PDF/ocr_dataset_1033.jsonl

{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 13, "total_pages": 29, "image_filename": "19930085879_p13.jpg", "text": "NACA RM L9D11\n\n11\n\n[Figure: A black-and-white photograph of a model rocket or test vehicle mounted on a wooden support frame. The vehicle has a pointed nose cone, a cylindrical body, and fins at the base. A label in the upper right corner of the photo reads “NACA I-57090”.]\n\nFigure 2.—PM-11 pilot-escape test vehicle.", "timestamp": "2026-07-22T05:17:24.356288+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 10, "total_pages": 96, "image_filename": "19930085880_p10.jpg", "text": "8\n\n1.667\n\n.208\n\n.083\n\nB\n\nA\n\nB\n\nA\n\nSection B-B\n\nView A-A\n\nAll longitudinal sections are circular arcs\nwith a height at the center of 5 percent of\nthe chord which forms the bottom of the\nsection.\n\nNACA\n\nFigure 1.- Details of rectangular planing surface with flat bottom (model 250A).\n(All dimensions are in feet.)\n\nNACA RM No. L9C03", "timestamp": "2026-07-22T05:17:25.006835+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 8, "total_pages": 37, "image_filename": "19930085889_p8.jpg", "text": "NACA RM L9F14 CONFIDENTIAL 7\n\n| q | M | R |\n|---|---|---|\n| 4 | 0.051 | 280,000 |\n| 8 | .073 | 395,000 |\n| 16 | .104 | 558,000 |\n| 25 | .131 | 718,000 |\n| 40 | .166 | 880,000 |\n| 65 | .211 | 1,116,000 |\n\nCORRECTIONS\n\nApproximate jet-boundary corrections (similar to those of reference 7) based on unswept-wing theory have been applied to the angle of attack, the drag coefficient, and the rolling-moment coefficient. Corrections for blocking or support-strut tares have not been applied to the results.\n\nRESULTS AND DISCUSSION\n\nStraight-Flow Characteristics\n\nThe lift, drag, and pitching-moment characteristics of the three wings, each tested in combination with the fuselage, are presented in figure 6. The pitching-moment results at low lift coefficients indicate that the aerodynamic center moved rearward, from 17.6 percent to 27.0 percent of the mean aerodynamic chord, as the angle of sweepback was increased from $3.6^\\circ$ to $46.7^\\circ$. The theoretical results given in reference 8 predict almost no change in the aerodynamic-center location of plain wings over this range of sweep angles for the particular aspect ratio and taper ratio of the wings investigated. The differences between theory and experiment probably resulted from the fact that a fuselage was used in the tests.\n\nBecause each of the wings was constructed in two semispan segments with mounting blocks at the inboard ends for attachment to a fuselage, true wing-alone characteristics could not be obtained. An attempt to simulate, as nearly as possible, the wing-alone condition was made, however, for the $46.7^\\circ$ sweptback wing. The wing segments were supported by cover plates and the entire root region was faired with balsa wood and clay. (See fig. 5.) Lift and pitching-moment results obtained with this model (wing alone) and with the same wing in\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:17:26.372189+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 8, "total_pages": 31, "image_filename": "19930085881_p8.jpg", "text": "6 CONFIDENTIAL NACA RM L9D12\n\ndrag rise of the double-wedge sections on the rectangular wings occurs at lower Mach numbers than for the other sections tested. The higher drag of this section at high subsonic Mach numbers is probably due to boundary-layer separation induced by the abrupt change of direction of the upper and lower surfaces at the midchord point.\n\nThe results for the sweptback configuration given in figure 8(b) are coincident within the experimental accuracy indicating that, for the Mach number range investigated, the effects of airfoil section on the drag of wings having $45^\\circ$ sweep are small.\n\nCONCLUSIONS\n\nThe main conclusions based on the results of the present investigation may be stated as follows:\n\n1. Reversal of rolling effectiveness of plain ailerons on the rectangular wings was obtained for trailing-edge angles of the order of $20^\\circ$ in the Mach number range from about 0.88 to 0.93. Reversal was not obtained for trailing-edge angles of the order of $10^\\circ$.\n\n2. Reversal of effectiveness of plain ailerons on the swept wings was obtained for trailing-edge angles of the order of $20^\\circ$ over a relatively large Mach number range. Reversal was not obtained for trailing-edge angles of the order of $10^\\circ$.\n\n3. The aileron effectiveness was not appreciably affected by changes in the shape of the forward part of the airfoil section.\n\n4. For the rectangular wings near Mach numbers of unity, the blunt-nose airfoil sections had slightly lower drag than did the sharp-nose sections. At the higher supersonic Mach numbers investigated, the sharp-nose airfoil sections had considerably less drag than did the blunt-nose sections.\n\n5. At Mach number 1.7, the drag of the unswept wing having the symmetrical double-wedge section was equal to that of the wings having $45^\\circ$ sweepback.\n\n6. For the sweptback wings, the various airfoil sections investigated produced no measurable differences in drag in the Mach number range investigated.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:17:29.517355+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 47, "total_pages": 49, "image_filename": "19930082498_p47.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:17:29.700372+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 57, "total_pages": 62, "image_filename": "19930082485_p57.jpg", "text": "NACA TN No. 1810\n\n.50\nRatio of\nspecific\nheats, $\\gamma$\n1.40\n1.36\n1.32\n.45\n.40\n.35\n.30\nf\n.25\n.20\n.15\n.10\n.05\n0\n.1\n.2\n.3\n.4\n.5\n.6\n$\\sqrt{z_m}$\n\n(a) Chart for finding $\\sqrt{z_m}$.\nFigure 14. - Chart for finding average velocity through blade channel.", "timestamp": "2026-07-22T05:17:30.162959+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 22, "total_pages": 46, "image_filename": "19930085542_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:17:31.897084+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 15, "total_pages": 36, "image_filename": "19930085869_p15.jpg", "text": "NACA RM L9D15\n13\n\nCONFIDENTIAL\n\n[Figure: Bottom view of an aircraft]\n\n(b) Bottom view.\nFigure 1.- Concluded.\nCONFIDENTIAL\nNACA\nL-56645.1", "timestamp": "2026-07-22T05:17:33.504780+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 39, "total_pages": 66, "image_filename": "19930082914_p39.jpg", "text": "38\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:17:44.482718+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 1, "total_pages": 23, "image_filename": "19930085906_p1.jpg", "text": "NACA RM E9F20\n\nCONFIDENTIAL\n\nCOPY\nRM E9F20\n\nNACA\n\nRESEARCH MEMORANDUM\n\nINVESTIGATION OF INTERNAL REGENERATIVE FUEL-HEATING SYSTEM\n\nFOR 20-INCH RAM JET\n\nBy Sol Baker and Eugene Perchonok\n\nLewis Flight Propulsion Laboratory\n\nCleveland, Ohio\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: RESEARCH ABSTRACT NO. 101\nDATE: MAY 25, 1956\nWHL\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nSeptember 1, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:17:44.806395+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 61, "total_pages": 66, "image_filename": "19930082245_p61.jpg", "text": ".8\n.7\n.6\n$M_{cr}$ .5\n.4\n.3\n.2\n(a) $c_n = 0$.\n\n.8\n.7\n.6\n.5\n.4\n.3\n.2\n(b) $c_n = 0.2$.\n\nTrue-contour aileron\nBeveled-trailing-edge aileron\n\n.8\n.7\n.6\n$M_{cr}$ .5\n.4\n.3\n.2\n(c) $c_n = 0.4$.\n\n.8\n.7\n.6\n.5\n.4\n.3\n.2\n(d) $c_n = 0.6$.\n\n-16 -12 -8 -4 0 4 8 12 16\nAileron deflection, $\\delta_a$, deg\n\n-16 -12 -8 -4 0 4 8 12 16\nAileron deflection, $\\delta_a$, deg\n\nFigure 13.-Variation of critical Mach number with aileron deflection for an NACA 66,1-115 airfoil section equipped with unsealed 0.20c plain ailerons.\n\n60\nNACA TN NO. 1596", "timestamp": "2026-07-22T05:17:50.833418+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 18, "total_pages": 24, "image_filename": "19930085626_p18.jpg", "text": "NACA RM No. L8K23\n17\n\nCONFIDENTIAL\nTip obtained by revolving\nairfoil section around chord\nline\n\nc/4 line\n40°\nA\nA\n4.05\n.93\n11.50\n4.50\n1.69\n8.48\n\nSection A-A\nCircular-arc section\nnormal to c/4 line.\nThickness ratio, 0.10\n\n$C_D$\n.08\n.04\n0\n\n$pb/2V$\n.08\n.04\n0\n-.04\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\n$M$\n\n(b) Inboard aileron configuration. $\\delta_a = 5^\\circ$.\nFigure 4.— Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:17:53.083163+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 20, "total_pages": 33, "image_filename": "19930085544_p20.jpg", "text": "NACA RM No. L8K26\n19\n\n[Figure: Diagram showing notation used for oscillating airfoil in unsteady flow (from reference 2).]\n\nFigure 2.-- Diagram showing notation used for oscillating airfoil in unsteady flow (from reference 2).", "timestamp": "2026-07-22T05:17:56.250181+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 51, "total_pages": 99, "image_filename": "19930082511_p51.jpg", "text": "NACA TN No. 1826\n49\n\n(4) $$ \\lim_{\\xi \\to +\\infty} \\frac{\\partial \\phi_A}{\\partial \\xi} = \\lim_{\\xi \\to -\\infty} \\frac{\\partial \\phi_A}{\\partial \\xi} $$\n\n(5) $$ \\left. \\frac{\\partial \\phi_A}{\\partial \\rho} \\right|_{\\rho=1} = 0 \\quad (\\xi = a) $$\n\nConditions (1) and (2) are satisfied by assuming $g_0(\\xi) = 0$ for $\\xi < a$ and $\\xi > b$. Thus\n\n$$ \\phi_A = \\frac{1}{\\pi} \\int_0^\\infty \\left[ \\frac{J_0(i\\rho q)}{iq J_0'(iq)} \\int_a^b g_0(\\beta) \\cos q(\\beta - \\xi) \\, d\\beta - \\frac{2}{q^2} \\int_a^b g_0(\\beta) \\, d\\beta \\right] dq \\quad (32) $$\n\nAgain it is desired to find $g_0(\\xi)$ for $a < \\xi < b$ so that $\\phi_A$ will satisfy conditions (3), (4), and (5). The representation of $g_0(\\xi)$ in the same form as before (equation 27)\n\n$$ g_0(\\xi) = h_{00} \\sin \\frac{\\pi}{2} \\frac{\\xi - a}{b - a} + \\sum_{n=1}^\\infty h_{0n} \\sin n\\pi \\frac{\\xi - a}{b - a} $$\n\nautomatically satisfies condition (5). Substituting this series in equation (32) gives\n\n$$ \\phi_A = \\sum_{n=0}^\\infty h_{0n} P_{0n}(\\xi, \\rho) $$\n\nwhere\n\n$$ P_{00}(\\xi, \\rho) = \\frac{1}{\\pi} \\int_0^\\infty \\left[ \\frac{J_0(i\\rho q)}{iq J_0'(iq)} \\int_a^b \\sin \\frac{\\pi}{2} \\frac{\\beta - a}{b - a} \\cos q(\\beta - \\xi) \\, d\\beta \\right. $$\n\n$$ \\left. - \\frac{2}{q^2} \\int_a^b \\sin \\frac{\\pi}{2} \\frac{\\beta - a}{b - a} \\, d\\beta \\right] dq $$", "timestamp": "2026-07-22T05:18:00.110422+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 14, "total_pages": 29, "image_filename": "19930085879_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:18:02.165376+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 31, "total_pages": 58, "image_filename": "19930082617_p31.jpg", "text": "30\nNACA TN 1962\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in.-lb)\n1 36.0 X $10^3$\n2 72.0 X $10^3$\n3 108.0 X $10^3$\n4 144.0 X $10^3$\n5 180.0 X $10^3$\n6 216.0 X $10^3$\n7 288.0 X $10^3$\n\n2.57\"\nA\nBand O\nA-A\n45°\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n20 16 12 8 4 0 -4 -8 -12 -16 -20 X $10^{-4}$\nStrain\n\n1 2 3 4 5 6 7\n\nNACA\n\nFigure 19.- Strain diagram of cylinder 77. Band O.", "timestamp": "2026-07-22T05:18:03.416148+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 29, "total_pages": 53, "image_filename": "19930082542_p29.jpg", "text": "28\nNACA TN No. 1867\n\nTABLE III.- STUDY OF EFFECT OF HOLDING TIME AT 1200° F IN TEST UNIT\nBEFORE LOADING ON RUPTURE LIFE OF LOW-CARBON N-155\nBAR STOCK AT 1200° F\n\n| Treatment (1) | Holding time (hr) | Stress (psi) | Rupture time (hr) | Elongation in 1 in. (percent) | Reduction of area (percent) |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| 2200° F 1 hr W.Q. | 0.8 | 45,000 | 60 | 4 | 17.8 |\n| | 24.0 | 45,000 | 108 | 4 | 13.3 |\n| 2200° F 1 hr W.Q.; 1600° F 24 hr | .9 | 50,000 | 15 | 28 | 43.7 |\n| | 24.0 | 50,000 | 48 | 26 | 36.2 |\n| 2050° F 2 hr W.Q. | .5 | 45,000 | 180 | 8 | 14.4 |\n| | 24.0 | 45,000 | 130 | 12.5 | 14.4 |\n| 2050° F 2 hr W.Q.; 1400° F 24 hr | .7 | 50,000 | 98 | 31 | 34 |\n| | 24.0 | 50,000 | 79 | 27.5 | 30.8 |\n| 2050° F 2 hr W.Q.; 1400° F 24 hr | .5 | 50,000 | 41 | 36 | 37.9 |\n| | 24.0 | 50,000 | 44 | 45 | 40.8 |\n\n1W.Q., water quenched.\nNACA", "timestamp": "2026-07-22T05:18:09.325417+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 11, "total_pages": 96, "image_filename": "19930085880_p11.jpg", "text": "NACA RM No. L9C03\n\n1.667\n\n.208\n\nB\n.083\nA\nB\nA\n\nAll longitudinal sections are circular arcs\nwith a height at the center of 5 percent of\nthe chord which forms the bottom of the\nsection.\n\n.208R\n.028\nSection B-B\n\nView A-A\n\nNACA\n\nFigure 2. - Details of rectangular planing surface with transverse bottom curvature (model 250B).\n(All dimensions are in feet.)\n\n9", "timestamp": "2026-07-22T05:18:10.110054+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 48, "total_pages": 49, "image_filename": "19930082498_p48.jpg", "text": "NACA TN No. 1838\n47\n\n<!-- Image (45, 162, 888, 726) -->\n\nFigure 7.- Radial survey of sound field 2 feet from the end of the tail pipe. Engine speed, 1650 rpm.", "timestamp": "2026-07-22T05:18:12.276809+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 16, "total_pages": 36, "image_filename": "19930085869_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:18:14.722036+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 23, "total_pages": 46, "image_filename": "19930085542_p23.jpg", "text": "NACA RM No. L5L29\n\n[Figure: Model 5 mounted in tunnel. A wing = 2.31; A fin = 0.77. Profile of wing, NACA 0012; profile of fin, flat plate.]\n\nFigure 5.- Model 5 mounted in tunnel. $A_{\\text{wing}} = 2.31$; $A_{\\text{fin}} = 0.77$. Profile of wing, NACA 0012; profile of fin, flat plate.\n\nNACA\nL-55510\n\n21", "timestamp": "2026-07-22T05:18:15.418880+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 58, "total_pages": 62, "image_filename": "19930082485_p58.jpg", "text": "NACA TN No. 1810\n\n.50\n.45\n.40\n.35\n.30\n.25\n.20\n.15\n.10\n.05\n0\n\nRatio of\nspecific\nheats, $\\gamma$\n1.40\n1.36\n1.32\n\nf\n\n0 .05 .10 .15 .20 .25 .30\ng\n\n(b) Chart for finding f.\nFigure 14. - Concluded. Chart for finding average velocity through blade channel.", "timestamp": "2026-07-22T05:18:15.984192+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 1, "total_pages": 36, "image_filename": "19930085912_p1.jpg", "text": "RESTRICTED\nCOPY NO.\nRM No. E9C16\n587\n\nNACA RM No. E9C16\n\nNACA\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION OF HOT-GAS BLEEDBACK\nFOR ICE PROTECTION OF TURBOJET ENGINES\nII - NACELLE WITH LONG STRAIGHT AIR INLET\nBy Edmund E. Callaghan and Robert S. Ruggeri\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\n\nCLASSIFIED DOCUMENT\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\nAUTHORITY J. W. CROWLEY\nDATE 12-14-53 CHANGE #1905 E. L. B.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nMay 26, 1949\n\nRESTRICTED", "timestamp": "2026-07-22T05:18:23.556035+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 40, "total_pages": 66, "image_filename": "19930082914_p40.jpg", "text": "NACA TN No. 1857\n39\n\n[Figure: Photograph of an interferometer apparatus. A ruler marked in inches is visible at the bottom. A label in the lower right corner reads \"NACA L-55231\".]\n\nFigure 6.- Interferometer. (Splitter plates, lower left and upper right; mirrors, upper left and lower right.)", "timestamp": "2026-07-22T05:18:23.835248+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 35, "total_pages": 78, "image_filename": "19930082618_p35.jpg", "text": "```markdown\nNACA TN 1945\n\nR\n0.7 x 10⁶\n1.0\n1.5\n2.0\n2.5\nFlagged symbols denote\nstandard roughness\n\nSection drag coefficient, c_d\n.028\n.024\n.020\n.016\n.012\n.008\n.004\n0\n-.8 -.4 0 .4 .8 1.2\nSection lift coefficient, c_l\n\nR\n3.0 x 10⁶\n6.0\n9.0\nFlagged symbols denote\nstandard roughness\n\nSection drag coefficient, c_d\n.020\n.016\n.012\n.008\n.004\n0\n-1.2 -.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, c_l\n\nMoment coefficient, c_m,c/4\n0\n-.1\n-.2\n-.3\n-.4\n-.5\n-.6\n-.7\n-.8\n-.9\n-1.0\n-1.1\n-1.2\n-1.3\n-1.4\n-1.5\n-1.6\n-1.7\n-1.8\n-1.9\n-2.0\n-2.1\n-2.2\n-2.3\n-2.4\n-2.5\n-2.6\n-2.7\n-2.8\n-2.9\n-3.0\n-3.1\n-3.2\n-3.3\n-3.4\n-3.5\n-3.6\n-3.7\n-3.8\n-3.9\n-4.0\n-4.1\n-4.2\n-4.3\n-4.4\n-4.5\n-4.6\n-4.7\n-4.8\n-4.9\n-5.0\n-5.1\n-5.2\n-5.3\n-5.4\n-5.5\n-5.6\n-5.7\n-5.8\n-5.9\n-6.0\n-6.1\n-6.2\n-6.3\n-6.4\n-6.5\n-6.6\n-6.7\n-6.8\n-6.9\n-7.0\n-7.1\n-7.2\n-7.3\n-7.4\n-7.5\n-7.6\n-7.7\n-7.8\n-7.9\n-8.0\n-8.1\n-8.2\n-8.3\n-8.4\n-8.5\n-8.6\n-8.7\n-8.8\n-8.9\n-9.0\n-9.1\n-9.2\n-9.3\n-9.4\n-9.5\n-9.6\n-9.7\n-9.8\n-9.9\n-10.0\n-10.1\n-10.2\n-10.3\n-10.4\n-10.5\n-10.6\n-10.7\n-10.8\n-10.9\n-11.0\n-11.1\n-11.2\n-11.3\n-11.4\n-11.5\n-11.6\n-11.7\n-11.8\n-11.9\n-12.0\n-12.1\n-12.2\n-12.3\n-12.4\n-12.5\n-12.6\n-12.7\n-12.8\n-12.9\n-13.0\n-13.1\n-13.2\n-13.3\n-13.4\n-13.5\n-13.6\n-13.7\n-13.8\n-13.9\n-14.0\n-14.1\n-14.2\n-14.3\n-14.4\n-14.5\n-14.6\n-14.7\n-14.8\n-14.9\n-15.0\n-15.1\n-15.2\n-15.3\n-15.4\n-15.5\n-15.6\n-15.7\n-15.8\n-15.9\n-16.0\n-16.1\n-16.2\n-16.3\n-16.4\n-16.5\n-16.6\n-16.7\n-16.8\n-16.9\n-17.0\n-17.1\n-17.2\n-17.3\n-17.4\n-17.5\n-17.6\n-17.7\n-17.8\n-17.9\n-18.0\n-18.1\n-18.2\n-18.3\n-18.4\n-18.5\n-18.6\n-18.7\n-18.8\n-18.9\n-19.0\n-19.1\n-19.2\n-19.3\n-19.4\n-19.5\n-19.6\n-19.7\n-19.8\n-19.9\n-20.0\n-20.1\n-20.2\n-20.3\n-20.4\n-20.5\n-20.6\n-20.7\n-20.8\n-20.9\n-21.0\n-21.1\n-21.2\n-21.3\n-21.4\n-21.5\n-21.6\n-21.7\n-21.8\n-21.9\n-22.0\n-22.1\n-22.2\n-22.3\n-22.4\n-22.5\n-22.6\n-22.7\n-22.8\n-22.9\n-23.0\n-23.1\n-23.2\n-23.3\n-23.4\n-23.5\n-23.6\n-23.7\n-23.8\n-23.9\n-24.0\n-24.1\n-24.2\n-24.3\n-24.4\n-24.5\n-24.6\n-24.7\n-24.8\n-24.9\n-25.0\n-25.1\n-25.2\n-25.3\n-25.4\n-25.5\n-25.6\n-25.7\n-25.8\n-25.9\n-26.0\n-26.1\n-26.2\n-26.3\n-26.4\n-26.5\n-26.6\n-26.7\n-26.8\n-26.9\n-27.0\n-27.1\n-27.2\n-27.3\n-27.4\n-27.5\n-27.6\n-27.7\n-27.8\n-27.9\n-28.0\n-28.1\n-28.2\n-28.3\n-28.4\n-28.5\n-28.6\n-28.7\n-28.8\n-28.9\n-29.0\n-29.1\n-29.2\n-29.3\n-29.4\n-29.5\n-29.6\n-29.7\n-29.8\n-29.9\n-30.0\n-30.1\n-30.2\n-30.3\n-30.4\n-30.5\n-30.6\n-30.7\n-30.8\n-30.9\n-31.0\n-31.1\n-31.2\n-31.3\n-31.4\n-31.5\n-31.6\n-31.7\n-31.8\n-31.9\n-32.0\n-32.1\n-32.2\n-32.3\n-32.4\n-32.5\n-32.6\n-32.7\n-32.8\n-32.9\n-33.0\n-33.1\n-33.2\n-33.3\n-33.4\n-33.5\n-33.6\n-33.7\n-33.8\n-33.9\n-34.0\n-34.1\n-34.2\n-34.3\n-34.4\n-34.5\n-34.6\n-34.7\n-34.8\n-34.9\n-35.0\n-35.1\n-35.2\n-35.3\n-35.4\n-35.5\n-35.6\n-35.7\n-35.8\n-35.9\n-36.0\n-36.1\n-36.2\n-36.3\n-36.4\n-36.5\n-36.6\n-36.7\n-36.8\n-36.9\n-37.0\n-37.1\n-37.2\n-37.3\n-37.4\n-37.5\n-37.6\n-37.7\n-37.8\n-37.9\n-38.0\n-38.1\n-38.2\n-38.3\n-38.4\n-38.5\n-38.6\n-38.7\n-38.8\n-38.9\n-39.0\n-39.1\n-39.2\n-39.3\n-39.4\n-39.5\n-39.6\n-39.7\n-39.8\n-39.9\n-40.0\n-40.1\n-40.2\n-40.3\n-40.4\n-40.5\n-40.6\n-40.7\n-40.8\n-40.9\n-41.0\n-41.1\n-41.2\n-41.3\n-41.4\n-41.5\n-41.6\n-41.7\n-41.8\n-41.9\n-42.0\n-42.1\n-42.2\n-42.3\n-42.4\n-42.5\n-42.6\n-42.7\n-42.8\n-42.9\n-43.0\n-43.1\n-43.2\n-43.3\n-43.4\n-43.5\n-43.6\n-43.7\n-43.8\n-43.9\n-44.0\n-44.1\n-44.2\n-44.3\n-44.4\n-44.5\n-44.6\n-44.7\n-44.8\n-44.9\n-45.0\n-45.1\n-45.2\n-45.3\n-45.4\n-45.5\n-45.6\n-45.7\n-45.8\n-45.9\n-46.0\n-46.1\n-46.2\n-46.3\n-46.4\n-46.5\n-46.6\n-46.7\n-46.8\n-46.9\n-47.0\n-47.1\n-47.2\n-47.3\n-47.4\n-47.5\n-47.6\n-47.7\n-47.8\n-47.9\n-48.0\n-48.1\n-48.2\n-48.3\n-48.4\n-48.5\n-48.6\n-48.7\n-48.8\n-48.9\n-49.0\n-49.1\n-49.2\n-49.3\n-49.4\n-49.5\n-49.6\n-49.7\n-49.8\n-49.9\n-50.0\n-50.1\n-50.2\n-50.3\n-50.4\n-50.5\n-50.6\n-50.7\n-50.8\n-50.9\n-51.0\n-51.1\n-51.2\n-51.3\n-51.4\n-51.5\n-51.6\n-51.7\n-51.8\n-51.9\n-52.0\n-52.1\n-52.2\n-52.3\n-52.4\n-52.5\n-52.6\n-52.7\n-52.8\n-52.9\n-53.0\n-53.1\n-53.2\n-53.3\n-53.4\n-53.5\n-53.6\n-53.7\n-53.8\n-53.9\n-54.0\n-54.1\n-54.2\n-54.3\n-54.4\n-54.5\n-54.6\n-54.7\n-54.8\n-54.9\n-55.0\n-55.1\n-55.2\n-55.3\n-55.4\n-55.5\n-55.6\n-55.7\n-55.8\n-55.9\n-56.0\n-56.1\n-56.2\n-56.3\n-56.4\n-56.5\n-56.6\n-56.7\n-56.8\n-56.9\n-57.0\n-57.1\n-57.2\n-57.3\n-57.4\n-57.5\n-57.6\n-57.7\n-57.8\n-57.9\n-58.0\n-58.1\n-58.2\n-58.3\n-58.4\n-58.5\n-58.6\n-58.7\n-58.8\n-58.9\n-59.0\n-59.1\n-59.2\n-59.3\n-59.4\n-59.5\n-59.6\n-59.7\n-59.8\n-59.9\n-60.0\n-60.1\n-60.2\n-60.3\n-60.4\n-60.5\n-60.6\n-60.7\n-60.8\n-60.9\n-61.0\n-61.1\n-61.2\n-61.3\n-61.4\n-61.5\n-61.6\n-61.7\n-61.8\n-61.9\n-62.0\n-62.1\n-62.2\n-62.3\n-62.4\n-62.5\n-62.6\n-62.7\n-62.8\n-62.9\n-63.0\n-63.1\n-63.2\n-63.3\n-63.4\n-63.5\n-63.6\n-63.7\n-63.8\n-63.9\n-64.0\n-64.1\n-64.2\n-64.3\n-64.4\n-64.5\n-64.6\n-64.7\n-64.8\n-64.9\n-65.0\n-65.1\n-65.2\n-65.3\n-65.4\n-65.5\n-65.6\n-65.7\n-65.8\n-65.9\n-66.0\n-66.1\n-66.2\n-66.3\n-66.4\n-66.5\n-66.", "timestamp": "2026-07-22T05:18:25.546003+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 9, "total_pages": 37, "image_filename": "19930085889_p9.jpg", "text": "8\nCONFIDENTIAL\nNACA RM L9F14\n\ncombination with the fuselage are compared in figure 7. The fuselage\nappeared to have very little effect on the general shapes of the lift\nand pitching-moment curves or on the aerodynamic-center location\ndetermined from the slope of the pitching-moment curve at zero lift.\nFor either the wing alone or the wing-fuselage combination, the aero-\ndynamic center was only about 1 percent of the mean aerodynamic chord\nbehind the location (27 percent of the mean aerodynamic chord) given by\nthe theory of reference 8. Apparently, for the 46.7° sweptback wing the\nforward location of the wing-fuselage juncture resulted in elimination\nof the usual unstable pitching-moment contribution of the fuselage.\nFor the wings with smaller sweep angles, the location of the wing-\nfuselage juncture was farther rearward and, in these cases, the\ncontribution of the fuselage to the pitching-moment characteristics\nseems to have been a destabilizing effect, as is normally expected.\nSuch an effect (an increase of the unstable pitching-moment contribution\nof the fuselage with a rearward shift of the wing-fuselage juncture)\nwas found in tests of midwing configurations with straight wings\nreported in reference 9. The results of reference 9 for a midwing\nconfiguration show that as the location of the quarter-chord line of\nthe wing with respect to the fuselage varied from 9 to 44 percent of\nthe fuselage length, the aerodynamic-center location of the configur-\nation varied from 0 to about 6 percent forward of the location for\nwing alone. For the 3.6° sweptback wing with fuselage, the\naerodynamic-center location (17.6 percent of the mean aerodynamic\nchord) was 7.4 percent forward of the location predicted by the theory\nof reference 8 for the wing alone.\n\nThe results presented in figure 7 show that removal of the fuselage\ncaused a reduction in lift-curve slope (from 0.062 to 0.054) near zero\nlift; but even with the fuselage removed, the lift-curve slope was\nslightly higher than the theoretical value (0.052) given in reference 8.\nThe small displacements of the lift and pitching-moment curves for the\nplain wing, relative to the curves for the wing-fuselage combination,\nprobably resulted from some camber introduced by the fairing of the\ncenter section of the wing.\n\nThe lift data presented in figure 6 indicate an increase in\nmaximum lift coefficient from 0.80 to 1.02 as the sweepback is\nincreased from 3.6° to 46.7°. This result is in agreement with the\nfindings of another low-scale investigation (reference 10) and has been\nconfirmed for Reynolds numbers as high as $12 \\times 10^6$ in a recent\ninvestigation (unpublished) of wings having geometric properties almost\nidentical to those used for the present investigation.\n\nAt lift coefficients below 0.8, the lift curves for the three\nwings are very nearly the same. Although the theories of references 8\nand 11 do predict a reduction in lift-curve slope of plain wings with\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:18:27.300371+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 33, "total_pages": 46, "image_filename": "19930085519_p33.jpg", "text": "32\nNACA RM No. L8K19\n\nPitching-moment coefficient, $C_m$\nDrag coefficient, $C_D$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\n\n| $\\delta_f$ (deg) | Flap span (percent) |\n| :--- | :--- |\n| 0 | 0 |\n| 50 | 51 |\n| 50 | 100 |\n\n[Figure: Graph showing aerodynamic characteristics with NACA logo]\n\nFigure 12.- The effect of span of the slotted flap on the aerodynamic characteristics in pitch of the 42° sweptback wing.", "timestamp": "2026-07-22T05:18:27.834230+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 62, "total_pages": 66, "image_filename": "19930082245_p62.jpg", "text": "NACA TN No. 1596\n\n$\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$\n\n.06  \n.05  \n.04  \n.03  \n.02  \n.01  \n0  \n\n(a) $c_n = 0.$\n\n$\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$\n\n.05  \n.04  \n.03  \n.02  \n.01  \n0  \n\n(c) $c_n = 0.4.$\n\nMach number, M  \n.1 .2 .3 .4 .5 .6 .7 .8 .9  \n\nTrue-contour aileron  \nBeveled-trailing-edge aileron  \n\n(b) $c_n = 0.2.$\n\n(d) $c_n = 0.6.$\n\nMach number, M  \n.1 .2 .3 .4 .5 .6 .7 .8 .9  \n\n[Figure: NACA logo in bottom-right graph]\n\nFigure 14.- Variation of $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ with Mach number for an NACA 66₁-115 airfoil section equipped with unsealed 0.20c plain ailerons.\n\n61", "timestamp": "2026-07-22T05:18:31.805377+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 19, "total_pages": 24, "image_filename": "19930085626_p19.jpg", "text": "18\nNACA RM No. L8K23\n\nCONFIDENTIAL\nTip obtained by revolving\nairfoil section around chord\nline.\n\nc/4 line\n40°\n8.48\n1.69\n11.50\n.93\n4.05\n\nSection A-A\nCircular-arc section\nnormal to c/4 line\nThickness ratio, .010\n\n$C_D$\n.06\n.04\n0\n\n$pb/2V$\n.12\n.08\n.04\n0\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\nM\n\n(c) Full-span aileron configuration. $\\delta_a = 5^\\circ$.\nFigure 4.— Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:18:33.958053+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 21, "total_pages": 33, "image_filename": "19930085544_p21.jpg", "text": "20\nNACA RM No. L8K26\n\n$$\n\\begin{array}{c}\n\\phi_o - \\phi_{90} \\quad \\text{———} \\\\\n\\phi_{270} - \\phi_o \\quad \\text{--- ---}\n\\end{array}\n$$\n\n$$\n\\left.\n\\begin{array}{c}\n\\alpha_T \\\\\n(deg) \\\\\n12\n\\end{array}\n\\right\\}\n$$\n\n$$\n\\left.\n\\begin{array}{c}\n8\n\\end{array}\n\\right\\}\n$$\n\n$$\n\\left.\n\\begin{array}{c}\n4\n\\end{array}\n\\right\\}\n$$\n\n$$\n\\left.\n\\begin{array}{c}\n2\n\\end{array}\n\\right\\}\n$$\n\n$$\n\\phi_o - \\phi_{90^\\circ} \\text{ and } \\phi_{270^\\circ} - \\phi_o\n$$\n\n$$\n\\frac{J}{\\pi \\lambda}\n$$\n\n[Figure: NACA logo]\n\nFigure 3.— Variation in amplitude of geometric angle of attack.", "timestamp": "2026-07-22T05:18:38.874440+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 15, "total_pages": 29, "image_filename": "19930085879_p15.jpg", "text": "```markdown\nNACA RM L9D11\n\ncylinder moves\nwith nose\n\npiston moves with\nrear body\n\n7\"\n\n4 1/2\"\n\nmain rocket\nhead\n\n2.109\"\n\n5.5 grams\nFFF-G powder\n\nheld rigid to airframe\nlocking spring\n\nnose section skin\n\nseparation station\n\nFigure 3.- RM-11 nose-section release mechanism.\n\n13\n```", "timestamp": "2026-07-22T05:18:46.212765+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 9, "total_pages": 31, "image_filename": "19930085881_p9.jpg", "text": "NACA RM L9D12 CONFIDENTIAL 7\n\nREFERENCES\n\n1. Sandahl, Carl A., and Marino, Alfred A.: Free-Flight Investigation of Control Effectiveness of Full-Span 0.2-Chord Plain Ailerons at High Subsonic, Transonic, and Supersonic Speeds to Determine Some Effects of Section Thickness and Wing Sweepback. NACA RM No. L7D02, 1947.\n\n2. Sandahl, Carl A.: Free Flight Investigation of Control Effectiveness of Full-Span, 0.2-Chord Plain Ailerons at High Subsonic, Transonic, and Supersonic Speeds to Determine Some Effects of Wing Sweepback, Taper, Aspect Ratio, and Section Thickness Ratio. NACA RM No. L7F30, 1947.\n\n3. Sandahl, Carl A., and Strass, H. Kurt: Additional Results in a Free-Flight Investigation of Control Effectiveness of Full-Span, 0.2-Chord Plain Ailerons at High Subsonic, Transonic, and Supersonic Speeds to Determine Some Effects of Wing Sweepback, Aspect Ratio, Taper, and Section Thickness Ratio. NACA RM No. L7L01, 1948.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:18:46.682836+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 50, "total_pages": 50, "image_filename": "19930082496_p50.jpg", "text": "NACA TN No. 1836\n49\n\n[Figure: A photograph showing a fractured surface of a ceramal-blade. A ruler labeled \"INCHES\" is placed above the blade for scale. Labels point to an \"Inner layer of oxidation\" and an \"Outer layer of oxidation\". In the bottom right corner of the image is a NACA logo with the text \"C-21268\" and \"4-22-48\".]\n\n(b) Fractured surface.\n\nFigure 13. - Concluded. Ceramal-blade failure after 12 hours and 13 minutes of operation.", "timestamp": "2026-07-22T05:18:49.128075+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 52, "total_pages": 99, "image_filename": "19930082511_p52.jpg", "text": "50\nNACA TN No. 1826\n\n$$\nP_{0n}(\\xi, \\rho) = \\frac{1}{\\pi} \\int_0^\\infty \\left[ \\frac{J_0(i\\rho q)}{iqJ_0'(iq)} \\int_a^b \\sin n\\pi \\frac{\\beta - a}{b - a} \\cos q(\\beta - \\xi) \\, d\\beta \\right.\n$$\n$$\n\\left. - \\frac{2}{q^2} \\int_a^b \\sin n\\pi \\frac{\\beta - a}{b - a} \\, d\\beta \\right] dq\n$$\n\nThe resulting infinite series for the $P_{0n}(\\xi, \\rho)$ obtained by means of contour integrations are given in appendix A.\n\nCondition (3) then becomes\n\n$$\n\\sum_{n=0}^\\infty h_{0n} \\frac{\\partial P_{0n}(\\xi, 1)}{\\partial \\xi} = - \\left. \\frac{\\partial (\\phi_0 + \\phi_c)}{\\partial \\xi} \\right|_{\\rho=1} + u \\quad (a \\le \\xi < b)\n$$\n\nBut (see appendix A) $\\lim_{\\xi \\to +\\infty} \\frac{\\partial P_{0n}}{\\partial \\xi} = - \\lim_{\\xi \\to -\\infty} \\frac{\\partial P_{0n}}{\\partial \\xi}$\n\nand so $\\lim_{\\xi \\to +\\infty} \\frac{\\partial \\phi_A}{\\partial \\xi} = - \\lim_{\\xi \\to -\\infty} \\frac{\\partial \\phi_A}{\\partial \\xi}$ so that condition (4) becomes\n\n$$\n\\lim_{\\xi \\to +\\infty} \\frac{\\partial \\phi_A}{\\partial \\xi} = 0.\n$$\n\nThus it is necessary that\n\n$$\n\\frac{h_{00}}{\\frac{\\pi}{2(b-a)}} + \\sum_{n=1}^\\infty \\frac{h_{0n} [1 - (-1)^n]}{\\frac{n\\pi}{b-a}} = 0\n$$\n\nThere exists a unique value of $u$ for which the coefficients $h_{0n}$ will satisfy this equation and it can be found as follows: Let\n\n$$\nh_{0n} = h_{0n}' + u h_{0n}''\n$$\n\nwhere\n\n$$\n\\sum_{n=0}^\\infty h_{0n}' \\frac{\\partial P_{0n}(\\xi, 1)}{\\partial \\xi} = - \\left. \\frac{\\partial (\\phi_0 + \\phi_c)}{\\partial \\xi} \\right|_{\\rho=1}\n$$", "timestamp": "2026-07-22T05:18:50.265570+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 12, "total_pages": 96, "image_filename": "19930085880_p12.jpg", "text": "```markdown\n10\n\n1.178\n\n.589\n\n.059\n\nAll longitudinal sections are circular arcs\nwith a height at the center of 5 percent of\nthe chord which forms the bottom of the\nsection.\n\nNACA\n\nFigure 3. - Details of triangular planing surface with flat bottom (model 250D).\n(All dimensions are in feet.)\n\nNACA RM No. L9C03\n```", "timestamp": "2026-07-22T05:18:54.725322+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 32, "total_pages": 58, "image_filename": "19930082617_p32.jpg", "text": "```markdown\nNACA TN 1962\n31\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X 10^3\n2 72.0 X 10^3\n3 108.0 X 10^3\n4 144.0 X 10^3\n5 180.0 X 10^3\n6 216.0 X 10^3\n7 288.0 X 10^3\n\n2.57\"\n|-|\n[Figure: Diagram of a cylinder cross-section labeled \"Band C'\" with a section view \"A-A\" showing a 45° angle]\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n20 16 12 8 4 0 -4 -6 -12 -16 -20 X 10^-4\nStrain\n\n1 2 3 4 5 6 7\nO O O O O O O\nX X X X X X X\n\n[Figure: NACA logo]\n\nFigure 20.- Strain diagram of cylinder 77. Band C'.\n```", "timestamp": "2026-07-22T05:18:55.592987+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 2, "total_pages": 23, "image_filename": "19930085906_p2.jpg", "text": "NACA RM E5F20 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nINVESTIGATION OF INTERNAL REGENERATIVE FUEL-HEATING SYSTEM\n\nFOR 20-INCH RAM JET\n\nBy Sol Baker and Eugene Perchonok\n\nSUMMARY\n\nAn investigation was conducted to evaluate the effectiveness of a simple internal regenerative fuel preheater for a 20-inch-diameter ram jet. Data obtained at subsonic sea-level conditions indicated that the fuel could be successfully preheated in this manner. The distance the preheater was located downstream of the flame holder was the primary variable affecting the final fuel temperature. Although approximately 5 minutes was required to attain a stable fuel temperature, the rate of fuel-temperature rise was maximum immediately after ignition and useful preheat temperatures ($>200^\\circ$ F) were approached within 2 minutes after ignition. The additional pressure loss caused by the introduction of the preheater in the combustion chamber may be considered negligible.\n\nINTRODUCTION\n\nResults of experiments reported in references 1 and 2 indicate that appreciable improvement in ram-jet performance and combustion efficiency can be realized by preheating the fuel. This improvement is the result of a decrease in the time required for the fuel-vaporization process preceding burning. For the investigation described in reference 1, the fuel was either regeneratively heated by being circulated through a copper coil wound around the combustion-chamber shell or was heated by circulation through a steam-heated heat exchanger mounted in the test cell. The use of the steam heat exchanger is merely an experimental technique. Although the external regenerative-heating system is useful in cooling the combustion-chamber shell, it requires an increase in engine frontal area and, if applied to a flight engine, is accompanied by an increase in engine aerodynamic drag. The additional hazard of fuel leaks also contributes to the undesirability of a fuel-preheating system in which the fuel is circulated through a passage around the combustion-chamber shell.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:19:01.176567+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 17, "total_pages": 36, "image_filename": "19930085869_p17.jpg", "text": "NACA RM L9D15\n15\n\nCONFIDENTIAL\n\n<!-- Image (101, 110, 912, 936) -->\n\nFigure 2.— General arrangement of Langley tank model 237-6SB. (All dimensions in inches.)\nCONFIDENTIAL", "timestamp": "2026-07-22T05:19:02.053377+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 59, "total_pages": 62, "image_filename": "19930082485_p59.jpg", "text": "NACA TN No. 1810\n\nJ\n\n$\\frac{C_1}{\\Delta C}$\n\n-4.0  \n-8.0  \n-16.0  \n∞  \n16.0  \n8.0  \n4.0  \n3.0  \n2.0  \n1.5  \n1.0  \n\n$\\frac{C_1}{\\Delta C}$  \n1.0  \n1.5  \n2.0  \n∞  \n3.0  \n4.0  \n8.0  \n∞  \n-8.0  \n-4.0  \n\n$C_1 n_0 / 2$\n\nFigure 15. - Chart for finding J from channel width and blade curvature.", "timestamp": "2026-07-22T05:19:06.836890+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 41, "total_pages": 66, "image_filename": "19930082914_p41.jpg", "text": "40\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:19:08.305105+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 2, "total_pages": 36, "image_filename": "19930085912_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:19:08.479932+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 30, "total_pages": 53, "image_filename": "19930082542_p30.jpg", "text": "NACA TN No. 1867\n\n29\n\nTABLE IV.- COMPARATIVE PROPERTIES OF LOW-CHROMIUM M-153 DISCS AND BAR STOCK\n\n| Material | Treatment (1) | Room-temperature properties | | | | | Rupture properties at 1200° F | | | | Data source (2) |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | Brinell hardness | Tensile strength (psi) | Offset yield strengths (psi) | | Elongation (percent) | 100 hr (psi) | 1000 hr (psi) | 100-hr elongation (percent) | | |\n| | | | | 0.02 percent | 0.2 percent | | | | | | |\n| Disc | As-forged, 2 hr at 1200° F | 235 | 120,350 | 58,750 | 72,650 | 34.5 | 52,000 | 42,000 | 11 | Reference 1 |\n| | | --- | --- | --- | 80,000 | 40 | 54,000 | 45,000 | 15 | A |\n| Disc | As-forged | 210 | 119,800 | --- | 65,000 | 44 | 46,000 | 38,500 | 9 | Reference 2 |\n| | | --- | --- | --- | 70,000 | 40 | 48,000 | 38,000 | 15 | A |\n| Bar | As-rolled | 233 | 128,500 | 72,500 | 78,500 | 40.5 | 49,500 | 37,500 | 17 | B |\n| Disc | As-forged, 24 hr at 1500° F | 232 | --- | --- | --- | --- | 47,000 | 33,500 | 36 | Reference 2 |\n| | | --- | --- | 60,000 | 70,000 | 35 | 45,000 | 36,000 | 15 | A |\n| Bar | As-rolled, 24 hr at 1500° F | 212 | 124,000 | 46,500 | 62,500 | 36 | 47,000 | 36,000 | 14 | B |\n| Disc | 2200° F W.Q.; 24 hr at 1350° F | 173 | 106,000 | --- | --- | --- | 41,500 | 36,000 | 12 | Reference 2 |\n| | | --- | --- | 40,000 | 53,500 | 38 | 44,500 | 37,000 | 12 | A |\n| Bar | 2200° F W.Q.; 24 hr at 1350° F | 205 | 120,125 | 54,000 | 64,000 | 38.5 | 53,000 | 39,500 | 15 | B |\n| Disc | 2200° F W.Q.; 24 hr at 1500° F | 207 | 114,000 | --- | 61,500 | 22.5 | 44,000 | 35,000 | 18 | Reference 2 |\n| | | --- | --- | 45,000 | 60,000 | 36 | 47,000 | 38,500 | 25 | A |\n| Bar | 2200° F W.Q.; 24 hr at 1500° F | 202 | 121,750 | 50,000 | 63,500 | 36.5 | 49,000 | 35,000 | 25 | B |\n| Disc | 2200° F W.Q.; 3-percent hot-cold-work at 1500° F; 24 hr at 1500° F | 222 | --- | --- | --- | --- | 46,000 | 34,000 | 21 | Reference 2 |\n| | | --- | --- | 52,000 | 70,000 | --- | 48,000 | 37,500 | 20 | A |\n\nW.Q., water-quenched.\n$^A$ estimated from Brinell hardness and data for bar stock in this report.\n$^B$ test data from this report.\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T05:19:13.783633+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 63, "total_pages": 66, "image_filename": "19930082245_p63.jpg", "text": "62\nNACA TN No. 1596\n\n<!-- Image (106, 100, 817, 810) -->\n\nFigure 15.- Variation of aileron hinge-moment-coefficient-curve slopes with Mach number for an NACA 66,1-115 airfoil section equipped with unsealed 0.20c plain ailerons.", "timestamp": "2026-07-22T05:19:15.084127+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 34, "total_pages": 46, "image_filename": "19930085519_p34.jpg", "text": "NACA RM No. L8K19\n33\n\nPitching-moment coefficient, $C_m$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n$\\delta_f$ (deg)\n0\n60\n\n[Figure: Graph showing aerodynamic characteristics with data points for $\\delta_f = 0$ and $\\delta_f = 60$ degrees]\n\nNACA\n\nFigure 13.- The effect of deflection of a half-span Zap flap on the aerodynamic characteristics in pitch of the 42° sweptback wing.", "timestamp": "2026-07-22T05:19:17.527901+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 20, "total_pages": 24, "image_filename": "19930085626_p20.jpg", "text": "NACA RM No. L8K23\n19\n\nCONFIDENTIAL\nTip obtained by revolving\nairfoil section around chord\nline.\n\n[Figure: Diagram of an extended-chord aileron configuration with dimensions: 5.21, 1.49, 7.00, 11.50, 1.69, 1.02, 8.48, 9.50. Angles: 40°, 10.7°. Labels: c/4 line, A, Flat sides, Section AA (Normal to c/4 line).]\n\n$C_D$\n[Graph: $C_D$ vs. $M$. Y-axis: 0, .04, .08. X-axis: .6, .8, 1.0, 1.2, 1.4, 1.6, 1.8, 2.0. Curve starts near 0.02, rises sharply around M=1.0 to ~0.06, then levels off.]\n\n$pb/2V$\n[Graph: $pb/2V$ vs. $M$. Y-axis: -.04, 0, .04, .08. X-axis: .6, .8, 1.0, 1.2, 1.4, 1.6, 1.8, 2.0. Curve starts near 0.05, drops sharply at M=1.0 to ~0.02, then gradually decreases to ~0.01. NACA logo in bottom right corner of graph.]\n\n(d) Extended-chord aileron configuration. $\\delta_a = 5^\\circ$.\n\nFigure 4.— Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:19:24.103522+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 10, "total_pages": 37, "image_filename": "19930085889_p10.jpg", "text": "NACA RM L9F14 CONFIDENTIAL 9\n\nincreased sweep angle, such a reduction, if it occurs, would be expected to be confined to a very small range of lift coefficients (from about -0.2 to 0.2) for the present models, because above a lift coefficient of 0.2 (somewhere between 0.2 and 0.3) partial separation appears to take place. The separation is indicated from the comparison of the experimental drag curves with the curve obtained by adding the drag at zero lift to the theoretical induced drag for elliptic wings of aspect ratio 4. (See fig. 6.) For each of the wings the experimental drag curve began to depart from the theoretical relation at a lift coefficient somewhere between 0.2 and 0.3. Partial flow separation above this lift coefficient therefore would be expected, which would invalidate the assumptions of the theory used to calculate the lift-curve slope. According to previous experience (see fig. 4 of reference 2, for example), the onset of flow separation, as indicated by an increase in the quantity $\\left(C_{D}-\\frac{C_{L}^{2}}{\\pi A}\\right)$, generally is accompanied by an increase in lift-curve slope for sweptback wings and a decrease in lift-curve slope for unswept wings. At the higher lift coefficients, therefore, sweptback wings may have lift-curve slopes as high or even higher than those of unswept wings of the same aspect ratio.\n\nThe results in figure 6 do not show an effect of sweep on lift-curve slope as large as was expected (on basis of references 8 and 11) for these models, even at the low lift coefficients. This difference may have resulted in part from the use of the fuselage. As has already been pointed out, removal of the fuselage caused a reduction in lift-curve slope from 0.062 to 0.054 for the 46.7° sweptback wing. Tests with straight wings (reference 9 and comparison of references 12 and 13) and tests of a 42° sweptback wing (reference 14) have indicated that the effect on the lift-curve slope of the addition of a fuselage of circular cross section depends, at least partly, on the wing geometry and on the longitudinal position of the wing-fuselage juncture. It is probable, therefore, that the usual effect of sweepback on the lift-curve slope was partially masked by a variable influence of the fuselage.\n\nIn order to determine how critically the wing characteristics were affected by changes in Reynolds number, in the range for which most of the tests had to be run, tests were made at various Reynolds numbers with and without transition strips on the leading edge of the wings. Plots to show the effect of transition strips and Reynolds numbers on the lift-curve slope and pitching-moment slope of the 46.7° sweptback wing tested alone are presented as figures 8 and 9, respectively. A summary of these results is presented in figure 10, which shows the variation of lift-curve slope and the variation of the location of the aerodynamic center with Reynolds number. Also presented in figure 10 is the theoretical value for lift-curve slope\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:19:26.520600+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 16, "total_pages": 29, "image_filename": "19930085879_p16.jpg", "text": "14\n\nseparation station\n\nattached to nose\nsection\n\nattached to thrust ring\non rear body\n\nNACA\n\nFigure 4.- Bearing plates.\n\nNACA RM L9D11", "timestamp": "2026-07-22T05:19:26.882130+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 22, "total_pages": 33, "image_filename": "19930085544_p22.jpg", "text": "NACA RM No. L9K26\n21\n\n$$\n\\epsilon = \\frac{W_{90}}{W_o} - 1 \\text{ or } \\frac{W_o}{W_{270}} - 1\n$$\n\n$$\n\\begin{array}{l}\n\\frac{W_{90}}{W_o} - 1 \\quad \\text{——} \\\\\n\\frac{W_o}{W_{270}} - 1 \\quad \\text{---}\n\\end{array}\n$$\n\n[Figure: Graph showing curves for $\\alpha_T$ (deg) values of 4, 8, 12, and 16 plotted against $J/\\pi x$ from 0 to 6, with $\\epsilon$ on the y-axis ranging from 0 to 0.16. Solid lines represent $\\frac{W_{90}}{W_o} - 1$, dashed lines represent $\\frac{W_o}{W_{270}} - 1$. NACA logo at bottom right of graph.]\n\nFigure 4.— Variation in fractional amplitude of resultant velocity.", "timestamp": "2026-07-22T05:19:27.807050+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 10, "total_pages": 31, "image_filename": "19930085881_p10.jpg", "text": "8\nCONFIDENTIAL\nNACA RM L9D12\n\nTABLE I\nPHYSICAL CHARACTERISTICS OF ALL TEST VEHICLES\n\n| | |\n| :--- | :--- |\n| Total exposed wing area, sq ft | 1.563 |\n| Aspect ratio, A | 3.71 |\n| Taper ratio | 1.00 |\n| Ratio of aileron span to exposed wing span | 1.00 |\n| Ratio of aileron span to total wing span | 0.81 |\n| Ratio of aileron chord to wing chord | 0.20 |\n| Moment of inertia about center line of test vehicle, slug-ft$^2$ | 0.10 |\n\n[Figure: NACA logo]\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:19:32.446878+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 1, "total_pages": 42, "image_filename": "19930085914_p1.jpg", "text": "RM A9D25\n\nNACA RM A9D25\n\n[Figure: NACA logo with wings]\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION  \nEMPLOYING A WING SWEPT BACK $63^\\circ$.– CHARACTERISTICS  \nTHROUGHOUT THE SUBSONIC SPEED RANGE WITH THE  \nWING CAMBERED AND TWISTED FOR A UNIFORM  \nLOAD AT A LIFT COEFFICIENT OF 0.25  \n\nBy J. Lloyd Jones and Fred A. Demele  \n\nAmes Aeronautical Laboratory  \nMoffett Field, Calif.\n\nNATIONAL ADVISORY COMMITTEE  \nFOR AERONAUTICS  \nWASHINGTON  \n\nAugust 15, 1949  \nDeclassified April 8, 1957", "timestamp": "2026-07-22T05:19:36.326772+00:00"}