2} \\frac{\\Delta \\varphi_o(x,y_1) dy_1}{(y-y_1)^2 + z^2} \\quad (59) $$\n\nIf the boundary-value problem is one of the first kind, the general expression for $\\varphi$ follows from a direct integration after noting that\n\n$$ \\Delta \\varphi_o(x,y) = \\int_{b_1}^{y} \\Delta u_o(x,y_1) dy_1 \\quad (60) $$\n\nSince, moreover, load coefficient $\\Delta p/q$ is related to $\\Delta u_o$ by means of the equation\n\n$$ \\frac{\\Delta p}{q} = \\frac{2 \\Delta u_o}{V_o} $$\n\nit follows that the velocity potential $\\varphi$ can be found for any prescribed load distribution of a given plan form. The value of vertical induced velocity, evaluated at $z = 0$, then suffices to calculate the twist and angle of attack of the wing.", "timestamp": "2026-07-22T06:34:21.633474+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 30, "total_pages": 31, "image_filename": "19930085859_p30.jpg", "text": "28\n\nM = 1.05\n$\\alpha = 10^\\circ$\n\nM = 1.10\n$\\alpha = 10^\\circ$\n\nM = 1.15\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\nWing alone\nWing-fuselage\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\n$\\alpha = 4^\\circ$\n$\\alpha = 4^\\circ$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\n$\\alpha = 0^\\circ$\n$\\alpha = 0^\\circ$\n\n- 80 - 40 0 40 80 - 80 - 40 0 40 80 - 80 - 40 0 40 80\n\nTail height, $h_t$, percent semispan\n\nNACA\n\nFigure 12.— Concluded.\n\nNACA RM No. 19D25", "timestamp": "2026-07-22T06:34:23.321661+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 35, "total_pages": 46, "image_filename": "19930085548_p35.jpg", "text": "```markdown\n34\nNACA RM No. E8L30\n\n1077\n\n<!-- Image (170, 268, 789, 714) -->\n\nFigure 12. - Effect of inlet-manifold temperature on power output of the experimental cylinder. Compression ratio, 4.5; inlet-manifold pressure, 100 pounds per square inch; fuel-air ratio, 0.025.\n```", "timestamp": "2026-07-22T06:34:23.599540+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 17, "total_pages": 29, "image_filename": "19930085899_p17.jpg", "text": "```markdown\n16\n\nReynolds number, R x 10⁻⁶\n\nMean\n\nMach number, M\n\nNACA\n\nFigure 6.— Variation of test Reynolds number with Mach number for model with 45° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\nNACA RM No. 19421\n```", "timestamp": "2026-07-22T06:34:24.642578+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 53, "total_pages": 65, "image_filename": "19930082546_p53.jpg", "text": "52\nNACA TN No. 1870\n\n<!-- Image (114, 166, 824, 833) -->\n\nFigure 13.- Free-space pressure distribution of the first harmonic of the NACA 4-(5)(08)-03 propeller.\nB = 2; $\\frac{d}{D}$ = 0.083; $F_{\\theta}$ = 0.6R.", "timestamp": "2026-07-22T06:34:34.326773+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 20, "total_pages": 92, "image_filename": "19930085870_p20.jpg", "text": "NACA RM No. L9D07 CONFIDENTIAL 19\n\nwhere\n\n$$\nG_2' = \\frac{1 - r}{1 + r} \\left\\{ \\frac{\\log n}{\\sqrt{n^2 - 1}} + \\frac{r \\cosh^{-1} n}{\\sqrt{n^2 - 1}} + \\frac{1}{\\sqrt{r^2 n^2 - 1}} \\log \\left[ 1 + \\frac{2 \\sqrt{r^2 n^2 - 1}}{n(1 - r) + \\sqrt{n^2 - 1} - \\sqrt{r^2 n^2 - 1}} \\right] \\right\\}\n\\tag{A5}\n$$\n\nand\n\n$$\nF' = \\frac{1 - r}{1 + r} \\left\\{ \\frac{\\log rn}{\\sqrt{r^2 n^2 - 1}} + \\frac{1}{\\sqrt{n^2 - 1}} \\log \\left[ \\frac{rn^2 - 1 + \\sqrt{(r^2 n^2 - 1)(n^2 - 1)}}{n(1 - r)} \\right] \\right\\}\n\\tag{A6}\n$$\n\n$$\n\\beta = \\sqrt{M_0^2 - 1}\n$$\n\nτ thickness ratio at root chord  \nr location of ridge-line apex in percent root chord from trailing edge  \n\n$$\nn = \\frac{\\tan m}{\\tan \\epsilon}\n$$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:34:35.129748+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 67, "total_pages": 114, "image_filename": "19930086061_p67.jpg", "text": "NACA RM L9J07\n63\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n(c) $\\psi = 20^\\circ$\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n(d) $\\psi = 35^\\circ$\n\nFigure 20.- Concluded.", "timestamp": "2026-07-22T06:34:37.061862+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 31, "total_pages": 32, "image_filename": "19930085847_p31.jpg", "text": "NACA RM A9D04 CONFIDENTIAL 29\n\nDrag-producing area\nThrust-producing area\n\n[Figure: Graph with y/c (upper and lower) vs. Δp/q, showing shaded regions labeled “Drag-producing area” and “Thrust-producing area”, and a curve marked P_crit]\n\n[Figure: Graph with Δp/q vs. x/c, showing data points for “upper” and “lower” surfaces, a horizontal line labeled P_crit, and a region labeled t_max; includes NACA logo]\n\n(d) M = 0.82 (C_L = 0.12).\n\nFigure 8.- Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:34:39.805569+00:00"}