PDF/ocr_dataset_1034.jsonl

{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 13, "total_pages": 96, "image_filename": "19930085880_p13.jpg", "text": "NACA RM No. 19C03\n\nTowing\ncarriage\n\nGondola housing\ntowing gear and\nobservers\n\nTowing staff\n\nTrimming-moment\ndynamometer\n\nTypical model\n(250 A)\n\nNACA\nL-58507\n\nFigure 4. - Photograph of test set-up.\n\n11", "timestamp": "2026-07-22T05:19:38.165884+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 36, "total_pages": 78, "image_filename": "19930082618_p36.jpg", "text": "```markdown\n34\n\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-1.6\n-2.0\n\nSection lift coefficient, $c_l$\n\n.1\n0\n-.1\n-.2\n\nMoment coefficient, $c_{m_{c/4}}$\n\n-24 -16 -8 0 8 16 24\nSection angle of attack, $\\alpha_0$, deg\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\nabla$ 2.0\n$\\triangle$ 2.5\n$\\triangleright$ 3.0\n$\\triangleleft$ 4.0\n$\\blacktriangle$ 9.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\nFigure 5.- Aerodynamic characteristics of the NACA 64$_1$-012 airfoil section, 24-inch chord.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:19:46.015294+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 53, "total_pages": 99, "image_filename": "19930082511_p53.jpg", "text": "NACA TN No. 1826\n51\n\nend\n\n$$\n\\sum_{n=0}^{\\infty} h_{0n} r \\frac{\\partial P_{0n}(\\xi, 1)}{\\partial \\xi} = 1\n$$\n\nthen if\n\n$$\n\\frac{h_{00}' + uh_{00}''}{\\frac{\\pi}{2(b-a)}} + \\sum_{n=1}^{\\infty} \\frac{(h_{0n}' + uh_{0n}'') [1 - (-1)^n]}{\\frac{n\\pi}{b-a}} = 0\n$$\n\n$$\nu = - \\frac{2h_{00}' + \\sum_{n=1}^{\\infty} \\frac{h_{0n}'}{n} [1 - (-1)^n]}{2h_{00}'' + \\sum_{n=1}^{\\infty} \\frac{h_{0n}''}{n} [1 - (-1)^n]}\n$$\n\nThe coefficients $h_{0n}'$ and $h_{0n}''$ are found by solution of sets of simultaneous linear equations, as described in the previous section.\n\nThe function $\\phi_A''' = \\sum h_{0n}''' P_{0n}(\\xi, \\rho)$ is the perturbation potential which, when added to that of a uniform flow, gives the potential of the disturbance-free expanding tunnel described in part I and indicated in figure 2. The corresponding perturbation velocities $\\frac{\\partial \\phi_A'''}{\\partial \\xi}$ have equal and opposite values at $\\infty$ and $-\\infty$.\n\nEXAMPLE\n\nAs a somewhat simplified illustration, the problem of a semi-infinite unit doublet distribution (degenerate horseshoe vortex) along the tunnel axis was considered. The tunnel was assumed to have an open jet, 3 tunnel radii in length. The tunnel interference was calculated for four different positions of the upstream end of the doublet distribution, these positions being 0.1, 0.4, 0.7, and 1.0 radii downstream from the entrance. If the upstream end is taken as the origin of the coordinate system, then (see appendix C)\n\n$$\nr_1^{(1)}(\\xi) = - \\frac{1}{(1 + \\xi^2)^{3/2}} - \\frac{2}{\\pi} \\int_0^{\\infty} \\frac{J_1(iq)}{iJ_1'(iq)} [qK_0(q) + K_1(q)] \\cos q\\xi \\, dq\n$$\n\n$$\nr_m^{(j)}(\\xi) \\equiv 0 \\quad \\quad (m \\neq 1, \\ j \\neq 1) \\tag{33}\n$$", "timestamp": "2026-07-22T05:19:47.710893+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 18, "total_pages": 36, "image_filename": "19930085869_p18.jpg", "text": "16\nNACA RM L9D15\n\nCONFIDENTIAL\n\n[Figure: Hull lines diagram with labels AP, 21, 15, 13, 11.25, 1, 2, 4, 6, 8, Chine]\n\n[Figure: Side view of hull with labels Unswept hull, Chine, FP, 1, 2, 4, 6, 8, 11.25, 13, 15, 21, AP, and NACA logo]\n\nFigure 3.- Hull lines of Langley tank model 237-6SB.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:19:47.910620+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 33, "total_pages": 58, "image_filename": "19930082617_p33.jpg", "text": "```markdown\n32\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\n\n[Figure: Cross-section diagram showing Band B, L-A, and A-A with dimensions 2.57\" and 45° angles]\n\nDistance from horizontal diameter, in.\nStrain\n\n| Distance from horizontal diameter, in. | 20 | 16 | 12 | 8 | 4 | 0 | -4 | -8 | -12 | -16 | -20 X 10^-4 |\n|----------------------------------------|----|----|----|---|---|---|----|----|-----|-----|-------------|\n| 10                                     | O  | O  | O  | O | O |   |    |    |     |     |             |\n| 9                                      | X  | X  | X  | X | X |   |    |    |     |     |             |\n| 8                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 7                                      | X  | X  | X  | X | X |   |    |    |     |     |             |\n| 6                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 5                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 4                                      |    |    |    | X | X | X | X  | X  |     |     |             |\n| 3                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 2                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 1                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 0                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 1                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 2                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 3                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 4                                      |    |    |    | O | O | O | O  | O  |     |     |             |\n| 5                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 6                                      |    |    |    | 1 | 2 | 3 | 4  | 5  | 6   |     |             |\n| 7                                      |    |    |    | X | X | X | X  | X  | X   |     |             |\n| 8                                      |    |    |    |   |   |   |    |    |     |     |             |\n| 9                                      |    |    |    | X | X | X | X  | X  | X   |     |             |\n| 10                                     |    |    |    | O | O | O | O  | O  | O   |     |             |\n\n[NACA logo]\n\nFigure 21.- Strain diagram of cylinder 78. Band B.\n```", "timestamp": "2026-07-22T05:19:50.807453+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 60, "total_pages": 62, "image_filename": "19930082485_p60.jpg", "text": "3.5\n1.2\nK/J 1.1\n3.0\n2.5\n1.0\nK/J\n2.0\n1.5\n1.0\n0\n.2\n.4\n.6\n.8\n1.0\nC₁ n₀/2\nC₁\nΔC\n-4.0\n-8.0\n-16.0\n∞\n16.0\n8.0\n4.0\n3.0\n2.0\n1.5\n1.0\nC₁\nΔC\n1.0, -4.0\n-8.0\n1.5\n2.0\n4.0\n∞\n0\n.1\n.2\n.3\nC₁ n₀/2\nC₁\nΔC\n-4.0\n-8.0\n∞\n16.0\n8.0\n4.0\n3.0\n2.0\n1.5\n1.0\nNACA\nFigure 16. - Chart for finding K/J from channel width and blade curvature.", "timestamp": "2026-07-22T05:19:55.032701+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 64, "total_pages": 66, "image_filename": "19930082245_p64.jpg", "text": "NACA TN No. 1596\n63\n\n<!-- Image (92, 110, 906, 999) -->\n\n(a) $\\delta_a = -4^\\circ$\nFigure 16.— Aileron pressure distribution for an NACA 66,1-115 airfoil section equipped with unsealed 0.20c plain ailerons. $\\alpha = 1^\\circ$.", "timestamp": "2026-07-22T05:19:55.783922+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 3, "total_pages": 23, "image_filename": "19930085906_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM E9F20\n\nA series of experiments was therefore undertaken to evalute\na simple light-weight internal regenerative fuel-preheating\nsystem that does not require a change in ram-jet external\ndimensions and is applicable to aircraft. The study was made at\nthe NACA Lewis laboratory with a 20-inch-diameter ram jet\noperating at simulated subsonic sea-level flight conditions.\nThe performance of the engine and the internal preheater was\ninvestigated for several positions of the preheater in the\ncombustion chamber. At each position, the variation of fuel\ntemperature with time was determined from ignition until stable\nfuel temperature was reached.\n\nAPPARATUS\n\nA schematic diagram of the circular, 20-inch-diameter, steady-\nflow ram jet used in this investigation is shown in figure 1. All\npertinent dimensions are given on the figure. Both the combustion\nchamber and the exhaust nozzle were water-cooled.\n\nBecause A.S.T.M. distillation curves may vary for different batches\nof the same fuel, the distillation curve for the fuel used (AN-F-28)\nis shown in figure 2. After passing through the preheater, the fuel\nwas injected in the liquid phase in an upstream direction near the\ndiffuser inlet. The fuel injector consisted of six equally spaced\n1/4-inch steel tubes in an 80°-V pattern with the base of the\nV 5 inches downstream of the diffuser inlet. A total of 68 No. 70\nholes were drilled on the upstream side of the fuel bars. These\norifices were equally spaced along the six bars; however, no holes\nwere drilled within 2 inches of the diffuser wall.\n\nThe flame holder (fig. 3) was a modification of the annular-V\ntype-B burner described in reference 3. This flame holder consisted\nof two concentric, 30°-V, 3-inch-chord, perforated annuli of 16-inch\nand 9-inch diameters connected by eight radial gutters. A 30°\nperforated center cone was supported by two radial gutters extending\nto the inner annulus. As originally designed, fuel was injected in\nthe V of the flame holder; however, for this investigation, the fuel\nnozzles were removed. A spark plug was used to ignite the burner.\n\nThe regenerative fuel preheater, made of coiled 3/4-inch\nInconel tubing, was located in the combustion chamber immediately\ndownstream of the flame holder. The design of this preheater was\nbased on preliminary experiments that indicated the length and the\nsize of fuel-preheating path required for useful preheat temper-\natures. The preheater weighed 9 pounds and consisted of three\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:19:55.980508+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 3, "total_pages": 36, "image_filename": "19930085912_p3.jpg", "text": "NACA RM No. E9C16 RESTRICTED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION OF HOT-GAS BLEEDBACK\n\nFOR ICE PROTECTION OF TURBOJET ENGINES\n\nII - NACELLE WITH LONG STRAIGHT AIR INLET\n\nBy Edmund E. Callaghan and Robert S. Ruggeri\n\nSUMMARY\n\nAerodynamic and icing investigations were conducted in the NACA Lewis icing research tunnel on a two-thirds-scale model of a turbojet-engine nacelle with a long straight air inlet in order to provide basic design criterions for hot-gas bleedback systems. An investigation of a hot-gas bleedback system consisting of several orifices peripherally located around the inlet opening was conducted for both dry-air and icing conditions. General rules for obtaining a satisfactory orifice configuration are presented.\n\nThe most uniform temperature distribution was obtained with a bleedback of 4.4 percent at a gas temperature of $1000^\\circ$ F and resulted in an average dry-air-temperature rise of $46^\\circ$ F. The maximum deviation from the average air-temperature rise for this condition was $6^\\circ$ F. Satisfactory agreement between calculated and measured heat requirements for icing conditions was obtained.\n\nINTRODUCTION\n\nAs part of a general program to provide icing protection for turbojet engines, a nacelle with several air inlets is being experimentally investigated at the NACA Lewis laboratory to establish a reasonable design criterion for hot-gas bleedback systems.\n\nThe investigation described herein is a continuation of the general program outlined in reference 1 and was conducted with the same turbojet-engine nacelle, but with a long straight air inlet. The nacelle was two-thirds full scale. The model was provided with orifices for introducing hot gas into the inlet. Data were obtained to determine the effect of gas temperature and pressure, tunnel velocity, and angle of attack on the temperature distribution at the\n\nRESTRICTED", "timestamp": "2026-07-22T05:19:56.280014+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 35, "total_pages": 46, "image_filename": "19930085519_p35.jpg", "text": "34\nNACA RM No. L8K19\n\nPlain wing\nFull-span slotted flap, $\\delta_f=30^\\circ$\n1/2-span slotted flap, $\\delta_f=50^\\circ$\nFull-span slotted flap, $\\delta_f=50^\\circ$\n1/2-span Zap flap, $\\delta_f=60^\\circ$\n1/2-span split flap, $\\delta_f=60^\\circ$\n(Unpublished data)\n\n<!-- Image (99, 110, 750, 866) -->\n\nFigure 14.- Gliding characteristics of the $42^\\circ$ sweptback wing in trimmed flight. Wing loading, 40 pounds per square foot; tail length, 3.08.", "timestamp": "2026-07-22T05:20:05.949130+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 21, "total_pages": 24, "image_filename": "19930085626_p21.jpg", "text": "20\nNACA RM No. L8K23\n\nCONFIDENTIAL\nTip obtained by revolving\nairfoil section around\nchord line.\n\nAileron\nc/4 line\n40°\n1.69\n8.48\n1.69\n4.05\n.93\n7.00\n11.50\nA\nA\n\nParallel-side aileron.\nThickness equal to\nthickness of basic\nsection at 0.8c point.\n\nFlat-sided aileron.\nTrailing edge thickness\none-half of that of\nparallel-side aileron.\n\nSection A-A\n(Normal to c/4 line)\n\n$C_D$\n.08\n.04\n0\n\n$pb/2v$\n.08\n.04\n0\n-.04\n.6\n.8\n1.0\n1.2\n1.4\n1.6\n1.8\n2.0\nM\nNACA\n\n(e) Blunt trailing-edge aileron configurations. $\\delta_a = 5^\\circ$.\nFigure 4.- Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:20:06.709190+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 23, "total_pages": 33, "image_filename": "19930085544_p23.jpg", "text": "22\nNACA RM No. L8K26\n\n<!-- Image (192, 138, 752, 874) -->\n\nFigure 5.- Variation of instantaneous thrust coefficient with blade position. B,2; B,26°; J,1.2; $\\omega n_1 4^\\circ$.", "timestamp": "2026-07-22T05:20:12.743666+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 11, "total_pages": 37, "image_filename": "19930085889_p11.jpg", "text": "10 CONFIDENTIAL NACA RM L9F14\n\nand the value for the location of the aerodynamic center as presented in reference 8. The effect of increasing Reynolds number was to decrease the lift-curve slope and to cause a forward shift of the aerodynamic center. The effect on the lift-curve slope and on the aerodynamic-center location of increasing the Reynolds number from 280,000 to 1,116,000 was approximately equivalent to fixing the transition at the nose of the airfoil. The fact that the characteristics of the wing were almost the same with transition strips, either on or off, at a Reynolds number of 1,116,000, is an indication that further increases in Reynolds number would not be particularly important, at least for the present test condition of surface smoothness and air-stream turbulence. Under conditions of extremely low turbulence and with highly polished wing surfaces, the results obtained with transition strips off probably would not approach those with strips on until a Reynolds number considerably higher than 1,116,000 had been attained. (See reference 15.)\n\nThe effect of sweepback on the static lateral stability characteristics is shown in figure 11. At low lift coefficients the rate of change of $C_{l_\\psi}$ with lift coefficient decreases as the sweep is decreased. For the $32.6^\\circ$ and $46.7^\\circ$ sweptback wings the values of $C_{l_\\psi}$ increase linearly for only a small range of lift coefficients after which there is an abrupt change in the initial trends, probably as a result of early partial stalling, mentioned previously. The $32.6^\\circ$ and $46.7^\\circ$ sweptback wings attain relatively small positive values of $C_{l_\\psi}$ (less than the values obtained for the unswept wing at lift coefficients greater than 0.6). There is little effect of sweepback on the values of $C_{n_\\psi}$ and $C_{Y_\\psi}$. The fuselage causes large positive contributions to both $C_{n_\\psi}$ and $C_{Y_\\psi}$. This contribution is shown in figure 12, which compares the values obtained for the $46.7^\\circ$ sweptback-wing and fuselage combination with those for the faired wing alone. Removing the fuselage causes, for the $46.7^\\circ$ sweptback wing, a small change in the variation of $C_{l_\\psi}$ with lift coefficient for low coefficients but has no effect on the maximum positive value of $C_{l_\\psi}$ attained with the combination.\n\nRolling-Flow Characteristics\n\nThe variations of the rolling derivatives $C_{Y_p}$, $C_{n_p}$, and $C_{l_p}$ with lift coefficient are presented in figure 13. As was explained in the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:20:18.609953+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 14, "total_pages": 96, "image_filename": "19930085880_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:20:19.395002+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 42, "total_pages": 66, "image_filename": "19930082914_p42.jpg", "text": "NACA TN No. 1857\n41\n\n[Figure: A black and white photograph of a piece of scientific equipment, an interferometer mirror mount. The device has a large, circular base and a rectangular frame holding a mirror. A small label on the equipment reads \"NACA L-55230\".]\n\nFigure 7.- Interferometer mirror mount.", "timestamp": "2026-07-22T05:20:21.644081+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 37, "total_pages": 78, "image_filename": "19930082618_p37.jpg", "text": "```markdown\nNACA TN 1945\n\nSection lift coefficient, $c_l$\nMoment coefficient, $c_{m_{0.25c}}$\nSection angle of attack, $\\alpha_0$, deg\n\nR\n$\\circ$ $0.7 \\times 10^6$\n$\\diamond$ $1.0$\n$\\square$ $2.0$\n$\\triangle$ $4.0$\n$\\nabla$ $6.0$\nFlagged symbols denote\nstandard roughness\n\n[Figure: NACA logo]\n\n(b) Section lift and pitching-moment characteristics of the NACA $64_1$-012 airfoil section\nwith a 0.20c simulated split flap deflected $60^\\circ$.\n\nFigure 5.— Continued.\n\n35\n```", "timestamp": "2026-07-22T05:20:24.986366+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 2, "total_pages": 42, "image_filename": "19930085914_p2.jpg", "text": "C\nNACA RM A9D25\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING A WING\nSWEPT BACK 63° — CHARACTERISTICS THROUGHOUT THE SUBSONIC\nSPEED RANGE WITH THE WING CAMBERED AND TWISTED FOR A\nUNIFORM LOAD AT A LIFT COEFFICIENT OF 0.25\n\nBy J. Lloyd Jones and Fred A. Demele\n\nSUMMARY\n\nWind-tunnel tests have been made to determine the independent\neffects of Mach and Reynolds numbers on the aerodynamic characteristics\nof a wing-fuselage combination employing a wing having the leading edge\nswept back 63° and having camber and twist. Tests were also made of\nthe fuselage alone.\n\nIncreasing the Mach number from 0.20 to 0.93 resulted in an\nincrease of lift-curve slope from about 0.049 to 0.055 per degree.\nThe abrupt forward movement of the aerodynamic center at the higher\nlift coefficients, typical of highly swept wings, decreased in\nseverity with increasing Mach number.\n\nThe principal effects of increasing Reynolds number from 0.8\nmillion to 9.0 million at a Mach number of 0.20 were a reduction of the\ndrag at positive lift coefficients above about 0.2 and elimination of\nminor irregularities in longitudinal stability up to a lift coefficient\nof about 0.55. These data indicate that certain important effects of\nboundary-layer separation which are evident from tests of highly swept-\nback wings at low Reynolds numbers may not be present under full-scale\nconditions.\n\nCharacteristics of the wing alone were calculated by subtracting\nthe forces and moments of the fuselage alone from those of the wing-\nfuselage combination, and no account was made either of wing-fuselage\ninterference or of the wing area enclosed by the fuselage. The charac-\nteristics thus obtained are compared with those of a wing of identical\nplan form but having no camber or twist. The effects of camber and\ntwist were a reduction of the drag at lift coefficients above about 0.1\nand an increase of about 33 percent in the lift coefficient at which\nloss of static longitudinal stability occurred.", "timestamp": "2026-07-22T05:20:28.044266+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 61, "total_pages": 62, "image_filename": "19930082485_p61.jpg", "text": "NACA TN No. 1810\n\n3.0\n\n$\\sqrt{z_1}/\\sqrt{z_m}$\n\n2.5\n\n2.0\n\n1.5\n\n1.0\n\n0 .2 .4 .6 .8 1.0\n\n$C_1 n_0 / 2$\n\n$\\frac{C_1}{\\Delta C}$\n\n-4.0\n\n-8.0\n\n-16.0\n\n$\\infty$\n\n16.0\n\n8.0\n\n4.0\n\n3.0\n\n2.0\n\n1.5\n\n1.0\n\nFigure 17. - Variation with channel width and blade curvature of ratio of velocities along blade suction surfaces to velocity at channel center.", "timestamp": "2026-07-22T05:20:31.630135+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 4, "total_pages": 23, "image_filename": "19930085906_p4.jpg", "text": "NACA RM E5F20 CONFIDENTIAL 3\n\nconvolutions of 16-inch diameter and $2\\frac{1}{2}$ convolutions of 9-inch diameter (fig. 4(a)) with adjacent turns spaced $1\\frac{1}{2}$ inches apart. Provision was made for varying the distance between the flame holder and the preheater (fig. 1) from 0 to 12 inches. In all positions, the preheater coils were located in the wake of the flame holder (fig. 4(b)). All the fuel passed through the preheater before injection. The inlet and outlet lines to the preheater extended upstream of the flame holder, then through the wall of the diffuser to the outside of the shell.\n\nThe diffuser inlet was connected to the outlet of a 500-horsepower, variable-speed, axial-flow blower having a rated delivery of 60,000 cubic feet per minute against a static-pressure rise of 45 inches of water. The ram jet exhausted directly to the atmosphere.\n\nPROCEDURE\n\nThe total and static pressures and the indicated temperature measured at the diffuser inlet were used to compute the air flow through the ram jet. From these measurements and the static pressure at the combustion-chamber inlet, the combustion-chamber-inlet velocity was determined. The gas temperature was computed from total and static pressures measured at the nozzle outlet with a water-cooled survey rake. The heat rejected to the shell cooling water amounted to approximately 3 percent of the lower heating value of the fuel and was not included in the evaluation of the combustion efficiency. A rotameter was used to measure the fuel flow. The fuel temperatures were measured at the preheater inlet and outlet. The methods employed in making the calculations are outlined in reference 3.\n\nFuel temperature at the preheater outlet was recorded at 0.2-minute intervals from the time of ignition until the fuel temperature reached a stable maximum value. The fuel flow and the air flow were maintained approximately constant during this entire period.\n\nThe effectiveness of the coils as a preheater was investigated at 0, 4, 8, and 12 inches downstream of the flame-holder trailing edge. At each preheater position the ram jet was operated over a range of fuel-air ratios and at combustion-chamber-inlet velocities from 62 to 95 feet per second.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:20:37.069656+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 17, "total_pages": 29, "image_filename": "19930085879_p17.jpg", "text": "NACA RM L9D11\n\n[Figure: Drag flaps retracted.]\n\nNACA\nL-57093\n\nFigure 5.- RM-11 jettisonable nose. Flaps retracted.\n\n15", "timestamp": "2026-07-22T05:20:41.710881+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 36, "total_pages": 46, "image_filename": "19930085519_p36.jpg", "text": "NACA RM No. L8K19\n35\n\nPitching-moment\ncoefficient, $C_m$\n0\n-.1\n-.2\n\nFlow-control\nvanes\n$\\diamond$ off\n$\\square$ on\n\nDrag coefficient, $C_D$\n6\n5\n4\n3\n2\n1\n0\n\nAngle of attack, $\\alpha$, deg\n24\n16\n8\n0\n-8\n-16\n-24\n\nLift coefficient, $C_L$\n-6 -4 -2 0 .2 .4 .6 .8 1.0 1.2 1.4\n\n[NACA logo]\n\nFigure 15.- The effect of flap-slot flow-control vanes A on the aerodynamic characteristics in pitch of the 42° sweptback wing with the full-span slotted flap deflected 50°. Flap position, 1 percent below lip and 1 percent ahead of lip. Tufts on.", "timestamp": "2026-07-22T05:20:47.884943+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 24, "total_pages": 33, "image_filename": "19930085544_p24.jpg", "text": "NACA RM No. L8K26\n23\n\n<!-- Image (164, 110, 882, 904) -->\n\nFigure 6.- Variation of instantaneous thrust coefficient with blade position. B.2; $\\beta$, 53°; J, 2.8; $\\alpha_T$, 40°.", "timestamp": "2026-07-22T05:20:50.229297+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 4, "total_pages": 36, "image_filename": "19930085912_p4.jpg", "text": "2\nNACA RM No. E9C16\n\nsimulated engine inlet. In addition, data were obtained to validate the\nuse of the jet-penetration equation (reference 2) as applied to a three-\ndimensional duct. The icing investigation to determine the minimum heat\nrequirements was conducted over a range of liquid-water contents from\n0.3 to 1.0 gram per cubic meter and at a free-stream total temperature\nof $0^\\circ$ F with the model at angles of attack of $0^\\circ$ and $6^\\circ$.\n\nA design procedure for obtaining a satisfactory orifice configu-\nration is presented in the appendix.\n\nAPPARATUS\n\nThe nacelle investigated was similar to the offset-air-inlet\nnacelle described in reference 1, but had a long, circular, and straight\nair inlet. The model as installed in the tunnel test section is shown\nin figure 1. The model was two-thirds full scale and was constructed\nof steel, Inconel, and aluminum. The inlet length from the nacelle lip\nto the accessory housing was 53 inches. The inlet area at stagnation\nwas 1.227 square feet and the area at the inlet minimum section was\n0.835 square foot. The orifices through which the hot gas was dis-\ncharged were located at the minimum section 5.75 inches from the\nnacelle lip. The hot gas was obtained by passing high-pressure air\nthrough a combustion heater and ducting the air to the model. The\nmodel was designed for a maximum air flow of 32 pounds per second,\ncorresponding to the flow through a full-scale axial-flow engine of\n4000 pounds static thrust at sea level with an 11-stage compressor,\neight cylindrical burners, and a single-stage turbine. A 1/4-inch\nmesh, 0.050-inch-diameter wire screen was mounted in the model\n(fig. 2) to simulate a protective screen installation and to provide\na means of indicating icing.\n\nINSTRUMENTATION\n\nThe model instrumentation used in the investigation is shown in\nfigure 2. Measurements were made of mass flow, ram-pressure recovery,\ntemperature distribution at the simulated engine inlet, pressure drop\nacross the screen, and air temperature ahead of and behind the screen.\nThe model aft of the temperature cross rake (fig. 2) is the same as that\nused in the offset-inlet investigation (reference 1). A detailed\ndescription of the instrumentation is presented in reference 1.\n\nThe inlet duct of the model was instrumented with three thermo-\ncouple rakes located 13.2, 23.1, and 33.1 inches downstream of a plane", "timestamp": "2026-07-22T05:20:50.844261+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 15, "total_pages": 96, "image_filename": "19930085880_p15.jpg", "text": "NACA RM No. L9C03\n13\n\n[Figure: Typical static waterline.]\n\n[Figure: Typical dynamic waterline. (Image includes an arrow labeled \"Wetted length\")]\n\nNACA\n\nFigure 5. - Underwater photographs of model 250B at trim of $4^{\\circ}$.", "timestamp": "2026-07-22T05:20:54.379303+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 19, "total_pages": 36, "image_filename": "19930085869_p19.jpg", "text": "NACA RM L9D15\n17\n\nCONFIDENTIAL\n\n<!-- Image (106, 176, 846, 393) -->\n\n(a) Spray strip similar to that on unswept model.\n\n<!-- Image (112, 531, 886, 775) -->\n\n(b) Final spray strip used throughout investigation.\n\nFigure 4.- Spray-strip arrangements. (All dimensions are in inches.)", "timestamp": "2026-07-22T05:20:58.389910+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 34, "total_pages": 58, "image_filename": "19930082617_p34.jpg", "text": "NACA TN 1962\n33\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 $10^3$\n\n2.57\"\n[Figure: Cross-section diagram labeled Band L]\n[Figure: Circular cross-section diagram labeled A-A with 45° angles]\n\nDistance from horizontal diameter, in.\nStrain\n\n16 12 8 4 0 -4 -8 -12 -16 -20 X $10^{-4}$\n\n1 2 3 4 5 6\n\n[Figure: NACA logo]\n\nFigure 22.- Strain diagram of cylinder 78. Band L.", "timestamp": "2026-07-22T05:20:59.678290+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 31, "total_pages": 53, "image_filename": "19930082542_p31.jpg", "text": "NACA TN No. 1867\n31\n\n[Figure: Two rectangular images showing metal cross-sections]\n\nCross section $1\\frac{1}{2}$X Longitudinal section near center of bar\n(a) Macrostructure of low-carbon N-155 hot-rolled bar stock.\nEtchant: 2 hours in Marble's Reagent at $160^\\circ$ F plus\n15 minutes in aqua regia in glycerine at $120^\\circ$ F.\n\n[Figure: Two rectangular images showing metal microstructures]\n\n1000X 100X [NACA logo]\n(b) Microstructure of low-carbon N-155 hot-rolled bar stock.\n(Electrolytic chromic acid etch.)\nFigure 1.- Structure of original bar stock.", "timestamp": "2026-07-22T05:21:00.643125+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 12, "total_pages": 37, "image_filename": "19930085889_p12.jpg", "text": "NACA RM L9F14 CONFIDENTIAL 11\n\nsection entitled \"Apparatus and Tests,\" measurements of forces and moments were obtained at four values of pb/2V. The derivatives were obtained from the average slopes of the data when plotted against pb/2V. In general, the slopes of the curves were well defined and the scatter of tests points was of the order of that obtained in other investigations which have utilized the rolling-flow technique. (See reference 7, for example.)\n\nAt low lift coefficients the results presented in figure 13 for the derivatives of lateral force caused by rolling $C_{Y_p}$ are in qualitative agreement with the approximate theory of reference 11 in that this derivative varied linearly with lift coefficient and the rate of variation increased with an increase in sweep angle. In general, $C_{Y_p}$ maintained its initial linear variation over about the same range of lift coefficients as the derivative $C_{l_\\psi}$.\n\nThe derivative of yawing moment caused by rolling $C_{n_p}$ was found to be either zero or positive from a lift coefficient of -0.2 to approximately the maximum positive lift coefficient for each of the wings tested. The approximate theory of reference 11, which is based on potential-flow considerations, indicates an initial negative slope of $C_{n_p}$ with $C_L$; however, this initial trend would be expected to be maintained only over the range of lift coefficients for which the total drag is approximately equal to the drag at zero lift plus the induced drag. (See reference 2.) As is indicated by the drag data of figure 6, this condition is satisfied only up to lift coefficients of about 0.2 or 0.3. At such low lift coefficients the magnitudes of the theoretical values of the yawing moment due to rolling probably are within the experimental accuracy of the measurements and, therefore, no initial negative slope could be detected.\n\nThe experimental results for the derivative $C_{n_p}$ are compared in figure 14 with results calculated by a method (presented in reference 2) which includes consideration of the drag measured under straight-flow conditions. In general, fair agreement is obtained, although the predicted values of $C_{n_p}$ at high lift coefficients are too highly positive for the 3.6° and 32.6° sweptback wings. In reference 2, through analysis of experimental data, the increment of $C_{n_p}$ due to profile drag was found to be proportional to the slope of the curve of profile drag plotted against angle of attack, and the constant of proportionality was found to vary with aspect ratio but to be essentially independent of the sweep angle. The comparison presented\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:21:02.448975+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 38, "total_pages": 78, "image_filename": "19930082618_p38.jpg", "text": "```markdown\n36\n\nSection drag coefficient, $c_d$\n.028\n.024\n.020\n.016\n.012\n.008\n.004\n0\n-.8\n-.4\n0\n.4\n.8\n1.2\nSection lift coefficient, $c_l$\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\triangle$ 1.5\n$\\diamond$ 2.0\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-1.2\n-.8\n-.4\n0\n.4\n.8\n1.2\n1.6\nSection lift coefficient, $c_l$\n\nR\n$\\nabla$ 3.0 x $10^6$\n$\\triangledown$ 6.0\n$\\blacktriangledown$ 9.0\nFlagged symbols denote\nstandard roughness\n\nMoment coefficient, $c_m$\n.04\n.02\n0\n-.02\n-.04\n-.8\n-.4\n0\n.4\n.8\n1.2\n1.6\nSection lift coefficient, $c_l$\n\nR\na.c. position\nx/c\ny/c\n$\\circ$ 0.7 x $10^6$\n.253\n.059\n$\\square$ 1.0\n.257\n.051\n$\\triangle$ 1.5\n.261\n.020\n$\\diamond$ 2.0\n.265\n.005\n$\\nabla$ 3.0\n.278\n.028\n$\\triangledown$ 6.0\n.289\n.017\n$\\blacktriangledown$ 9.0\n.282\n-.002\n\nNACA\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic center of the plain NACA 64$_1$-012 airfoil section.\n\nFigure 5.— Concluded.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:21:08.537860+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 54, "total_pages": 99, "image_filename": "19930082511_p54.jpg", "text": "52\nNACA TN No. 1826\n\nwhere $K_0$ and $K_1$ are the Bessel functions as defined in reference 15.\n\nThe points $\\{\\xi_i\\}$ were taken as $a, a + 0.3, a + 0.6, \\dots, a + 2.7$, that is, a set of ten points, at 0.3 intervals, starting with the entrance lip of the tunnel. The coefficients $h_{1n}^{(1)}$ were found by the method of least squares for $N = 0, 1, \\dots, 5$ and also so as to satisfy the equations $r_1^{(1)}(\\xi_i) = \\sum_{n=0}^{N} h_{1n}^{(1)} \\frac{\\partial P_{1n}(\\xi, 1)}{\\partial \\xi}$ at all ten points. Plots of the resulting functions $g_1^{(1)}(\\xi)$ for the different values of $N$ indicated that convergence was essentially complete for $N$ between 3 and 5. This simplification results in appreciable saving in the amount of computation. Not only is it necessary to solve a smaller set of simultaneous equations, but also $P_{1n}$ and $\\partial P_{1n}/\\partial \\xi$ need be found for fewer values of $n$.\n\nThe computation was fairly straightforward. In the determination of $r_1^{(1)}(\\xi)$, $K_0$ and $K_1$ were obtained from the tables of reference 15, and $J_1$ and $J_1'$ from the tables and from the relations between the Bessel functions and their derivatives (references 15 and 16). Weddle's formula (reference 17) is convenient for performing the integrations. In the case of $P_{1n}$, the values of $y_{s1}$ appearing in the formula for $Q_{1n}^{(s)}(\\rho)$ were found from the formula in appendix III of reference 15, and $J_1$ and $J_1'$ as just noted. In the evaluation of $\\lim_{\\rho \\to 0} \\frac{1}{\\rho} P_{1n}$ it is noted that the value of $\\lim_{x \\to 0} \\frac{1}{x} J_1(x) = \\frac{1}{2}$.\n\nThe results of these computations, together with those for the completely open and completely closed tunnels and those given in reference 3 are shown in figures 24, 25, and 26. In figure 24, the vertical tunnel-induced velocity along the axis for the four different positions of the lifting element together with the results for the open and closed tunnels are plotted against distance from the lifting element. The same results are plotted against the longitudinal distance from the entrance lip in figure 25. Figure 26 shows the results of reference 3 compared with the results of this paper for the same case - that of the lifting element 1 radius downstream from the entrance lip.\n\nThe tunnel-induced velocity in the upstream regions and in the neighborhood of the lifting element, although only slightly less than that for an open tunnel for the lifting element 1 radius downstream, falls off more and more rapidly as the lifting element is moved towards the entrance lip. The maximum induced velocity is attained about 1 radius upstream from the exit, and is never more than 78 percent of the maximum value for a completely open tunnel. After the maximum the values fall rapidly and approach the values for a closed tunnel in the downstream regions. The results of reference 3 (see fig. 26) are consistently below the present results especially in the region behind the lifting element.", "timestamp": "2026-07-22T05:21:10.363281+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 65, "total_pages": 66, "image_filename": "19930082245_p65.jpg", "text": "64\nNACA TN No. 1596\n\nPressure coefficient, P\n-6\n-4\n-2\n0\n.2\n.4\n.6\nM=0.25\n\n-6\n-4\n-2\n0\n.2\n.4\n.6\nM=0.35\n\nUpper surface Lower surface\nTrue-contour aileron\nBeveled-trailing-edge aileron\n\n-6\n-4\n-2\n0\n.2\n.4\n.6\nM=0.55\n\n-6\n-4\n-2\n0\n.2\n.4\n.6\nM=0.65\n\n-6\n-4\n-2\n0\n.2\n.4\n.6\nM=0.70\n\n-6\n-4\n-2\n0\n.2\n.4\n.6\nM=0.75\nM=0.74\nNACA\n\n-1 0 1 2 3 4 5 6\nAileron chord, in.\n\n-1 0 1 2 3 4 5 6\nAileron chord, in.\n\n(b) $\\delta_a = 4^\\circ$.\nFigure 16.—Concluded.", "timestamp": "2026-07-22T05:21:10.625676+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 62, "total_pages": 62, "image_filename": "19930082485_p62.jpg", "text": "NACA TN No. 1810\n\n1.0\n\n.9\n\n.8\n\n.7\n\n$\\sqrt{z_2}/\\sqrt{z_m}$ .6\n\n.5\n\n.4\n\n.3\n\n.2\n\n0 .2 .4 .6 .8 1.0\n\n$C_1 r_0 / 2$\n\n$-\\frac{C_1}{\\Delta C}$\n\n1.0\n\n1.5\n\n2.0\n\n3.0\n\n4.0\n\n8.0\n\n16.0\n\n$\\infty$\n\n-16.0\n\n-8.0\n\n-4.0\n\nFigure 18. - Variation with channel width and blade curvature of ratio of velocities along blade pressure surfaces to velocity at channel center.", "timestamp": "2026-07-22T05:21:12.147326+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 3, "total_pages": 42, "image_filename": "19930085914_p3.jpg", "text": "2\nNACA RM A9D25\n\nINTRODUCTION\n\nThe advantages of wings having large amounts of sweepback for efficient flight at supersonic speeds up to Mach numbers of approximately 1.5 have been pointed out by R. T. Jones in reference 1. A coordinated program was formulated for investigation in various facilities of the Ames Aeronautical Laboratory of a wing-fuselage combination designed according to the indications of that study.\n\nTests to date have shown that the rate of drag increase with lift coefficient was greater than theory predicted, the discrepancy being attributed to boundary-layer separation resulting from an adverse chord-wise pressure gradient due to lift, especially severe at the wing tips where the induced upwash is large. Camber and twist have been suggested (reference 1) as possible means of decreasing this adverse pressure gradient. A discussion of the design of a wing incorporating such camber and twist is presented in reference 2, along with the results of tests of this wing at a Mach number of 1.53.\n\nThe present report presents the results of tests in the Ames 12-foot pressure wind tunnel of a sting-mounted model of a cambered and twisted wing having the leading edge swept back $63^\\circ$ in combination with a slender fuselage. The model was similar to the model used for the tests reported in reference 2. The effects of the independent variation of Mach and Reynolds numbers on the subsonic characteristics of the wing-fuselage combination and of the fuselage alone are presented. A comparison is made with a wing of identical plan form, but having no camber or twist. Data were obtained at the lowest Reynolds number (0.8 million) to aid in evaluating other data on highly swept wings obtained at comparable Reynolds numbers.\n\nSYMBOLS\n\nThe following coefficients and symbols are used in this report:\n\na speed of sound, feet per second\n\nb wing span measured perpendicular to plane of symmetry, feet\n\nc local chord measured parallel to plane of symmetry, feet\n\n$\\overline{c}$ wing mean aerodynamic chord $\\left( \\frac{\\int_{0}^{b/2} c^2 dy}{\\int_{0}^{b/2} c \\ dy} \\right)$, feet\n\n$C_D$ drag coefficient $\\left( \\frac{\\text{drag}}{qS} \\right)$", "timestamp": "2026-07-22T05:21:12.269139+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 18, "total_pages": 29, "image_filename": "19930085879_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:21:14.406443+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 22, "total_pages": 24, "image_filename": "19930085626_p22.jpg", "text": "$8 \\times 10^6$\n\nCONFIDENTIAL\n\nNACA RM No. L8E23\n\nReynolds number, R\n\nMach number, M\n\nNACA\n\nFigure 5.- Variation of Reynolds number with Mach number for range of test conditions.\n\nCONFIDENTIAL\n\n21", "timestamp": "2026-07-22T05:21:18.225247+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 37, "total_pages": 46, "image_filename": "19930085519_p37.jpg", "text": "36\nNACA RM No. 18K19\n\nPitching-moment coefficient, $C_m$\nDrag coefficient, $C_D$\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\n\nFlow control vanes\n$\\Delta$ off\n$\\square$ on\n\n<!-- Image (129, 80, 796, 791) -->\n\nFigure 16. The effect of flap-slot flow-control vanes B on the aero-dynamic characteristics in pitch of the 42° sweptback wing with the full-span slotted flap deflected 50°. Flap position, 1 percent below lip.", "timestamp": "2026-07-22T05:21:22.112435+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 5, "total_pages": 23, "image_filename": "19930085906_p5.jpg", "text": "4 CONFIDENTIAL NACA RM E5F20\n\nRESULTS AND DISCUSSION\n\nThe data presented herein are characteristic of the installation and should be considered only as an indication of the potentialities of this type of fuel-preheating system. Because a fuel injector with fixed orifices was used, the fuel flow was changed by varying the fuel pressure. Thus the minimum fuel flow at which the ram jet could be operated was determined by the minimum fuel pressure required to keep the preheated fuel from vaporizing in the fuel lines. Additional restriction to the range over which data could be obtained was imposed by the flame-holder burning characteristics.\n\nThe range of fuel-air ratios over which the engine could be operated at each preheater position is shown in figure 5. In figure 5 and several succeeding figures, data are presented in range form rather than as actual data points because it was impossible from the data available to separate the effects of combustion-chamber-inlet velocity and fuel-air ratio on the final fuel temperature. The rich fuel-air-ratio limit was established by flame-holder blow-out, whereas the lean limit was due to either flame-holder blow-out or fuel vapor lock. The lean fuel-air-ratio limit can probably be lowered if fuel vapor lock is eliminated. The introduction of the preheater in the combustion chamber and the preheater position had no great effect on the operable fuel-air-ratio range, inasmuch as this range was approximately the same as was obtained with the fuel preheated by an external preheater (reference 4).\n\nFor aircraft of short flight duration, the time required for the fuel temperature to rise to a steady value may be an appreciable portion of the total flight time. The rate of fuel-temperature increase from the time of ignition to the time at which a stable fuel temperature is finally attained is therefore important. The variation of fuel temperature with time after ignition and preheater position is presented in figure 6 for several typical operating conditions. Fuel was stored at a temperature of approximately $63^\\circ$ F.\n\nAll the data shown in figure 6 can be reduced to a single curve (fig. 7) by plotting the ratio of the temperature rise at any instant to the maximum temperature rise. All the data obtained in this investigation, irrespective of ram-jet operating condition or preheater position, are within $\\pm 6$ percent of this curve. The variation indicated by the curve in figure 7 may be approximated by a first-order exponential equation of general form $y = 1 - e^{-at}$ having a time constant $t = 1/a$ of 1.17 minutes. An inflection of\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:21:23.736724+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 43, "total_pages": 66, "image_filename": "19930082914_p43.jpg", "text": "42\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:21:24.714873+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 25, "total_pages": 33, "image_filename": "19930085544_p25.jpg", "text": "24\nNACA RM No. L5K26\n\n<!-- Image (110, 131, 796, 903) -->\n\nFigure 7.- Variation of instantaneous thrust coefficient with blade position. B,2; $\\beta$,53°; J,3.1; $\\alpha_{T_1}$,40°.", "timestamp": "2026-07-22T05:21:26.298741+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 16, "total_pages": 96, "image_filename": "19930085880_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:21:32.625802+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 5, "total_pages": 36, "image_filename": "19930085912_p5.jpg", "text": "NACA RM No. E9C16\n\nthrough the orifice center lines. In each case the probes were spaced 1/2 inch apart with the outer probe coinciding with the model center line. The rakes were located directly behind a 3/4-inch-diameter orifice. The average air temperature and temperature distribution inside the model were measured by means of a thermocouple cross rake mounted in the duct 43 inches downstream of the orifices and just ahead of the simulated accessory housing, as shown in figure 2. The rake consisted of 29 total-temperature thermocouples spaced 1 inch apart and mounted in two streamlined struts intersecting at $90^\\circ$. Temperatures on the surface of the inlet duct were measured by 47 flush-type thermocouples, 30 of which were mounted in the duct wall and 17 on the accessory housing.\n\nThe state of the gas in the plenum chamber was measured by four thermocouple probes located $90^\\circ$ apart in the plane of the orifices and by four static-pressure taps in the rear wall of the plenum chamber.\n\nThe nacelle-lip pressure distribution was measured by means of pressure belting cemented to the lip surface. Lip-temperature distribution was measured by thermocouples welded to the nacelle lip.\n\nAir flow through the model and inlet-velocity ratio were controlled by an electrically driven tail cone (fig. 2).\n\nSYMBOLS\n\nThe following symbols are used in this report:\n\n| Symbol | Description |\n|--------|-------------|\n| $A_1$ | inlet area at orifices, square feet |\n| $A_j$ | total area of orifices (jet area), square feet |\n| $A_s$ | free area through screen, 0.732 square foot |\n| $c_p$ | specific heat of air, Btu per pound $^\\circ$F |\n| $c_{p,g}$ | specific heat of gas, Btu per pound $^\\circ$F |\n| $c_{p,w}$ | specific heat of water, Btu per pound $^\\circ$F |\n| $D_j$ | diameter of orifice (jet diameter), inches |\n| $g$ | acceleration due to gravity, 32.2 feet per second per second |\n| $J$ | mechanical equivalent of heat, 778 foot-pounds per Btu |", "timestamp": "2026-07-22T05:21:32.806221+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 20, "total_pages": 36, "image_filename": "19930085869_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:21:35.562837+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 32, "total_pages": 53, "image_filename": "19930082542_p32.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:21:41.071492+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 13, "total_pages": 37, "image_filename": "19930085889_p13.jpg", "text": "12 CONFIDENTIAL NACA RM L9F14\n\nin figure 14 indicates that the constant of proportionality probably should be somewhat lower than that given in reference 2 for wings having sweep angles less than about $45^\\circ$.\n\nAt low lift coefficients the negative value of $C_{l_p}$ (fig. 13) of the wing-fuselage combination decreases as the sweep is increased. At some lift coefficient which decreases with an increase in sweep, there is a sudden increase in $C_{l_p}$. The magnitude of the increase is greatest for the wing with the largest sweep. At the higher lift coefficients the damping decreases for all three wings.\n\nA comparison of values of $C_{l_p}$ obtained by the rolling-flow technique of the Langley stability tunnel with those obtained from the free-rotation tests of the models in the Langley 7- by 10-foot tunnel (reference 6) is presented in figure 15. In general, the variation of $C_{l_p}$ with lift coefficient is similar, and the values of $C_{l_p}$ are in good agreement. The Langley 7- by 10-foot tunnel results are slightly higher, but this difference can be attributed almost entirely to the difference in Mach number of the tests, as is indicated in figure 16, which compares experimental results obtained by the two techniques with theoretical results (from reference 16) corresponding to the two test Mach numbers. The difference between the two theoretical curves is almost exactly equal to the difference between the two experimental curves. Both experimental techniques yield values that are consistently larger than the theoretical values, although the experimental variation of $C_{l_p}$ with sweep angle is in good agreement with theory.\n\nThe $46.7^\\circ$ sweptback wing was also tested in rolling flow with the fuselage removed. The effect on $C_{Y_p}$, $C_{n_p}$, and $C_{l_p}$ of removing the fuselage was small. (See fig. 17.) The values of $C_{l_p}$ obtained with the wing alone were slightly larger than those obtained for the combination, although the slope of the lift curve for wing alone was lower than that obtained with the combination. A similar result was obtained in the tests reported in reference 17 which gives as a possible explanation the fact that the loading on the fuselage during roll would act normal to the surface of the circular cross-section fuselage and would contribute little to the damping in roll.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:21:42.657065+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 11, "total_pages": 31, "image_filename": "19930085881_p11.jpg", "text": "NACA RM L9D12 CONFIDENTIAL 9\n\nTABLE II\n\nPHYSICAL CHARACTERISTICS OF CONFIGURATION TESTED\n\n| Configuration | Wing sweepback (deg) | $^{\\text{a}}$NACA airfoil section | $^{\\text{d}}$Nominal trailing-edge angle (deg) |\n| :--- | :--- | :--- | :--- |\n| 50 | 0 | 65A009 | 10.6 |\n| 115 | 0 | 16-009 | 21.0 |\n| 119 | 0 | (b) | 10.2 |\n| 117 | 0 | (c) | 20.4 |\n| 53 | 45 | 65A009 | 10.6 |\n| 116 | 45 | 16-009 | 21.0 |\n| 120 | 45 | (b) | 10.2 |\n| 118 | 45 | (c) | 20.4 |\n\n[NACA logo]\n\n$^{\\text{a}}$In plane parallel to test vehicle center line.\n$^{\\text{b}}$Double-wedge section of 9-percent-thickness ratio symmetrical with respect to chord line and to line normal to chord line at midchord point.\n$^{\\text{c}}$Circular-arc section of 9-percent-thickness ratio symmetrical with respect to chord line and to line normal to chord line at midchord point.\n$^{\\text{d}}$In plane parallel to test vehicle center line. Actual measured values agreed with nominal values within $\\pm 0.2^{\\circ}$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:21:46.542004+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 35, "total_pages": 58, "image_filename": "19930082617_p35.jpg", "text": "34\nNACA TN 1962\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X 10³\n2 72.0 X 10³\n3 108.0 X 10³\n4 144.0 X 10³\n5 180.0 X 10³\n6 216.0 X 10³\n\n[Figure: Cross-section diagram labeled \"Band V\" with dimensions \"2.57\" and \"A-A\" showing a 45° angle]\n\nDistance from horizontal diameter, in.\nStrain\n\nFigure 23.- Strain diagram of cylinder 78, Band V.", "timestamp": "2026-07-22T05:21:47.882035+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 66, "total_pages": 66, "image_filename": "19930082245_p66.jpg", "text": "NACA TN No. 1596\n65\n\n<!-- Image (117, 110, 959, 790) -->\n\nFigure 17.- Variation of wing drag coefficient with Mach number for an NACA 66,1-115 airfoil section equipped with unsealed 0.20c plain ailerons. $\\delta_a=0^\\circ$. Drag data from force-test measurements.", "timestamp": "2026-07-22T05:21:50.986509+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 39, "total_pages": 78, "image_filename": "19930082618_p39.jpg", "text": "```markdown\nNACA TN 1945\n\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-1.6\n-2.0\nSection lift coefficient, $c_l$\n\n+1.2\n.1\n0\n-.1\n-.2\n-.3\n-.4\nMoment coefficient, $c_{m_{c/4}}$\n\n-16 -8 0 8 16 24 0 0 0 0 0 0\nSection angle of attack, $\\alpha_0$, deg\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.7\n$\\diamond$ 2.9\n$\\triangle$ 4.0\n$\\nabla$ 5.2\nFlagged symbols denote\nstandard roughness\n\n[Figure: Airfoil cross-section diagram]\n\n[NACA logo]\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\n\nFigure 6.— Aerodynamic characteristics of the NACA 64$_1$A212 airfoil section, 24-inch chord.\n\n37\n```", "timestamp": "2026-07-22T05:21:53.316790+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 19, "total_pages": 29, "image_filename": "19930085879_p19.jpg", "text": "NACA RM L9D11\n17\n\nDrag flaps extended\n\n[Figure: A model of a jettisonable nose cone mounted on a wooden stand. The nose cone has extended drag flaps.]\n\nNACA\nL-57092\n\nFigure 6.- RM-11 jettisonable nose. Flaps open.", "timestamp": "2026-07-22T05:21:56.217891+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 1, "total_pages": 23, "image_filename": "19930085934_p1.jpg", "text": "FILE COPY\nNO 3\nRM E9G12\nNACA BM E9G12\n\nNACA\n\nRESEARCH MEMORANDUM\n\nMETHOD OF DETERMINING CENTRIFUGAL-FLOW-COMPRESSOR\nPERFORMANCE WITH WATER INJECTION\n\nBy Joseph T. Hamrick and William L. Beede\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nDOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE A\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nREVIEWED BUT NOT\nEDITED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nSeptember 7, 1949", "timestamp": "2026-07-22T05:21:57.760515+00:00"}