PDF/ocr_dataset_1068.jsonl

{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 32, "total_pages": 47, "image_filename": "19930083221_p32.jpg", "text": "```markdown\n30\nNACA TN No. 1824\n\nIn the wake of a lifting wing the function $g^{*'}(x_1, \\mu) = 0$ and if, moreover,\n\n$$ \\int_{0}^{c_0} g^{*''}(x_1, \\mu) dx_1 = g^{*'}(c_0, \\mu) - g^{*'}(0, \\mu) = 0 $$\n\nreversal of integration in equation (51) yields the simpler expression\n\n$$ D = \\frac{\\rho_0 \\beta^2}{8\\pi^2} \\int_{0}^{2\\pi} \\sin^2 \\theta \\, d\\theta \\int_{0}^{l} \\int_{0}^{l} g^{*''}(x_1, \\mu) g^{*''}(x_2, \\mu) \\ln |x_1 - x_2| \\, dx_1 \\, dx_2 \\quad (52) $$\n\nIt is possible to draw some general conclusions from equations (46) and (52) regarding the wave drag of wings and bodies of revolution without the necessity of detailed applications to particular configurations. It is apparent immediately from equation (46a) that the wave drag of a body of revolution at zero angle of attack is independent of Mach number. This conclusion does not apply, however, to the nonlifting wing since the distribution function $f(x, \\mu)$ in equation (46) contains the variable $\\mu$ which, in turn, is a function of both $\\theta$ and $\\beta$. As $M_0$ approaches one, the study of the non-lifting wing is divided most conveniently into two parts, depending on the behavior of $f(x, \\mu)$.\n\nConsider first the more general situation in which $f(x, \\mu)$ is not zero; that is, the case in which the number of sources does not equal the number of sinks along the line $\\xi = \\text{constant}$. This means, when $M_0$ is 1, that an unequal number of sources and sinks appear in the transverse or $yz$ plane and, if equation (46) is applied, either a finite or an infinite value of drag can result. The limiting value of drag at sonic speed, obtained from integrations of surface pressures, was given by Stewart and Puckett in reference (16) for several wing plan forms, all of which had nonvanishing values of $f(x, \\mu)$. If the pressure distribution is calculated, however, the local pressure coefficients are seen to become infinitely large as sonic speed is reached, even for the body of revolution, so that the assumptions of the linear theory are violated and the reliability of the drag predicted by equation (46) can in no case be assessed even though the predicted values remain finite. Equation (43) shows also that when control-surface methods are used to compute drag at $M_0 = 1$,\n```", "timestamp": "2026-07-22T06:31:34.377129+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 42, "total_pages": 46, "image_filename": "19930082613_p42.jpg", "text": "NACA TN 1938\n41\n\n[Figure: Microstructure of dead-soft Inconel. Etchant, 10-percent sodium cyanide, electrolytic; magnification, X750.]\n\n(a) Not treated (dead soft).\n\n(b) After heating at $900^\\circ$ F for 1 hour and air-cooling.\n\nFigure 15. - Microstructure of dead-soft Inconel. Etchant, 10-percent sodium cyanide, electrolytic; magnification, X750.", "timestamp": "2026-07-22T06:31:35.255826+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 34, "total_pages": 72, "image_filename": "19930085491_p34.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 33\n\n(c) The difference between the theoretical and experimental values of maximum lift-drag ratio was found to be a result of higher values of both minimum drag coefficient and drag due to lift. These higher values of drag as well as the large experimental center-of-pressure travel were associated with relatively large areas of separated boundary-layer flow.\n\n2. The following effects of Reynolds number were observed in tests with the $63^\\circ$ swept-back wing configuration:\n\n(a) Increasing the Reynolds number to 0.84 million increased the maximum lift-drag ratio to 7.2 and reduced the total center-of-pressure travel to approximately 12 percent of the mean aerodynamic chord.\n\n(b) The improvement in maximum lift-drag ratio resulted from decreases in both minimum drag coefficient and drag due to lift. These reductions as well as the decrease in total center-of-pressure travel with lift coefficient were attributed to reductions in the areas of separated flow as the Reynolds number was increased.\n\n3. Tests at a Mach number of 1.53 and Reynolds number of 0.62 million of four additional sweep angles of $57.0^\\circ$, $60.4^\\circ$, $67.0^\\circ$, and $69.9^\\circ$ obtained by rotating the wing panels about the midpoint of the root chord afforded the following conclusions:\n\n(a) A maximum lift-drag ratio of 7.1 was obtained at the optimum leading-edge sweep angle of $67^\\circ$. The optimum leading-edge sweep angle resulted from the opposing effects of increasing sweep in decreasing the minimum drag coefficient and in increasing the drag due to lift.\n\n(b) The effect of sweep in decreasing the minimum drag coefficient was associated with the decrease in wing pressure drag resulting from the increased angle behind the Mach cone and the decreased streamwise thickness-chord ratio. The increase in drag due to lift with increasing sweep was primarily due to the decrease in lift-curve slope.\n\n(c) The total center-of-pressure travel increased with increase in sweep angle but no abrupt changes in pitching-moment characteristics were found as the complement of the trailing-edge sweep angle became less than the Mach angle for a Mach number of 1.53.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:31:35.940873+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 16, "total_pages": 92, "image_filename": "19930085870_p16.jpg", "text": "NACA RM No. L9D07\nCONFIDENTIAL\n15\n\nline between upper and lower surfaces with increase in angle of attack\nis quite obvious. At 4.20° angle of attack, for example, the initially\nsteep adverse pressure gradient on the lower surface favors transition\nimmediately aft of the ridge line while the lower and more uniform\nadverse pressure gradient on the upper surface would, by comparison,\nindicate a delay in transition. The liquid-film tests have shown this\nto be the actual result. At the 60.3-percent-semispan station similar\ntrends in the pressure distributions occur. However, the position of\nthe steep adverse pressure gradients on upper and lower surfaces indi-\ncate that the point of transition on the lower surface would be nearer\nthe ridge line than was the case at the inboard station and, conversely,\nthe point of transition on the upper surface would be further removed\nfrom the ridge line. As before, the liquid-film tests exhibit such a\npattern. Thus, the characteristics of the chordwise pressure distri-\nbution with varying angle of attack bear out the liquid-film observa-\ntions in regard to the curvature of the shocks arising on the wing\nsurfaces and their position.\n\nThe pressure distributions for wedge-leading-edge wing 11 indicate\nthat the adverse pressure gradient originates immediately aft of the Mach\nlines from the ridge-line apex, except at the outboard station where the\ntest results show the pressure rise to begin aft of the ridge line. The\npressure distributions indicate the same effects as shown for wing 5, an\nappreciable forward movement of the shocks arising on the lower surface\nand little rearward shift of the shocks on the upper surface. At the\n22.5-percent-semispan station it is interesting to note the change in\nshape of the curve ahead of the ridge line for the upper surface at\n10.75° angle of attack. Although the initial wedge angle of the wing\nstill produces a positive angle with respect to stream direction, the\ninitial negative pressure followed by a positive pressure, both points\nahead of the ridge line, may possibly be due to the detached shock and\nthe resulting subsonic nature of the flow accompanied by the tendency\nof the high pressure on the lower surface to relieve itself by flow\naround the leading edge and over the upper surface.\n\nThe pressure distributions for elliptical-leading-edge wing 11 show\nsimilar trends to the wedge-leading-edge wing though not quite so marked.\nA delay in the transition point as shown by the liquid-film tests would\nbe expected from the very gradual rise of the adverse pressure gradient.\nThe difference in location of the shocks on the wing surface with change\nin angle of attack is still evident on the curves.\n\nGeneral Remarks\n\nIt appears that the peaks and breaks in the curves of this paper\ncalculated by the linear theory will not in most instances be realized\nexperimentally. The theoretical pressure-distribution curves for the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:31:38.738930+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 2, "total_pages": 52, "image_filename": "19930085911_p2.jpg", "text": "NACA RM E9F22 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nFREE-FLIGHT PERFORMANCE OF 16-INCH-DIAMETER SUPERSONIC RAM-JET UNITS\n\nI - FOUR UNITS DESIGNED FOR COMBUSTION-CHAMBER-INLET MACH NUMBER OF 0.12 AT FREE-STREAM MACH NUMBER OF 1.6 (UNITS A-2, A-3, A-4, AND A-5)\n\nBy William W. Carlton and Wesley E. Messing\n\nSUMMARY\n\nFree-flight investigations have been conducted on four 16-inch-diameter ram-jet units to determine the performance at high subsonic and supersonic velocities. The units were released from an airplane at high altitudes. The engine thrust and the force of gravity accelerated the ram-jet units to high subsonic and supersonic Mach numbers. Data for evaluating the performance were obtained from radio-telemetering and radar-tracking equipment.\n\nThe effects of free-stream Mach number and gas total-temperature ratio on diffuser total-pressure recovery, thrust coefficient, and external drag coefficient are correlated. Also included are the performance data of the individual ram-jet units for a range of free-stream Mach numbers from 0.38 to 1.73 and for gas total-temperature ratios between 1.0 and 6.6.\n\nA maximum combustion efficiency of 91 percent occurred in one unit at a free-stream Mach number of 1.70, with a diffuser total-pressure recovery of 0.90. The corresponding gas total-temperature ratio of 5.1 was equivalent to an exhaust-gas total temperature of $4050^\\circ$ R. A net acceleration (excluding gravity) of 2.0 g's and a net thrust coefficient of 0.56 were produced.\n\nINTRODUCTION\n\nAs part of an extensive study of the performance of ram jets, the NACA Lewis laboratory is conducting a free-flight investigation\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:31:44.209076+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 11, "total_pages": 33, "image_filename": "19930085900_p11.jpg", "text": "CONFIDENTIAL\n\n(a) Chine configuration.\n\nSta. 0\n10.00\n18.00\n26.00\n34.00\n42.22\n\n(b) Multiple step configuration, V's pointed forward.\n\nFigure 2.— Bottom views of model showing location of chines and multiple steps.\n\nCONFIDENTIAL\n\nNACA\n\nNACA RM L9D20\n\n10", "timestamp": "2026-07-22T06:31:44.430519+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 25, "total_pages": 60, "image_filename": "19930085862_p25.jpg", "text": "NACA RM No. L9A07\n23\n\n<!-- Image (274, 157, 810, 457) -->\n\n$\\delta_s$\n(deg)\n$\\circ$ -25\n$\\circ$ 0\n$\\square$ 25\n\n<!-- Image (274, 542, 810, 797) -->\n\n(c) $C_L$ and $C_m$ against $\\alpha$.\nFigure 5.- Concluded.", "timestamp": "2026-07-22T06:31:46.716506+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 52, "total_pages": 78, "image_filename": "19930082483_p52.jpg", "text": "50\nNACA TN No. 1807\n\nAt $180^\\circ$ admission, a corrected rotor speed of 8650 rpm, and a total-pressure ratio of 2.0,\n\n(power observed)$_{360^\\circ}$ = 365 hp, from figure 8\n(pumping loss)$_{360^\\circ}$ = 0, from figure 9(c)\n(bearing loss)$_{360^\\circ}$ = 6.9 hp, from figure 9(b)\n(pumping loss)$_{180^\\circ}$ = 1.5 hp, from figure 9(c)\n(driving-fluid loss)$_{180^\\circ}$ = 19.8 hp, from figure 9(d)\n\nTherefore\n\n(estimated power)$_{180^\\circ}$ = $\\frac{1}{2}$ (365 + 6.9) - (6.9 + 1.5) - 19.8\n= 157.8 hp\n\nBecause all quantities used were corrected to standard sea-level conditions, the estimated power for $180^\\circ$ is at standard sea-level conditions.\n\nEstimation of Efficiency at $180^\\circ$ Admission\n\nThe over-all efficiency that may be expected at a given degree of admission may be calculated from equation (33) once the power for the same operating conditions has been estimated for that degree of admission.\n\nAt $180^\\circ$ admission, equation (33) becomes", "timestamp": "2026-07-22T06:31:48.061702+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 31, "total_pages": 149, "image_filename": "19930083192_p31.jpg", "text": "NACA TN 1976\n27\n\nof the wing, then at peak acceleration in a sharp gust, the wing has traveled about 4 chords while the fuselage has traveled 1. On this basis, the lift on the fuselage might be considered zero or at most about half of its steady-flow value. If this hypothesis is compared with the material given in reference 19 (some of the values were in error and have been recomputed), the results shown in table VI are obtained. The results in table VI indicate that, although Jones' unsteady functions are much too high on the basis of gross area and the functions from reference 4 are somewhat high, the use of a net wing area plus one-half the fuselage intercept brings all calculations into closer agreement. In fact, for the infinite-aspect-ratio functions, the discrepancies are less than the precision of the data. If the net wing area is assumed, then both finite- and infinite-aspect-ratio functions differ from experiment by about the same amount, the results based on reference 14 being high and those based on reference 4 being low. Similar corrections may apply to Keuthe's results (fig. 27). Although the evidence indicates that the unsteady-lift functions for infinite aspect ratio should be used for conventional airplanes in conjunction with the net wing area plus half the fuselage intercept, the use of net area is recommended and is discussed subsequently.\n\nUnpublished tests of a skeleton airplane equipped with a wing swept back $45^\\circ$ showed that the installation of the fuselage had no appreciable effect on the maximum acceleration increment. In this particular case, however, the length of the wing from the leading edge of the root section to the trailing edge of the tip section was almost equal to the length of the fuselage. Since one effect of sweep would be to modify the rate of development of lift on the wing, the difference between the lift developed on the wing and that on the fuselage for a given gust penetration may have been too small to be noted during the test. Although the evidence is still scant and conflicting in some respects, the use of net wing area appears to be better than the use of gross wing area for the sharp-edge gust except when the side projections of the wing and the fuselage are about the same length.\n\nThe determination of the proper wing area for the gradient gust is complicated by the introduction of the pitching motion of the airplane, and comparisons must be made between the detailed calculations and experiment. In the investigations that have been made (reference 15) the results have indicated that the use of the net wing area yields the best over-all agreement between calculations and experiment for gradient distances between 0 and 16 chords.\n\nThe use of tapered and swept wings for modern aircraft leads to problems in the computation of the gust load factor. Relatively little information is available on either problem, but the data of reference 19 indicated that moderate amounts of taper (up to 2:1) have little or no effect and can be neglected at least until experimental evidence of a", "timestamp": "2026-07-22T06:31:49.335776+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 27, "total_pages": 31, "image_filename": "19930085859_p27.jpg", "text": "```markdown\nNACA RM No. L9B25\n\nWing alone\nWing-fuselage\n$C_L = 0$\n\n<!-- Image (33, 117, 909, 765) -->\n\nTail height, $h_t$, percent semispan\n\nFigure 11.- Variation of downwash gradient with tail height and Mach number for a model with $35^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\n50\n```", "timestamp": "2026-07-22T06:31:50.743591+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 1, "total_pages": 34, "image_filename": "19930085913_p1.jpg", "text": "NACA RM L9F24\nFILE COPY\nNO. 9\nRESTRICTED\nCopy\nRM L9F24\n228\n\nNACA\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION OF THE EFFECTS OF SWEEPBACK\nON THE FLUTTER OF A UNIFORM CANTILEVER WING WITH\nA VARIABLY LOCATED CONCENTRATED MASS\nBy Herbert C. Nelson and John E. Tomassoni\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nAUTHORITY CROWLEY CHANGE #1903\nDATE 12-11-53\nCLASSIFIED DOCUMENT\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\nT.C.F.\n\nRETURN TO THE ABOVE ADDRESS\n\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\n\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nAugust 31, 1949\n\nRESTRICTED", "timestamp": "2026-07-22T06:31:52.404779+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 24, "total_pages": 32, "image_filename": "19930085847_p24.jpg", "text": "22\nCONFIDENTIAL\nNACA RM A9D04\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (252, 101, 716, 883) -->\n\n(e) M = 0.80 ($C_L$ = 0.14).\nFigure 7- Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:31:55.166396+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 12, "total_pages": 29, "image_filename": "19930085899_p12.jpg", "text": "NACA RM No. L9J21\n\nFloating-tail geometry\nTwice semispan area 0.0178 sq ft\nAspect ratio 4.0\nTaper ratio 0.60\n\nChord plane\n$\\alpha = 0^\\circ$\n\n1.0\n1.0\n1.0\n1.0\n\nSection B-B\n$\\frac{1}{16}$\nDouble scale\n\nStation A\n.453\n(MAC).80\n.25(MAC) Model\nB\nB\n.25-chord line\n.60\n1.60\n.76\nBump surface .10\n$45^\\circ$\n$\\frac{1}{8}$ Diameter\nPivot center\n\n0 1 2\nScale, inches\n\nNACA\n\nFigure 3.- Details of free-floating tails used in surveys behind model with $45^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\n11", "timestamp": "2026-07-22T06:32:01.067372+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 43, "total_pages": 46, "image_filename": "19930082613_p43.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:32:01.263560+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 62, "total_pages": 114, "image_filename": "19930086061_p62.jpg", "text": "58\nNACA RM L9J07\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n-3 -3\n-2 -2\n-1 P -1\n0 0\n1 1\n\n(a) $\\psi = 0^\\circ$\n\nUpper\nLower\n\nLeft semispan\nRight semispan\n\n-3 -3\n-2 -2\n-1 P -1\n0 0\n1 1\n$10^\\circ$\n\n(b) $\\psi = 10^\\circ$\n\nNACA\n\nFigure 18.- Pressure distribution about wing 2 at various angles of yaw;\n$\\alpha = 14.1^\\circ$.", "timestamp": "2026-07-22T06:32:02.770878+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 13, "total_pages": 26, "image_filename": "19930085890_p13.jpg", "text": "12\nNACA RM No. E9C11\n\nin that the small tube containing the cap blew apart, resulting in\na slight bulging of the large brass container. In each experiment\nwith the tetranitromethane and nitrobenzene mixture, which is known\nto be sensitive to the shock produced by a number 6 cap, the entire\nassembly disintegrated with the explosion of the cap. The results\nof the shock tests with gaseous and liquid diborane were similar to\nthose obtained with the cap alone or with the cap and water or with\nthe cap and alcohol.\n\nWith the limited number of sensitivity determinations made\nwith diborane, no explosions nor detonations were produced by heat\nor by shock.\n\nREFERENCES\n\n1. Huff, Vearl N., Calvert, Clyde S., and Erdmann, Virginia C.:\nTheoretical Performance of Diborane as a Rocket Fuel. NACA\nRM No. E8I17a, 1949.\n\n2. Rowe, William H., Ordin, Paul M., and Diehl, John M.: Investigation of the Diborane - Hydrogen Peroxide Propellant\nCombination. NACA RM No. E7K07, 1948.\n\n3. Malina, Frank J.: Characteristics of the Rocket Motor Unit\nBased on the Theory of Perfect Gases. Jour. Franklin Inst.,\nvol. 230, no. 4, Oct. 1940, pp. 433-454.\n\n4. Laubengayer, A. W., Ferguson, R. P., and Newkirk, A. E.: The\nDensities, Surface Tensions and Parachors of Diborane, Boron\nTriethyl and Boron Tribromide: The Atomic Parachor of Boron.\nJour. Am. Chem. Soc., vol. 63, no. 2, Feb. 1941, pp. 559-561.\n\n5. Anon.: Handbook of Chemistry and Physics. Charles D. Hodgman,\ned., Chem. Rubber Pub. Co. (Cleveland), 29th ed., 1945,\npp. 425, 1764.", "timestamp": "2026-07-22T06:32:06.785593+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 12, "total_pages": 33, "image_filename": "19930085900_p12.jpg", "text": "NACA RM L9D20\n\nCONFIDENTIAL\n\n(c) Multiple step configuration, V's pointed aft.\n\nSta. 0\n10.00\n18.00\n26.00\n34.00\n42.22\n\n(d) Multiple step configuration, transverse.\n\nFigure 2.- Concluded.\nCONFIDENTIAL\n\n11", "timestamp": "2026-07-22T06:32:08.424894+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 36, "total_pages": 37, "image_filename": "19930082646_p36.jpg", "text": "NACA TN 1980\n35\n\n<!-- Image (163, 172, 874, 392) -->\n\nWarped forebody and extended afterbody ———\nBasic forebody and basic afterbody - - - - - -\n\n<!-- Image (163, 450, 889, 688) -->\n\nFigure 18.- Variation of maximum and minimum trim and rise with wave length,\nfor landings in waves 4 feet high.", "timestamp": "2026-07-22T06:32:10.291793+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 33, "total_pages": 47, "image_filename": "19930083221_p33.jpg", "text": "NACA TN No. 1824\n31\n\nthe x component of induced velocity increases indefinitely when $f(x,\\mu)$ is not zero and that the theory is, therefore, no longer consistent.\n\nIn the very special second case, that is, when $f(x,\\mu)$ vanishes for all values of $\\theta$, the analysis just presented breaks down at equation (40). It is clear, however, that in this case there are equal numbers of sources and sinks in the $\\xi = \\text{constant}$ plane and the behavior of the flow field at infinity is, therefore, exactly the same as that which would have been produced by a distribution of doublets. Equations (49) and (50) give the velocities induced at infinity by an arbitrary doublet distribution. These induced velocity components are, in terms of $\\beta$, one degree higher than the similar components for the nonlifting case. The values of both $u$ and $v_n$ can thus be expected to approach zero for all values of $M_0$ as $r$ approaches infinity for any flow field generated entirely by doublets or by an equal number of sources and sinks. It follows then that the linearized theory for lifting surfaces (generated entirely by doublets) and for bodies with thickness distributions such that $f(x,\\mu)$ vanishes (generated by an equal number of sources and sinks in all $\\xi = \\text{constant}$ planes) is entirely consistent as $M_0$ approaches one and, in particular, for $M_0$ equal to one. This being true, it follows immediately from equation (52) that the wave drag of a lifting system is zero at sonic speed.\n\nThickness solutions at $M_0 = 1$.— A swept-back wing of constant chord and infinite aspect ratio is an example of a practical aerodynamic shape for which an equal number of sources and sinks occur in every $yz$ plane. (See fig. 13.) Consider the case in which the wing cross section is diamond shaped with a slope equal to $\\lambda$ in a plane normal to the leading edge. Then, in a transverse plane, (section BB of fig. 13) $w_0$ equals $\\pm V_0 \\lambda \\cos \\psi$, the minus and plus signs applying, respectively, to the left and right of the ridge line. Accordingly, the solution of the problem can be written in terms of a distribution of sources, thus\n\n$$\n\\begin{aligned}\n\\Phi = & - \\frac{1}{2\\pi} \\int_{x \\cot \\psi}^{x} \\frac{c_0}{\\left(x - \\frac{c_0}{2 \\cos \\psi}\\right) \\cot \\psi} V_0 \\lambda \\cos \\psi \\ln \\left[(y-y_1)^2 + z^2\\right] dy_1 \\\\\n& + \\frac{1}{2\\pi} \\int_{x \\cot \\psi}^{x + \\frac{c_0}{2 \\cos \\psi}} \\cot \\psi \\quad V_0 \\lambda \\cos \\psi \\ln \\left[(y-y_1)^2 + z^2\\right] dy_1 \\quad (53)\n\\end{aligned}\n$$", "timestamp": "2026-07-22T06:32:11.674566+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 32, "total_pages": 46, "image_filename": "19930085548_p32.jpg", "text": "NACA RM No. EBL30\n31\n\n1077\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:32:11.897274+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 35, "total_pages": 72, "image_filename": "19930085491_p35.jpg", "text": "34 CONFIDENTIAL NACA RM No. A8J04\n\n(d) At the optimum leading-edge sweep angle of $67^\\circ$, increasing the Reynolds number to 0.95 million resulted in a value of maximum lift-drag ratio of 7.4.\n\nIn all cases where it was possible to compare experimental values of lift, drag, and pitching moment with those calculated by the linear theory, the experimental values were, respectively, lower, higher, and less stable than those indicated by theory. These differences were due to both the low scale of test and the partial exclusion of viscous effects in the theory. The experimental and theoretical trends with sweep, however, were in good agreement.\n\nBecause of the influence of the adverse lifting pressure gradients that caused boundary-layer separation close to the leading edges of the wings in the present study, the theoretical values of maximum lift-drag ratio may not be realized at full scale with this wing. These results indicate that the use of camber and wing twist may be necessary as a means of reducing the gradient to improve the boundary-layer flow if the maximum value of lift-drag ratio is to be attained.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nREFERENCES\n\n1. Jones, Robert T.: Estimated Lift-Drag Ratios at Supersonic Speed. NACA TN No. 1350, 1947.\n\n2. Jones, Robert T.: Thin Oblique Airfoils at Supersonic Speed. NACA TN No. 1107, 1946.\n\n3. McCormack, Gerald M., and Walling, Walter C.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.—Investigation of a Large-Scale Model at Low Speed. NACA RM No. A8D02, 1948.\n\n4. Reynolds, Robert M., and Smith, Donald W.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back $63^\\circ$.—Subsonic Mach and Reynolds Number Effects on the Characteristics of the Wing and on the Effectiveness of an Elevon. NACA RM No. A8D20, 1948.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:32:13.188115+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 3, "total_pages": 52, "image_filename": "19930085911_p3.jpg", "text": "```markdown\n2\nCONFIDENTIAL\nNACA RM E9F22\n\nof 16-inch-diameter ram-jet units. The units are released from an\nairplane at high altitudes and accelerated to supersonic velocities\nby the engine thrust and the force of gravity.\n\nThe purpose of the investigation is to provide performance\ndata on full-scale units operating under actual atmospheric condi-\ntions at high subsonic and supersonic Mach numbers. In addition to\nsubsonic and supersonic data obtainable from wind-tunnel research,\nthe flight investigation provides data throughout the transonic\nrange. Data are also being obtained under conditions of rapid\nacceleration with accompanying changes in inlet conditions due to\nlarge variations in altitude and Mach number.\n\nThe investigation is being conducted off the Virginia coast\nnear the NACA Langley laboratory. Four ram-jet designs (designated\n16-A, 16-B, 16-C, and 16-D) of different inlet and outlet diameters\nare used in order to obtain data over a range of combustion-chamber\nvelocities. Data are obtained at different values of fuel-air ratio\nby presetting the fuel regulator. Continuous data records are\nobtained by radio-telemetering and radar-tracking equipment during\nthe flight.\n\nData obtained from the first ram-jet unit investigated (desig-\nnated 16-A-1) are discussed in reference 1. Data obtained with the\nsucceeding four A-type ram-jet units are presented herein. Time\nhistories of the performance are presented for altitudes between\n36,000 feet and sea level and free-stream Mach numbers from 0.38 to\n1.73. Also included are the effects of free-stream Mach number and\ngas total-temperature ratio on diffuser total-pressure recovery,\nthrust coefficient, and external drag coefficient.\n\nInsufficient data are available from the four ram-jet units\ndiscussed herein to permit correlation of the variables affecting\ncombustion efficiency. Time histories of these variables are there-\nfore presented, showing only simultaneous values.\n\nAPPARATUS\n\nThe ram-jet unit consisted of an outer shell with four stabi-\nlizing fins at the rear and a centrally located body in the diffuser\nsection that housed the telemetering equipment and the fuel system.\nThe gross weight of each unit was approximately 525 pounds. A ram-\njet unit suspended from an airplane is shown in figure 1. A cutaway\nview of a typical ram jet is shown in figure 2.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:32:19.882204+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 25, "total_pages": 32, "image_filename": "19930085847_p25.jpg", "text": "NACA RM A9D04 CONFIDENTIAL 23\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (273, 100, 748, 874) -->\n\n(f) M = 0.81 (C_L = 0.17).\n\nFigure 7.- Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:32:22.556057+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 13, "total_pages": 29, "image_filename": "19930085899_p13.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:32:22.869871+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 17, "total_pages": 92, "image_filename": "19930085870_p17.jpg", "text": "16 CONFIDENTIAL NACA RM No. L9D07\n\nwings of angular or abrupt ridge line are possibly an exception. Much of the discrepancy between test and theoretical values may be attributed to two factors omitted in the linear theory: viscosity and shocks resulting from second-order compressibility effects. Certainly the presence of the shocks observed on the wing surfaces and their movement with angle of attack influence the lift and drag results. The transition line in the boundary layer is obviously determined by the position of these shocks and the associated adverse pressure gradient. It follows that a greater or lesser turbulent area will affect the drag accordingly. Thus the lower minimum drag of the elliptical-leading-edge wings for values of $\\tan \\epsilon / \\tan m$ less than 1.6 may be attributed to their lesser areas of turbulent boundary layer. Furthermore it appears that, regardless of whether the leading edge is supersonic, until complete attachment of the shock is realized along the wing leading edge, the flow at or near the leading edge is physically similar to the flow over two-dimensional wings at high subsonic Mach numbers. At the lower values of $\\tan \\epsilon / \\tan m$ it is possible that an increased lift may be experienced at the leading edge of sufficient magnitude to raise the total lift above the predicted theoretical value. Of course at extremely low values of $\\tan \\epsilon / \\tan m$ such an effect would diminish. At the larger values of $\\tan \\epsilon / \\tan m$ the effect of boundary layer and shock interaction may be blamed for the reduced lift with respect to theory; but as $\\tan \\epsilon / \\tan m$ approached the value for complete attachment of the shock to the leading edge, the transonic nature of the flow in the vicinity of the ridge line would give way to entirely supersonic flow and the actual lift would be expected to attain a value somewhat near the theoretical. It is possible that a wing having a sharp leading edge and a ridge line of easy curvature might retain the smaller region of turbulent boundary layer associated with the elliptical-leading-edge series. This configuration would also favor early attachment of the leading-edge shock with the consequent higher lift and lower drag exhibited by the wedge-leading-edge series at values of $\\tan \\epsilon / \\tan m$ much greater than 1.\n\nCONCLUSIONS\n\nSupersonic tests at Mach numbers of 1.62, 1.92, and 2.40 of 22 triangular wings having 8 percent thickness ratio, an 18-percent location of maximum-thickness point, and representing two leading-edge configurations, wedge and elliptical, for each apex angle indicate the following conclusions:\n\n1. For a given wing series the ratio of the actual lift-curve slope to the theoretical two-dimensional value was, for any given ratio of the tangent of the vertex half-angle to the tangent of the Mach angle ($\\tan \\epsilon / \\tan m$), relatively independent of Mach number.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:32:23.762827+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 44, "total_pages": 46, "image_filename": "19930082613_p44.jpg", "text": "NACA TN 1938\n43\n\n[Figure: Microstructure image showing grain boundaries and precipitates]\n\n(c) After annealing at $1600^\\circ$ F for 2 hours and air-cooling.\n\n[Figure: Microstructure image showing grain boundaries and precipitates]\n\nNACA\nC-22634\n12-9-48\n\n(d) After annealing at $2200^\\circ$ F for 2 hours and air-cooling.\n\nFigure 15. - Concluded. Microstructure of dead-soft Inconel. Etchant, 10-percent sodium cyanide, electrolytic; magnification, X750.", "timestamp": "2026-07-22T06:32:26.319924+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 26, "total_pages": 60, "image_filename": "19930085862_p26.jpg", "text": "24\nNACA RM No. L9A07\n\n$C_{Na}$\n.6\n.4\n.2\n0\n-.2\n-.4\n\n$\\delta_a$\n(deg)\n$\\nabla$ -15\n$\\diamond$ -6\n$\\circ$ 0\n$\\triangle$ 6\n$\\Delta$ 15\n\n$C_n$\n.01\n0\n-.01\n\n$C_l$\n.02\n.01\n0\n-.01\n-.02\n\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\nNACA\n\n(a) $C_l$, $C_n$, and $C_{Na}$ against $\\alpha$.\nFigure 6.— Aileron characteristics of wing with split flaps.", "timestamp": "2026-07-22T06:32:26.585709+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 2, "total_pages": 34, "image_filename": "19930085913_p2.jpg", "text": "NACA RM L9F24\nRESTRICTED\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nRESEARCH MEMORANDUM\nEXPERIMENTAL INVESTIGATION OF THE EFFECTS OF SWEEPBACK\nON THE FLUTTER OF A UNIFORM CANTILEVER WING WITH\nA VARIABLY LOCATED CONCENTRATED MASS\nBy Herbert C. Nelson and John E. Tomassoni\nSUMMARY\nThe results obtained from 95 subsonic flutter tests which were\nconducted in the Langley 4.5-foot flutter research tunnel on untapered\ncantilever wings with sweepback angles of $0^\\circ$, $45^\\circ$, and $60^\\circ$ and carrying\na single concentrated weight are presented. The weight used throughout\nthe series of tests was 14 percent heavier than each wing. A primary\npurpose of the investigation was to present experimental information to\nbe used as a basis for evaluating analytical procedures for determining\nthe flutter speed of weighted sweptback wings.\nThe weight was mounted at a series of spanwise positions on the\nleading edges and on the midchord lines of the wings. The results of\nthe tests in which the wings were weighted at the leading edge indicated that the flutter speed was greatly affected by the spanwise\nposition of the weight and, in these cases, the change in sweepback\ndid not appreciably alter the flutter speed. For the cases in which\nthe wings were weighted at the midchord, an increase in sweepback\ngenerally caused an increase in the flutter speed and, as the sweep\nangle was increased, the effect of the spanwise weight position became\nmore pronounced. The results are presented in the form of plots of\nflutter speed and frequency as a function of spanwise weight position\nfor the sweepback angles tested.\nINTRODUCTION\nThe purpose of this paper is to present experimental data on the\nflutter characteristics of sweptback untapered cantilever wings\ncarrying concentrated weights. These data were obtained from 95\nflutter tests conducted in the Langley 4.5-foot flutter research tunnel\non wings, each carrying a single weight at a series of spanwise\nRESTRICTED", "timestamp": "2026-07-22T06:32:26.782921+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 63, "total_pages": 114, "image_filename": "19930086061_p63.jpg", "text": "NACA RM L9J07\n59\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\n-3\n-2\n-1\n0\n1\n-3\n-2\n-1\n0\n1\n20°\n(c) $\\psi = 20^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\n-3\n-2\n-1\n0\n1\n-3\n-2\n-1\n0\n1\n35°\n(d) $\\psi = 35^\\circ$\n\nFigure 18.- Concluded.", "timestamp": "2026-07-22T06:32:30.237685+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 13, "total_pages": 33, "image_filename": "19930085900_p13.jpg", "text": "12\n\nStation 0 10.00 CONFIDENTIAL 42.22\n\nBottom view\n\nStrip\n\n2.50\" R\n\n45° 45°\n\n1\"\n16\n\nMaximum section\n\nNACA\n\nFigure 3.— Size of strips relative to model.\nCONFIDENTIAL\n\nNACA RM L9D20", "timestamp": "2026-07-22T06:32:30.448468+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 53, "total_pages": 78, "image_filename": "19930082483_p53.jpg", "text": "```markdown\nNACA TN No. 1807\n51\n\n$$\n\\eta'_{180^\\circ} = \\frac{1}{2} \\left( \\frac{P_{180^\\circ}}{P_{360^\\circ}} \\right) \\eta'_{360^\\circ}\n$$\n\n$$\nP_{180^\\circ} = 157.6 \\text{ hp}\n$$\n\n$$\nP_{360^\\circ} = 365 \\text{ hp from figure 8}\n$$\n\n$$\n\\eta'_{360^\\circ} = 0.755 \\text{ from figure 8}\n$$\n\n$$\n\\eta'_{180^\\circ} = \\frac{1}{2} \\left( \\frac{157.6}{365} \\right) 0.755\n$$\n\n$$\n= 0.653 \\text{ estimated efficiency}\n$$\n\nBy comparison with figure 11, it may be seen that the estimated efficiency 0.653 is in close agreement with the value actually obtained by test for $180^\\circ$ admission.\n\n### REFERENCES\n\n1. Kent, Robert Thurston: Mechanism and Mechanics. Kent's Mechanical Engineers' Handbook, sec. 8. John Wiley & Sons, Inc., 11th ed., 1938, p. 21.\n2. Goudie, William G.: Steam Turbines. Longmans, Green and Co. (London), 2d ed., 1922, p. 535.\n3. Culver, E. P.: Investigation of a Simple Form of Hydraulic Dynamometer. Mech. Eng., vol. 59, no. 10, Oct. 1937, pp. 749-753.\n4. Stodola, A.: Steam and Gas Turbines. Vol. I. McGraw-Hill Book Co., Inc., 1927, pp. 199-200, 201, 221. (Reprinted, Peter Smith (New York), 1945.)\n5. Moore, Charles S., Biermann, Arnold E., and Voss, Fred: The NACA Balanced-Diaphragm Dynamometer-Torque Indicator. NACA RB No. 4C28, 1944.\n```", "timestamp": "2026-07-22T06:32:33.516735+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 34, "total_pages": 47, "image_filename": "19930083221_p34.jpg", "text": "32\nNACA TN No. 1824\n\n<!-- Image (101, 99, 843, 447) -->\n\nFigure 13.- Views of infinite swept wing showing coordinates.\n\nThe value of $\\partial\\phi/\\partial x$ can immediately be found to be\n$$\n\\frac{\\partial\\phi}{\\partial x} = \\left(\\frac{V_o\\lambda \\cos \\psi}{2\\pi \\tan \\psi}\\right) \\times\n$$\n$$\n\\ln \\frac{\\left\\{\\left[y-\\left(x+\\frac{c_o}{2 \\cos \\psi}\\right)\\frac{1}{\\tan \\psi}\\right]^2+z^2\\right\\} \\left\\{\\left[y-\\left(x-\\frac{c_o}{2 \\cos \\psi}\\right)\\frac{1}{\\tan \\psi}\\right]^2+z^2\\right\\}}{\\left\\{\\left[y-\\left(\\frac{x}{\\tan \\psi}\\right)\\right]^2+z^2\\right\\} \\left\\{\\left[y-\\left(\\frac{x}{\\tan \\psi}\\right)\\right]^2+z^2\\right\\}} \\quad (54)\n$$\nfrom which it is apparent that as $r = \\sqrt{y^2+z^2}$ becomes infinitely large, $\\partial\\phi/\\partial x$ approaches zero. In the plane of the airfoil, that is, for $z = 0$, $\\partial\\phi/\\partial x$ becomes", "timestamp": "2026-07-22T06:32:35.845860+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 34, "total_pages": 62, "image_filename": "19930082918_p34.jpg", "text": "NACA TN 1940\n\n33\n\nTABLE 5\n\nRUPTURE CHARACTERISTICS AT 1200° F OF LOW-CARBON S-195 ALLOY\n\nSOLUTION-TREATED 10 HOURS AT 2000° F, WATER-QUENCHED,\n\nAND AGED AS INDICATED\n\n| Aging temperature (°F) | Aging time (hr) | Stress (psi) | Rupture time (hr) | Elongation (percent) | Reduction of area (percent) | Maximum true strain (1) |\n|---|---|---|---|---|---|---|\n| 1400 | 0.5 | 70,000 | 0.27 | 23.8 | 22.6 | 0.126 |\n| | | 60,000 | 1.21 | 14.5 | 16.8 | .091 |\n| | 1.0 | 75,000 | .22 | 26.1 | 28.2 | .165 |\n| | | 70,000 | .38 | 20.6 | 22.6 | .126 |\n| | | 65,000 | 1.12 | 15.9 | 19.7 | .109 |\n| | | 60,000 | 2.71 | 13.2 | 19.0 | .109 |\n| | | | 2.78 | ---- | 20.5 | .114 |\n| | | 50,000 | 22.00 | 11.6 | 18.0 | .100 |\n| | 3.0 | 75,000 | .37 | 24.2 | 26.8 | ---- |\n| | | 60,000 | 5.65 | 12.7 | 16.2 | ---- |\n| | 10.0 | 78,000 | .21 | 27.7 | 27.5 | .162 |\n| | | 70,000 | 1.45 | 22.0 | 21.8 | .122 |\n| | | 60,000 | 7.80 | 19.1 | 16.2 | .087 |\n| | | | 10.05 | 13.0 | 13.1 | .069 |\n| | | | 6.25 | ---- | 15.4 | .084 |\n| | | 50,000 | 29.5 | 10.2 | 11.0 | .058 |\n| | 30.0 | 70,000 | 1.83 | 24.6 | 28.2 | ---- |\n| | | 60,000 | 15.73 | 23.4 | 28.8 | ---- |\n| | 100.0 | 80,000 | .05 | 30.6 | 43.7 | .285 |\n| | | 75,000 | .81 | 27.1 | 29.5 | .174 |\n| | | 60,000 | 2.73 | 22.9 | 31.5 | .194 |\n| | | | 32.5 | 21.0 | 26.8 | .157 |\n| | | | 8.0 | ---- | 32.2 | .195 |\n| | | 50,000 | 63.0 | 23.5 | 24.5 | .140 |\n| | 1000.0 | 75,000 | 1.21 | 26.5 | 37.9 | .240 |\n| | | 60,000 | 10.5 | 21.7 | 34.8 | .213 |\n| | | | 6.75 | ---- | 36.6 | .227 |\n| | | 50,000 | 82.75 | 27 | 30.7 | .182 |\n| | | | 59.5 | 27 | 34.2 | .208 |\n| 1600 | .5 | 70,000 | 2.25 | 19.8 | 23.4 | .131 |\n| | | | 1.81 | 21.6 | 20.5 | .113 |\n| | | 60,000 | 10.00 | ---- | 16.8 | .092 |\n| | 1.0 | 75,000 | .34 | 23.0 | 34.2 | .207 |\n| | | 65,000 | 3.73 | 20.0 | 22.6 | .128 |\n| | | 60,000 | 9.12 | 12.3 | 15.4 | .083 |\n| | | | 7.83 | ---- | 17.6 | .096 |\n| | | 50,000 | 41.0 | 4.5 | 10.7 | .058 |\n| | 10.0 | 65,000 | 4.38 | 26.1 | 29.5 | .174 |\n| | | 60,000 | 10.83 | 22.5 | 26.2 | .150 |\n| | | | 6.25 | ---- | 21.1 | .120 |\n| | | 50,000 | 49.75 | 15.7 | 21.3 | .120 |\n| | 100.0 | 75,000 | .45 | 35.4 | 39.8 | .254 |\n| | | 70,000 | 1.25 | 29.4 | 34.2 | .207 |\n| | | 65,000 | 4.57 | 27.5 | 33.2 | .199 |\n| | | 60,000 | 12.50 | 30.8 | 35.6 | .219 |\n| | | | 8.0 | ---- | 38.6 | .242 |\n| | | 50,000 | 80 | 26.2 | 32.3 | .192 |\n| | 1000.0 | 73,000 | .78 | 26.8 | 38.6 | .243 |\n| | | 60,000 | 13.5 | 33.8 | 42.9 | .279 |\n| | | | 6.33 | ---- | 44.6 | .296 |\n| | | 50,000 | 57.5 | 45.0 | 30.0 | .178 |\n| Unaged | | 70,000 | .083 | 19.9 | 23.4 | .131 |\n| | | 65,000 | .192 | 16 | 21.8 | .122 |\n| | | 60,000 | .85 | 14 | 16.8 | .092 |\n| | | 55,000 | 34.25 | 10.8 | 11.7 | .062 |\n| | | 50,000 | 74.0 | 6.1 | 10.9 | .058 |\n| | | | 23.0 | 9.2 | ---- | ---- |\n\n(1) $\\epsilon_{max} = \\log_e \\frac{A_0}{A}$, at fracture section.\n\nNACA", "timestamp": "2026-07-22T06:32:36.328147+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 49, "total_pages": 65, "image_filename": "19930082546_p49.jpg", "text": "48\nNACA TN No. 1870\n\nPressure, p, dynes/cm$^2$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:32:36.718909+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 37, "total_pages": 37, "image_filename": "19930082646_p37.jpg", "text": "36\nNACA TN 1980\n\n2° amplitude of porpoising\nElevator deflection, deg\n-35 -30 -25 -20 -15 -10 -5 0\nUnstable Stable\n\nGreatest cycle of oscillation\nTrim, deg\n16 12 8 4 0\nRise, ft\n2 0 -2\n\nCenter of gravity, percent M.A.C.\n20 22 24 26 28 30 32 34 36 38 40\n(a) Stable take-off range.\n\nMaximum\nMinimum\n\nMaximum\nMinimum\n\nContact trim, deg\n3 4 5 6 7 8 9 10 11 12\n(b) Landing behavior.\nUnstable\nUpper limit\ndecreasing trim\nUpper limit\ndecreasing trim\nLower limit\nUnstable Stable\n\nTrim, deg\n8 6 4 2\nSpeed, mph\n40 50 60 70 80 90 100\n(d) Trim limits of stability.\n\nSpray in propellers\nHeavy\nLight\nPropellers clear\nClear\n\nSpray on flaps\nLight\nFlaps clear\nClear\n\nGross load, percent of design gross load\n90 100 110 120 130\nSpeed, mph\n10 20 30 40 50 10 20 30 40 50\n(c) Effect of gross load on spray.\n\nFigure 19.-Summary chart of principal hydrodynamic qualities of a flying boat having a hull of high length-beam ratio, a warped forebody, and an extended afterbody. Gross load, 75,000 pounds, power loading, 11.5 pounds per brake horsepower, wing loading, 41.1 pounds per square foot; flap deflection, 20°.\n\nNACA-Langley - 11-7-49 - 875", "timestamp": "2026-07-22T06:32:40.261450+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 14, "total_pages": 29, "image_filename": "19930085899_p14.jpg", "text": "NACA RM No. L9A21\n\n[Figure: Photograph of a model aircraft wing with sweptback design and free-floating tails, mounted on a test surface. The NACA logo is visible in the lower right corner of the image.]\n\nFigure 4.— Photograph of model with $45^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil showing free-floating tails.\n\n13", "timestamp": "2026-07-22T06:32:42.762307+00:00"}
{"citation_id": "19930085847", "source_url": "https://ntrs.nasa.gov/api/citations/19930085847/downloads/19930085847.pdf", "page_number": 26, "total_pages": 32, "image_filename": "19930085847_p26.jpg", "text": "24\nCONFIDENTIAL\nNACA RM A9D04\n\nDrag-producing area\nThrust-producing area\n\n<!-- Image (252, 100, 720, 497) -->\n\n<!-- Image (252, 508, 616, 837) -->\n\n(g) M = 0.82 ($C_L = 0.21$).\n\nFigure 7.- Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:32:44.217065+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 45, "total_pages": 46, "image_filename": "19930082613_p45.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:32:48.569607+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 4, "total_pages": 52, "image_filename": "19930085911_p4.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 3\n\nThe four ram-jet units investigated were designated 16-A-2, 16-A-3, 16-A-4, and 16-A-5. (The 16-A refers to the maximum diameter and the model design and the numeral is the unit number.) Model A was designed for a combustion-chamber-inlet velocity of 165 feet per second (Mach number, 0.12) at a free-stream Mach number of 1.60 and a gas total-temperature ratio of 4.0. This heat addition is equivalent to operation at a fuel-air ratio of 0.067 and a combustion efficiency of 60 percent at sea-level altitude. The diffuser was a single oblique-shock type with no internal contraction. The spike cone angle was 50° and the diffuser was designed for a normal shock at the inlet lip at a free-stream Mach number of 1.60 and a combustion-chamber-inlet Mach number of 0.12. The lip of the outer shell was positioned to intercept the oblique shock at a free-stream Mach number of 1.80. A schematic cross-sectional diagram of a ram-jet unit, including the dimensions for model A, is presented in figure 3.\n\nThe fuel system (fig. 4) included a fuel tank, a fuel regulator, and a fuel-spray ring. Helical tubing was coiled inside the fuel tank to store helium at a pressure of 3200 pounds per square inch. Fuel was stored in a flexible synthetic-rubber fuel cell that has a capacity of $8\\frac{1}{2}$ gallons. The fuel used was 73-octane gasoline (AN-F-23a). Free-stream total pressure actuated the fuel regulator and controlled the pressure of helium on the fuel cell. This helium pressure forced the fuel into three fuel lines, each of which contained a spring-loaded reducing valve and a separate set of spray nozzles. Only one set of these nozzles operated at the start of each flight, which permitted the use of high fuel pressures at low fuel-flow rates. As the free-stream total pressure increased, the regulator increased the pressure in the fuel tank and thus provided greater rates of fuel flow. Increased fuel pressure successively opened the second and third reducing valves, which brought more nozzles into operation at the desired values of free-stream total pressure. The regulator could be adjusted to alter the fuel pressure and different reducing valve springs could be used to change the opening pressure of each set of nozzles.\n\nA ducted-airfoil-type flame holder with intermediate gutters (fig. 5) was used in units A-2, A-3, and A-4. Two electrically ignited magnesium flares mounted upstream of the flame holder initiated the combustion process. This type of flame holder was previously investigated in a test stand at low combustion-chamber-inlet velocities and at pressures corresponding to nearly sea-level altitude. Ram-jet unit A-5 employed a rake-type flame holder with\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:32:51.247993+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 36, "total_pages": 72, "image_filename": "19930085491_p36.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 35\n\n5. Vincenti, Walter G., Nielsen, Jack N., and Matteson, Frederick H.: Investigation of Wing Characteristics at a Mach Number of 1.53. I – Triangular Wings of Aspect Ratio 2. NACA RM No. A7I10, 1947.\n\n6. Van Dyke, Milton D.: Aerodynamic Characteristics Including Scale Effect of Several Wings and Bodies Alone and in Combination of a Mach Number of 1.53. NACA RM No. A6K22, 1947.\n\n7. Haack, W.: Geschossformen Kleinsten Wellenwiderstandes Bericht 139 der Lilienthal Gesselschaft.\n\n8. Grey, W.E.: A Simple Visual Method of Recording Boundary Layer Transition (Liquid Film) Tech. Note Aero. 1816, R.A.E. (British/U.S. Restricted), Aug. 1946.\n\n9. Perkins, Edward W.: Experimental Investigation of the Effects of Support Interference on the Drag of Bodies of Revolution at a Mach Number of 1.5. NACA RM No. A8B05, 1948.\n\n10. Cohen, Doris: The Theoretical Lift of Flat Swept-Back Wings at Supersonic Speeds. NACA TN No. 1555, 1948.\n\n11. Kleissas, John: Charts of the Zero-Lift Drag of Supersonic Swept Back Wings for Various Taper Ratios. Northrop Aircraft Inc., Rep. No. GM-109, AM-51, Sept. 1947.\n\n12. Vincenti, Walter G., Van Dyke, Milton D., and Matteson, Frederick H. Investigation of Wing Characteristics at a Mach Number of 1.53. II – Swept Wings of Taper Ratio 0.5. NACA RM No. A8E05, 1948.\n\n13. Sauer, R.: Method of Characteristics for Three-Dimensional Axially Symmetrical Supersonic Flows. NACA TM No. 1133, 1947.\n\n14. Chapman, Dean R., and Perkins, Edward W.: Experimental Investigation of the Effects of Viscosity on the Drag of Bodies of Revolution at a Mach Number of 1.5. NACA RM No. A7A31a, 1947.\n\n15. Stewart, H.J.: The Lift of a Delta Wing at Supersonic Speeds. Quart. App. Math., vol. IV, no. 3, Oct. 1946, pp. 246-254.\n\n16. von Kármán, T.H., and Millikan, C.B.: On the Theory of Laminar Boundary Layers Involving Separation. NACA Rep. No. 504, 1934.\n\n17. Jones, Robert T.: Effect of Sweepback on Boundary Layer and Separation. NACA TN No. 1402, 1947.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:32:52.333130+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 27, "total_pages": 60, "image_filename": "19930085862_p27.jpg", "text": "NACA RM No. L9A07\n25\n\n$P_R$\n.8\n.4\n0\n-.4\n-.8\n\n$\\delta_a$ (deg)\n$\\triangle$ -15\n$\\diamond$ -5\n$\\circ$ 0\n$\\nabla$ 5\n$\\square$ 15\n\n$C_{ha}$\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\n(b) $C_{ha}$ and $P_R$ against $\\alpha$.\nFigure 6.— Continued.", "timestamp": "2026-07-22T06:32:53.452527+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 64, "total_pages": 114, "image_filename": "19930086061_p64.jpg", "text": "60\nNACA RM L9J07\n\n<!-- Image (37, 109, 874, 874) -->\n\nFigure 19.- Pressure distribution about wing 2 at various angles of yaw;\n$\\alpha = 24.1^\\circ$.", "timestamp": "2026-07-22T06:32:54.762076+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 66, "total_pages": 98, "image_filename": "19930086073_p66.jpg", "text": "```markdown\n64\n\nLift coefficient, $C_L$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:32:55.418892+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 32, "total_pages": 149, "image_filename": "19930083192_p32.jpg", "text": "28\nNACA TN 1976\n\nmore accurate nature is available. Large values of taper (for example,\nof about 3:1) and large sweep angles ($30^\\circ$ to $40^\\circ$) lead to considerable\nconcern as to the adequacy of the unsteady-lift functions, particularly\nthe Küssner function $C_{L_g}$. The concern arises from the fact that for\nwings of high taper, the root section with the larger chord will have a\nrate of development of lift for a given distance penetrated into a gust\nconsiderably less than that of the narrower tip chord. In the case of\nsweep, the root of a sweptback wing penetrates the gust first and the\nlift may develop an appreciable value before the tip sections ever enter\nthe gust.\n\nWith regard to the effect of wing taper on the unsteady-lift\nfunction, no theoretical developments are available to permit accurate\ncomputation, and calculations have been made by utilizing strip theory\nto develop the unsteady-lift functions for the finite wing from the two-\ndimensional functions of Küssner and Wagner. Experimental data available\nto check the effect of taper on the unsteady-lift functions are indirect\nand are the result of tests of two specific airplanes. In both cases,\ncalculations based on the results presented in reference 4 were in agree-\nment with the experimental data for the sharp-edge gust in which pitch\ncan be neglected. In the case of the large flying boat with $3\\frac{1}{2}:1$ taper\nratio, the discrepancy amounted to about 2 percent. The maximum dis-\ncrepancy was well within the experimental error. On the basis of these\nlimited results, the effect of taper, at least up to $3\\frac{1}{2}:1$, appears to be\nnegligible insofar as the calculation of total loads is concerned.\n\nNo theoretical studies are available for the sweptback wing, but\nrecent test results for a straight and a $45^\\circ$ sweptback wing are available.\nTests were made on a straight wing with a 2:1 taper and on the equivalent\nof the straight wing where each half-wing was rotated about the midchord\npoint at the root so that the span changed with the angle of sweep. The\nsweep was such that the midchord line was at an angle of $45^\\circ$ to its\noriginal position. Flights were made through a sharp-edge gust, and an\naverage time history of acceleration increments as a fraction of the\nmaximum value is shown in figure 29. For comparison with the experi-\nmental results, two curves are shown, one based on the theory presented\nin reference 4 which disregards the effect of sweep on the unsteady-lift\nfunctions and the other represents the calculated time history of accele-\nration based on the assumption that strip theory could be applied to derive\nthe Küssner function for the swept wing.\n\nFigure 29 shows that strip theory is in good agreement with experi-\nment throughout the entire history, but the neglect of the effect of\nsweep on the lag in lift is in error at the start of the motion, although\ngood agreement is obtained for distances greater than 4 chords. Both", "timestamp": "2026-07-22T06:32:56.023646+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 28, "total_pages": 31, "image_filename": "19930085859_p28.jpg", "text": "```markdown\n26\n\nM = 0.70\n$\\alpha = 10^\\circ$\nM = 0.80\n$\\alpha = 10^\\circ$\nM = 0.85\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\nTail height, $h_t$, percent semispan\n\nNACA\n\nFigure 12.- Dynamic-pressure surveys in region of tail plane for a model with $35^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\nNACA RM No. L9B25\n```", "timestamp": "2026-07-22T06:32:56.112609+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 14, "total_pages": 33, "image_filename": "19930085900_p14.jpg", "text": "CONFIDENTIAL\n\nNACA RM L59D20\n\nTrim indicator\n\nDashpot piston linkage\n(secured to model)\n\nFlexible rubber hose\n\nPressure gauge connection\n\nDashpot cylinder\n(fixed to staff)\n\nAir inlet connection\n\nNACA\nL-53357.2\n\nFigure 4.— Model mounted for testing.\nCONFIDENTIAL\n\n13", "timestamp": "2026-07-22T06:32:56.413588+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 14, "total_pages": 26, "image_filename": "19930085890_p14.jpg", "text": "```markdown\nNACA RM No. E9C11\n\nTABLE I - SUMMARY OF PERFORMANCE OF LIQUID DIBORANE AND LIQUID OXYGEN IN\nROCKET ENGINE\n\n| Chamber pressure (lb/sq in. abs.) | Thrust (lb) | Total propellant flow (lb/sec) | Specific impulse, I (lb-sec/lb) | Ratio of fuel weight to total propellant weight | Characteristic length, L* (in.) | Specific impulse, corrected (lb-sec/lb) | Volume specific impulse ($\\frac{\\text{lb-sec}}{\\text{cu ft}} \\times \\frac{1}{62.4}$) | Characteristic velocity, C* (ft/sec) | Thrust coefficient $C_F$ |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Injection system, eight hole** | | | | | | | | | |\n| 315 | 88.9 | 0.402 | 221 | 0.581 | 325 | 223 | 123 | 5960 | 1.20 |\n| 343 | 105.0 | .425 | 247 | .267 | 325 | 261 | 190 | 6120 | 1.30 |\n| 315 | 89.2 | .360 | 248 | .349 | 325 | 280 | 174 | 6560 | 1.20 |\n| 290 | 88.0 | .359 | 245 | .329 | 325 | 272 | 175 | 6150 | 1.28 |\n| 295 | 85.3 | .346 | 247 | .391 | 325 | 287 | 171 | 6480 | 1.22 |\n| 285 | 96.1 | .387 | 248 | .401 | 325 | 266 | 164 | 5610 | 1.43 |\n| 290 | 95.5 | .438 | 218 | .219 | 325 | 240 | 178 | 5040 | 1.39 |\n| 297 | 99.8 | .435 | 229 | .253 | 325 | 247 | 179 | 5200 | 1.42 |\n| 305 | 101.6 | .417 | 244 | .414 | 325 | 258 | 159 | 5560 | 1.41 |\n| 310 | 104.2 | .422 | 247 | .388 | 325 | 262 | 166 | 5580 | 1.42 |\n| 285 | 96.1 | .416 | 232 | .392 | 159 | 251 | 150 | 5220 | 1.42 |\n| 295 | 97.0 | .411 | 244 | .408 | 159 | 259 | 155 | 5470 | 1.39 |\n| 295 | 95.4 | .408 | 234 | .372 | 159 | 237 | 155 | 5510 | 1.37 |\n| **Injection system, four hole** | | | | | | | | | |\n| 290 | 97.2 | 0.389 | 250 | 0.350 | 325 | 273 | 170 | 5680 | 1.42 |\n| 290 | 95.4 | .367 | 260 | .400 | 325 | 274 | 167 | 6020 | 1.39 |\n| 305 | 99.5 | .415 | 240 | .367 | 325 | 259 | 160 | 5600 | 1.38 |\n\nNACA\n13\n```", "timestamp": "2026-07-22T06:32:58.724825+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 18, "total_pages": 92, "image_filename": "19930085870_p18.jpg", "text": "NACA RM No. 19D07 CONFIDENTIAL 17\n\n2. The experimental lift-curve slopes for both the elliptical- and wedge-leading-edge configurations were essentially the same, but slightly higher than theory for wings with leading edges well behind the Mach cone. With the Mach cone in the vicinity of the leading edge, the lift-curve slopes were considerably lower than theory. With the leading edge well ahead of the Mach cone the wedge-leading-edge configuration approached very close to the theoretical two-dimensional lift-curve slope.\n\n3. Except for cases with the Mach cone well behind the leading edge, the elliptical-leading-edge configuration gave lower minimum drag. This advantage was attributed to the lesser area of turbulent boundary layer on these wings.\n\n4. The linear theory applied to the wedge-leading-edge series was quite inadequate for prediction of the drag.\n\n5. The maximum lift-drag ratios for the elliptical-leading-edge configuration were higher up to a value of $\\tan \\epsilon / \\tan m$ equal approximately to 1.3, from which point the wedge-leading-edge configuration exhibited the greater value.\n\n6. The location of center of pressure was relatively independent of Mach number for a given wing series and approached the center of area. An essentially linear variation of location of center of pressure with $\\tan \\epsilon / \\tan m$ occurred with the over-all travel being approximately 10 percent. For the elliptical-leading-edge wings the center of pressure lay 3 to 4 percent ahead of its location for the wedge-leading-edge wings.\n\n7. Any leading-edge suction achieved by the elliptical-leading-edge wings was evidently of such magnitude as to be overshadowed by other effects.\n\n8. The position of shocks arising on the wing surfaces, the line of boundary-layer transition, and the steep adverse pressure gradient were found to be practically coincident.\n\n9. The agreement of the theoretical with experimental pressure distributions was much better for the wing of subsonic leading edge than for the wing having supersonic leading edge.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:32:59.643719+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 54, "total_pages": 78, "image_filename": "19930082483_p54.jpg", "text": "52\nNACA TN No. 1807\n\n6. Gabriel, David S., Carman, L. Robert, and Trautwein, Elmer E.:\nThe Effect of Inlet Pressure and Temperature on the Efficiency\nof a Single-Stage Impulse Turbine Having an 11.0-Inch Pitch-\nLine Diameter Wheel. NACA ACR No. E5E19, 1945.\n\n7. Yates, A. H.: 'Carpets' and 'Lattices'. Aircraft Eng.,\nvol. XVIII, no. 203, Jan. 1946, pp. 8-9.\n\n8. Keenan, Joseph H., and Kaye, Joseph: Thermodynamic Properties\nof Air. John Wiley & Sons, Inc., 1945.\n\n9. Allen R. C.: Steam-Turbine Blading. Trans. A.S.M.E., vol. 62,\nno. 8, Nov. 1940, pp. 689-705; discussion, pp. 705-710.", "timestamp": "2026-07-22T06:33:02.029765+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 15, "total_pages": 29, "image_filename": "19930085899_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:33:02.824010+00:00"}