PDF/ocr_dataset_1036.jsonl

{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 41, "total_pages": 46, "image_filename": "19930085519_p41.jpg", "text": "40\nNACA RM No. L8K19\n\n$\\delta p$\n(percent\nlocal wing\nchord)\n$\\nabla$ -1/2\n$\\nabla$ -1\n$\\square$ -2\n$\\diamond$ -3\n$\\circ$ -5\n$\\cup$ -7\n\n<!-- Image (132, 284, 830, 768) -->\n\nFigure 19.- Variation of rolling-moment coefficient and yawing-moment\ncoefficient with angle of attack for various plug-aileron\nprojections on the 42° sweptback wing. Faired plug-slot lower\nlip; full-span slotted flap at $\\delta_F = 30^\\circ$.", "timestamp": "2026-07-22T05:24:04.528123+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 24, "total_pages": 36, "image_filename": "19930085869_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:24:05.135842+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 7, "total_pages": 36, "image_filename": "19930085912_p7.jpg", "text": "NACA RM No. E9C16\n\n$W_a$ mass flow through model, pounds per second  \n$W_d$ design mass flow through model, pounds per second  \n$W_g$ hot-gas flow through orifices, pounds per second  \n$W_s$ saturated vapor content, pounds water per pound dry air  \n$\\beta$ bleedback ($W_g/W_a \\times 100$), percent  \n$\\eta$ ram-pressure recovery, $1 - \\left(\\frac{P_0 - P_f}{q_0}\\right)$  \n$\\eta_{calc}$ calculated ram-pressure recovery  \n$\\rho_a$ density of air in model, slugs per cubic foot  \n$\\rho_d$ design free-stream density, slugs per cubic foot  \n$\\rho_i$ mass density of inlet stream at orifices, slugs per cubic foot  \n$\\rho_j$ mass density of jet at vena contracta, slugs per cubic foot  \n$\\rho_0$ mass density of free stream, slugs per cubic foot  \n\nPROCEDURE\n\nAerodynamic investigation without bleedback. - An aerodynamic investigation of the model without orifices was conducted to determine air-flow characteristics, lip-pressure distribution, and ram-pressure recovery as a function of inlet-velocity ratio and angle of attack. The range of tunnel-air velocities was from approximately 200 to 400 feet per second. At each tunnel velocity investigated, the angle of attack was varied from $0^\\circ$ to $8^\\circ$ and for each angle of attack the inlet-velocity ratio was varied from 0.64 to 0.82.\n\nAerodynamic investigation with cold-gas bleedback. - An aerodynamic investigation of the model with orifices was conducted to determine the effect of cold-gas bleedback on air-flow characteristics, lip-pressure distribution, and ram-pressure recovery. This investigation was conducted at an angle of attack of $0^\\circ$ at a fixed tail-cone position corresponding to an inlet-velocity ratio of 0.82 without bleedback. The tunnel-air velocities ranged from 200 to 460 feet per second and bleedback ranged from 2.4 to 10.4 percent.", "timestamp": "2026-07-22T05:24:05.911713+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 47, "total_pages": 66, "image_filename": "19930082914_p47.jpg", "text": "46\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:24:06.896482+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 20, "total_pages": 96, "image_filename": "19930085880_p20.jpg", "text": "18\nNACA RM No. L9C03\n\n<!-- Image (100, 299, 817, 655) -->\n\nFigure 7.- Variation of wetted lengths at chine and at model centerline\nwith wetted area for model 250D.", "timestamp": "2026-07-22T05:24:13.523036+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 15, "total_pages": 31, "image_filename": "19930085881_p15.jpg", "text": "NACA RM L9D12 CONFIDENTIAL 13\n\n[Figure: A photograph of a white rocket model with a pointed nose cone and four fins at the base, standing vertically against a dark background. A ruler marked in inches is placed at the base for scale. A label on the rocket reads \"510-F\". A NACA identification tag in the lower right corner reads \"NACA I-57478\".]\n\n(a) $\\Lambda = 0^\\circ$.\n\nFigure 2.- Photographs of typical test vehicles.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:24:20.348592+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 23, "total_pages": 29, "image_filename": "19930085879_p23.jpg", "text": "NACA RM L9D11\n21\n\n[Figure: A line graph plotting Acceleration (Gs) against Time from launching, sec. The Y-axis ranges from 0 to -5. The X-axis ranges from 99 to 102. Two lines are plotted: a solid line labeled \"Nose\" and a dashed line labeled \"Rear body\". The NACA logo is present at the bottom right of the graph.]\n\nFigure 9.- RM-11B longitudinal acceleration of nose and rear body during separation.", "timestamp": "2026-07-22T05:24:29.475178+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 36, "total_pages": 53, "image_filename": "19930082542_p36.jpg", "text": "**Page intentionally left blank**\n\nPage 37—52 (every other page is blank)\n\n**Page intentionally left blank**", "timestamp": "2026-07-22T05:24:31.316890+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 4, "total_pages": 23, "image_filename": "19930085934_p4.jpg", "text": "NACA RM E9G12\n\nv specific volume, (cu ft/lb)\n\nW air weight, (lb)\n\nw/a water-air ratio due to water injection\n\n$\\eta_{\\text{ad}}$ adiabatic efficiency\n\n$\\varphi = \\int_{T_0}^{T} \\frac{c_p}{T} \\, dT, \\quad (\\text{Btu}/(\\text{lb})(^\\circ\\text{F}))$\n\nThe following subscripts refer to states of working fluid through the compressor:\n\n0 NACA standard sea-level conditions\n\n1 compressor inlet before injection of water\n\n2 compressor outlet\n\na constituent of mixture based on 1 pound of dry air\n\nd dry air\n\nf saturated liquid\n\ng saturated vapor\n\nm mixture\n\ns superheated vapor\n\nt stagnation\n\nANALYSIS\n\nIn order to enable comparisons of compressor performance with wet- and dry-compression, in the following analysis adiabatic efficiency is based on the isentropic-adiabatic compression process used for dry compression. The adiabatic efficiency of a compressor is defined as the ratio of the isentropic (reversible adiabatic) compression work to the actual compression work. Although the actual work can be determined by either thermodynamic or mechanical methods, the isentropic work must be computed by use of thermodynamics.", "timestamp": "2026-07-22T05:24:31.783793+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 39, "total_pages": 58, "image_filename": "19930082617_p39.jpg", "text": "38\nNACA TN 1962\n\nStrain\n20 X 10^-4\n18\n16\n14\n12\n10\n8\n6\n4\n2\n0\n\nLength of cutout, 51.4 in.\n\nDistance along edge stringer, in.\n10\n20\n30\n40\n50\n60\n70\n80\n\nM=216,000\nM=180,000\nM=144,000\nM=108,000\nM=72,000\nM=36,000\nM=288,000\n\nNACA\n\nFigure 27.- Variation of strain along stringer at edge of cutout. Cylinder 77. M, moment, inch-pounds.", "timestamp": "2026-07-22T05:24:35.603525+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 9, "total_pages": 23, "image_filename": "19930085906_p9.jpg", "text": "8\nCONFIDENTIAL\nNACA RM E9F20\n\nThe results presented are characteristic of the configuration investigated and other configurations may not give similar performance.\n\nLewis Flight Propulsion Laboratory,\nNational Advisory Committee for Aeronautics,\nCleveland, Ohio.\n\nREFERENCES\n\n1. Perchonok, Eugene, Wilcox, Fred A., and Sterbentz, William H.: Investigation of the Performance of a 20-Inch Ram Jet Using Preheated Fuel. NACA RM E6I23, 1946.\n\n2. Anon.: Summary Report on Jet Engine Development. Continental Aviation and Eng. Corp. (Detroit), Aug. 1, 1948. (Air Forces Contract W33-038-ac-13371.) (Available from Central Air Documents Office (Wright-Patterson Air Force Base, Ohio), under ATI No. 52077.)\n\n3. Perchonok, Eugene, Wilcox, Fred A., and Sterbentz, William H.: Preliminary Development and Performance Investigation of a 20-Inch Steady-Flow Ram Jet. NACA ACR E6D05, 1946.\n\n4. Sterbentz, W. H., Perchonok, E., and Wilcox, F. A.: Investigation of Effects of Several Fuel-Injection Locations on Operational Performance of a 20-Inch Ram Jet. NACA RM E7L02, 1948.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:24:35.981587+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 30, "total_pages": 33, "image_filename": "19930085544_p30.jpg", "text": "NACA RM No. L8K26\n29\n\n<!-- Image (185, 124, 879, 866) -->\n\nFigure 12.- Variation of force coefficient with k. $\\epsilon = 0$.", "timestamp": "2026-07-22T05:24:41.897439+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 3, "total_pages": 42, "image_filename": "19930085938_p3.jpg", "text": "2\nNACA RM No. L9B04\n\nvolume and a substantial decrease in the aerodynamic drag below that of\nthe hulls reported in reference 1. The two configurations, which had lower\ndrag than the others, were a configuration with an \"afterbody\" that was\nsimply a tapered boom of circular cross section (see fig. 1(a)) and a\nconfiguration that had two \"afterbodies\" consisting of tapered booms fairing\nout of the engine nacelles (see fig. 1(b)).\n\nThe results of the hydrodynamic investigation of these two configu-\nrations conducted in Langley tank no. 2 are given in the present paper.\nBecause of the large portion of the total volume forward of the center of\ngravity, the problem of airplane balance may limit the application of these\nhulls to special-purpose, high-performance airplanes.\n\nThere was some doubt that a small conical boom would be a hydrody-\nnamically adequate substitute for an afterbody, although tests of refer-\nence 4 had indicated that a small cylindrical boom might be sufficient.\nConsequently, there was included in the wind-tunnel investigation a\nhull in which a small tail float was faired into the end of the tail\nboom (see fig. 1(c)). Exploratory tank tests were made with the tail\nfloat on the single-boom configuration but these tests showed that the\ntail float actually impaired take-off performance and tank tests were\ndiscontinued in favor of the simpler hulls having lower drag.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| $C_{\\Delta_o}$ | gross load coefficient $\\left(\\Delta_o/wb^3\\right)$ |\n| $C_{\\Delta}$ | load coefficient $\\left(\\Delta/wb^3\\right)$ |\n| $C_V$ | speed coefficient $\\left(V/\\sqrt{gb}\\right)$ |\n| $C_R$ | resistance coefficient $\\left(R/wb^3\\right)$ |\n| $\\Delta/R$ | load-resistance ratio |\n| $\\Delta_o$ | gross load on water, pounds |\n| $\\Delta$ | load on water, pounds |\n| R | resistance, pounds |\n| V | speed, feet per second |\n| $\\tau$ | trim, measured between forebody keel and horizontal, degrees |\n| g | acceleration of gravity, feet per second per second |", "timestamp": "2026-07-22T05:24:42.615177+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 58, "total_pages": 99, "image_filename": "19930082511_p58.jpg", "text": "56\nNACA TN No. 1826\n\nwhere the number of terms in the summation depends on the radius of the outer semicircle.\n\nFor $k \\ge 0$, if the radius of the outer semicircle is allowed to approach infinity in discrete steps so as to avoid the poles of the integrand, the integral over the semicircle approaches zero. The limiting values of the integrals along the two inner semicircles, as their radii approach zero, are readily determined by the usual process as\n\n$$\ni \\frac{J_m(-iph)e^{-ikh}}{4h^2 J_m'(-ih)}\n$$\n\nand\n\n$$\ni \\frac{J_m(iph)e^{ikh}}{4h^2 J_m'(ih)}\n$$\n\nThese two terms may be combined, after noting that reversing the sign of the argument in $J_m$ and $J_m'$ merely reverses the sign of their ratio, to\n\n$$\n- \\frac{J_m(iph)}{2h^2 J_m'(ih)} \\sin kh\n$$\n\nEquating the total integral along the infinite contour to the sum of the residues thus gives\n\n$$\n\\frac{1}{2\\pi i} \\int_{-\\infty}^{\\infty} \\frac{J_m(ipq)e^{ikq} dq}{iq J_m'(iq)(q^2 - h^2)} = \\frac{J_m(iph)}{2h^2 J_m'(ih)} \\sin kh - i \\sum_{s=0}^{\\infty} \\frac{J_m(i\\gamma_{sm})e^{-k\\gamma_{sm}}}{(\\gamma_{sm}^2 + h^2)(m^2 - \\gamma_{sm}^2)J_m(\\gamma_{sm})}\n$$\n\nwhere the $\\gamma_{sm}$ terms are now defined as the positive zeros of $J_m'$ instead of the negative zeros (if $J_m'(x) = 0$, so also does $J_m'(-x)$). By equating the imaginary part of the left-hand term to the right-hand side, which is a pure imaginary, there results, finally", "timestamp": "2026-07-22T05:24:43.840962+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 25, "total_pages": 36, "image_filename": "19930085869_p25.jpg", "text": "NACA RM L9D15 CONFIDENTIAL 23\n\n[Figure: (a) Spray strip similar to that on unswept model.]\n\n[Figure: (b) Final spray strip used throughout investigation. NACA L-60570]\n\nFigure 7.— Worst propeller spray condition for two spray-strip arrangements.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:24:46.052688+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 7, "total_pages": 42, "image_filename": "19930085914_p7.jpg", "text": "6\nNACA RM A9D25\n\nBase-Pressure Corrections\n\nThe pressure on the base of the model fuselage was measured and, in an effort to correct for support interference, the drag data were corrected to correspond to a base pressure equal to the static pressure of the free stream. The effect of longitudinal pressure gradient on drag was calculated and found to be negligible.\n\nTares\n\nSince the balance was within the model, there were no tares due to direct air forces on the model-support equipment. Corrections were made for the change in static tares due to the weight of the model and the variation of model attitude throughout the angle-of-attack range.\n\nPRECISION\n\nThe several sources of error affecting the accuracy of the results presented herein are listed below, along with an estimate of their magnitude.\n\nThe principal source of error in the data arises from the fact that the percentage accuracy of a given wire-resistance strain gage varies linearly with the absolute magnitude of the force imposed, and that the greatest percentage error occurs with the smallest applied force. The capacities of the gages used were governed by the large variation of forces encountered, and it was not practicable to change gages during the tests to improve the accuracy of the balance. The following table gives an estimate of the precision of the force and moment coefficients as determined from strain-gage calibrations for the limiting values of Mach number and Reynolds number:\n\n| | | Wing-fuselage combination | | | Fuselage alone |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| M | R | $0.8 \\times 10^6$ (percent) | $2.0 \\times 10^6$ (percent) | $9.0 \\times 10^6$ (percent) | $2.0 \\times 10^6$ (percent) |\n| | | $C_D$ | | | |\n| 0.20 | | 9 | 4 | 1 | 7 |\n| .93 | | 2 | 1 | - | 1 |\n| | | $C_L$ | | | |\n| .20 | | 1 | 1 | 0 | 4 |\n| .93 | | 0 | 0 | - | 3 |\n| | | $C_m$ | | | |\n| .20 | | 3 | 1 | 0 | 1 |\n| .93 | | 0 | 0 | - | 0 |", "timestamp": "2026-07-22T05:24:46.792015+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 17, "total_pages": 37, "image_filename": "19930085889_p17.jpg", "text": "```markdown\n16\nCONFIDENTIAL\nNACA RM L9F14\n\nREFERENCES\n\n1. MacLachlan, Robert, and Letko, William: Correlation of Two\nExperimental Methods of Determining the Rolling Characteristics\nof Unswept Wings. NACA TN 1309, 1947.\n\n2. Goodman, Alex, and Fisher, Lewis R.: Investigation at Low Speeds\nof the Effect of Aspect Ratio and Sweep on Rolling Stability\nDerivatives of Untapered Wings. NACA TN 1835, 1949.\n\n3. Brewer, Jack D., and Fisher, Lewis R.: Effect of Taper Ratio on\nthe Low-Speed Rolling Stability Derivatives of Swept and Unswept\nWings of Aspect Ratio 2.61. NACA RM L8H18, 1948.\n\n4. Queijo, M. J., and Jaquet, Byron M.: Calculated Effects of\nGeometric Dihedral on the Low-Speed Rolling Derivatives of Swept\nWings. NACA TN 1732, 1948.\n\n5. Letko, William, and Brewer, Jack D.: Effect of Airfoil Profile of\nSymmetrical Sections on the Low-Speed Rolling Derivatives of\n$45^\\circ$ Sweptback-Wing Models of Aspect Ratio 2.61. NACA RM L8L31a,\n1949.\n\n6. Kuhn, Richard E., and Myers, Boyd C., II: Effects of Mach Number\nand Sweep on the Damping-In-Roll Characteristics of Wings of\nAspect Ratio 4. NACA RM L9E10, 1949.\n\n7. Feigenbaum, David, and Goodman, Alex: Preliminary Investigation at\nLow Speeds of Swept Wings in Rolling Flow. NACA RM L7E09, 1947.\n\n8. DeYoung, John: Theoretical Additional Span Loading Characteristics\nof Wings with Arbitrary Sweep, Aspect Ratio, and Taper Ratio.\nNACA TN 1491, 1947.\n\n9. Jacobs, Eastman N., and Ward, Kenneth E.: Interference of Wing and\nFuselage from Tests of 209 Combinations in the N.A.C.A. Variable-\nDensity Tunnel. NACA Rep. 540, 1935.\n\n10. Goodman, Alex, and Brewer, Jack D.: Investigation at Low Speeds of\nthe Effect of Aspect Ratio and Sweep on Static and Yawing\nStability Derivatives of Untapered Wings. NACA TN 1669, 1948.\n\n11. Toll, Thomas A., and Queijo, M. J.: Approximate Relations and\nCharts for Low-Speed Stability Derivatives of Swept Wings. NACA\nTN 1581, 1948.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:24:47.472715+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 48, "total_pages": 66, "image_filename": "19930082914_p48.jpg", "text": "NACA TN No. 1857\n47\n\n[Figure: Interferogram of jet showing horizontal fringe patterns with distortions and discontinuities, including a region of compressed fringes near the center and vertical markers along the bottom edge.]\n\n(a) 0 to $1\\frac{1}{2}$ inches from nozzle.\n\nFigure 11.— Interferogram of jet.\n\nNACA", "timestamp": "2026-07-22T05:24:49.101113+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 1, "total_pages": 18, "image_filename": "19930085964_p1.jpg", "text": "RESTRICTED\nCopy\nRM E9G25\n227\n\nNACA RM E9G25\n\nNACA\nRESEARCH MEMORANDUM\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nAUTHORITY CROWLEY CHANGE #1919\nDATE 12-14-53\nT.C.F.\n\nVIBRATIONAL MODES OF SEVERAL HOLLOW TURBINE BLADES AND OF\nSOLID TURBINE BLADE OF SIMILAR AERODYNAMIC DESIGN\n\nBy R. H. Kemp and J. Shifman\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\ntitle 50 U.S.C. 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\nOctober 3, 1949\n\nRESTRICTED", "timestamp": "2026-07-22T05:24:52.786398+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 42, "total_pages": 78, "image_filename": "19930082618_p42.jpg", "text": "```markdown\n2.8\n2.4\nSection lift coefficient, $c_l$\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-1.6\nMoment coefficient, $c_{m_{c/4}}$\n-16\n-8\n0\n8\n16\nSection angle of attack, $\\alpha_c$, deg\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\nabla$ 6.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(b) Section lift and pitching-moment characteristics of the NACA 64$_1$-612 airfoil section with a\n0.20c simulated split flap deflected 60$^\\circ$.\n\nFigure 7.— Continued.\n\n40\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:24:53.717986+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 21, "total_pages": 96, "image_filename": "19930085880_p21.jpg", "text": "NACA RM No. L9C03\n19\n\n.03\nArea forward of observed\nwetted length, sq ft\n.02\n.01\n0\n0\n4\n8\n12\n16\n20\nTrim, deg\n(a) Model 250B.\n\n.03\nArea forward of observed\nwetted length, sq ft\n.02\n.01\n0\n0\n.1\n.2\n.3\n.4\nTotal wetted area, sq ft\n(b) Model 250D.\n\nFigure 8.- Wetted area forward of the observed wetted length for\nmodels 250B and 250D.", "timestamp": "2026-07-22T05:24:55.034053+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 16, "total_pages": 31, "image_filename": "19930085881_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:24:55.206641+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 1, "total_pages": 51, "image_filename": "19930085962_p1.jpg", "text": "```markdown\nCopy 357\nRM A9E05\n\nNACA RM A9E05\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nTHE AERODYNAMIC CHARACTERISTICS THROUGHOUT THE\nSUBSONIC SPEED RANGE OF A THIN, SHARP-EDGED\nHORIZONTAL TAIL OF ASPECT RATIO 4 EQUIPPED\nWITH A CONSTANT-CHORD ELEVATOR\n\nBy Angelo Bandettini and Verlin D. Reed\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nCHANGE # 2323\nAUTHORITY OF J.W. CROWLEY\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 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.\n29/54\nJ.L.C.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nJune 30, 1949\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:24:59.806062+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 8, "total_pages": 36, "image_filename": "19930085912_p8.jpg", "text": "6\nNACA RM No. E9C16\n\nAerodynamic investigation with hot-gas bleedback. - Several\norifice configurations were investigated in order to obtain a con-\nfiguration that would give a uniform temperature distribution inside\nthe model for a range of values of tunnel velocity, angle of attack,\ngas flow, and gas temperature. The configuration selected is shown\nin figure 3 and consists of three $\\frac{3}{4}$-inch, three $\\frac{1}{2}$-inch, and six $\\frac{13}{32}$-inch\norifices. Calculated jet outlines and penetrations are also shown in\nfigure 3 and the method by which these outlines and penetrations were\nobtained is discussed in the appendix.\n\nThe effect of hot-gas bleedback on the temperature distribution\ninside the model, on the air-flow characteristics, and on the ram-\npressure recovery for the optimum orifice configuration was determined\nas a function of tunnel velocity, angle of attack, gas flow, and gas\ntemperature. The investigation was conducted at a fixed tail-cone\nposition corresponding to an inlet-velocity ratio of 0.82 without\nbleedback and at a free-stream total temperature of $0^\\circ$ F for tunnel\nvelocities from 200 to 470 feet per second and angles of attack from\n$0^\\circ$ to $8^\\circ$. Gas flows and plenum-chamber gas temperatures ranged from\n0.52 to 1.73 pounds per second and from approximately $600^\\circ$ to $1000^\\circ$ F,\nrespectively. For each plenum-chamber gas temperature, the gas pressure\nwas varied from 2900 to 6000 pounds per square foot absolute.\n\nIcing with hot-gas bleedback. - An investigation to determine the\ncritical-icing criterion as a function of mass air flow, gas flow,\nangle of attack, and liquid-water content for a constant free-stream\ntotal temperature of $0^\\circ$ F was conducted in the same manner as that of\nreference 1. This investigation was conducted at tunnel velocities of\n200, 280, 360, and 410 feet per second at angles of attack of $0^\\circ$ and\n$8^\\circ$. The liquid-water content ranged from 0.3 to 1.0 gram per cubic\nmeter at an average drop diameter of 15 microns. The range of gas\nflows and plenum-chamber gas temperatures was the same as that employed\nfor the aerodynamic investigation with hot-gas bleedback.\n\nRESULTS AND DISCUSSION\n\nAerodynamic Investigation without Bleedback\n\nMass-flow characteristics. - The mass flow through the model\nincreased nearly linearly with tunnel velocity for a fixed tail-cone\nposition and angle of attack. A maximum flow of approximately 31.4\npounds per second was obtained at an inlet-velocity ratio of 0.82, a\ntunnel velocity of 470 feet per second, and an angle of attack of $0^\\circ$.", "timestamp": "2026-07-22T05:25:02.259027+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 24, "total_pages": 46, "image_filename": "19930085542_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:25:02.837891+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 24, "total_pages": 29, "image_filename": "19930085879_p24.jpg", "text": "```markdown\n22\n\nDrag coefficient, $C_D$\n\nVelocity\n902 fps\n\nRear body\n\nVelocity\n9415 fps\n\nVelocity\n925 fps\n\nNose\n\nVelocity - 965 fps\n\nNACA\n\n0 4 8 12 16 20 24 28 32 36 40 44 48 52\nSeparation distance, ft\n\nFigure 10.- RM-11B drag coefficient curves during separation.\n\nNACA RM L9D11\n```", "timestamp": "2026-07-22T05:25:03.087427+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 37, "total_pages": 53, "image_filename": "19930082542_p37.jpg", "text": "NACA TN No. 1867\n37\n\n<!-- Image (84, 126, 890, 454) -->\n\n100X\n(a) As-rolled.\n1000X\n\n<!-- Image (84, 516, 881, 857) -->\n\n100X\n(b) 1800° F 2 hours.\n1000X\nNACA\n\nFigure 6.- Effect of solution-treating temperature on the micro-\nstructures of low-carbon N-155 bar stock. All specimens were\nwater-quenched from temperature. Electrolytic chromic acid\netch.", "timestamp": "2026-07-22T05:25:04.353308+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 40, "total_pages": 58, "image_filename": "19930082617_p40.jpg", "text": "NACA TN 1962\n39\n\nMoment at\nfailure\nCylinder (in. - lb)\n75 373,600\n77 324,700\n\nCylinder 75, band L\nCylinder 77, band O\n(longer cutout)\n\nDistance from horizontal diameter, in.\nStrain\n\n[Figure: Graph showing strain distribution for Cylinder 75 and Cylinder 77]\n\nNACA\n\nFigure 28.- Effect of length of cutout on strain distribution. Load, 500 pounds.", "timestamp": "2026-07-22T05:25:10.105460+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 5, "total_pages": 23, "image_filename": "19930085934_p5.jpg", "text": "4\nNACA RM E9G12\n\nIf conditions at the outlet are such that the entropy at the\noutlet equals that at the inlet, the second law of thermodynamics\nprecludes the possibility of any change in entropy during the proc-\ness; therefore only the end points of the process are considered.\nEvaluation of the entropy of the mixture at the inlet is difficult\nbecause an increase in entropy due to adiabatic saturation of the\nair at the inlet cannot be avoided. Only if the air is saturated\nbefore the water is injected and is at the same temperature as the\ninjected water can the process conceivably be ideally reversible.\nFor a fixed set of inlet conditions and given compression ratio,\nthe pressure at the outlet is fixed. The temperature at the outlet\nwill therefore determine not only the state of the injected liquid\nbut also the entropy at the outlet. A temperature that will give\nan entropy at the outlet equal to that at the inlet must therefore\nbe found by trial and error. When this temperature is found, the\nenthalpy of the outlet mixture can be determined.\n\nFrom the Gibbs-Dalton law it may be deduced that the entropy\nand the enthalpy of a mixture of gases and vapors can be regarded\nas the sum of the individual entropies and enthalpies of the con-\nstituents when each occupies the same space as the mixture at the\ntemperature of the mixture. This law is not extended to liquid\nparticles suspended in a gas. For the method presented herein, how-\never, the liquid-water particles are assumed to be homogeneously\ndistributed throughout the mixture but to exert no pressure. The\nentropy and the enthalpy for the particles are calculated in the\nsame manner as for gases.\n\nIn computing the entropy of the mixture at the inlet, no attempt\nis made to estimate the amount of water evaporated into the air on\ninjection. Instead, the entropy of the mixture is based on condi-\ntions existing before any degree of vaporization of injected water\noccurs. The assumption of no vaporization at the compressor inlet\nsimplifies the calculation procedure. Data based on no saturation\nand complete saturation of the inlet air indicate only a small dif-\nference in enthalpy change; however, the condition of no saturation\nresults in a more conservative value of efficiency. The temperature\nof the water particles is assumed equal to that of the water before\ninjection in order to approximate more closely the value of the\ninlet entropy of the mixture. All values of entropy are taken from\nthe steam and air tables of references 1 and 2, respectively. The\nentropy of each constituent in the mixture is based on 1 pound of\ndry air, with the result that the sum of all the entropies gives\nthe entropy of the mixture per pound of dry air. The enthalpy of\nthe mixture at the inlet is determined by the same procedure.\n\n1164", "timestamp": "2026-07-22T05:25:13.318850+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 10, "total_pages": 23, "image_filename": "19930085906_p10.jpg", "text": "NACA RM E9F20\n\nCONFIDENTIAL\n\n[Figure: Schematic diagram of 20-inch-diameter ram jet]\n\nFuel preheater\nFlame holder\nFuel injector\n\n17\"\n20\"\n64½\"\n42¾\"\n14\"\n4\"\n\nNozzle\nCombustion chamber\nDiffuser\n\nDistance between preheater\nand flame holder\n\nAir flow\n\nFigure 1. - Schematic diagram of 20-inch-diameter ram jet.\n\nNACA\n\nCONFIDENTIAL\n\n9", "timestamp": "2026-07-22T05:25:15.293487+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 31, "total_pages": 33, "image_filename": "19930085544_p31.jpg", "text": "30\nNACA RM No. L8K26\n\n<!-- Image (93, 110, 804, 836) -->\n\nFigure 13.- Variation of force coefficient with k and with $\\epsilon \\frac{\\alpha}{\\alpha_{P_0}}$.\n($\\epsilon < 0.1$).", "timestamp": "2026-07-22T05:25:16.821001+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 26, "total_pages": 36, "image_filename": "19930085869_p26.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:25:17.780448+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 49, "total_pages": 66, "image_filename": "19930082914_p49.jpg", "text": "48\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:25:23.389525+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 4, "total_pages": 42, "image_filename": "19930085938_p4.jpg", "text": "```markdown\nNACA RM No. L9B04\n3\n\nb maximum beam of hull (6.43 ft, full size)\n\nw specific weight of water (63.0 lb/cu ft in these tests)\n\n$\\bar{c}$ mean aerodynamic chord\n\nDESCRIPTION OF MODELS\n\nThe model having an afterbody consisting of a single boom was designated Langley tank model 237-7B. Photographs of this model are shown in figures 2(a) and 2(b). The general arrangement and hull lines are shown in figures 3 and 4, respectively. The model with the twin booms was designated Langley tank model 237-7TB. Photographs of this model are shown in figures 5(a) and 5(b). General arrangement and hull lines are shown in figures 6 and 7, respectively. Offsets for both configurations are given in reference 3.\n\nThe forebody plan form was a modified 16-series symmetrical airfoil section with length-beam ratio of 7.0. The upper portion of the hull as used in the wind tunnel was not reproduced. In order to provide adequate spray control, chine strips of 0.05b depth were used on both configurations. On the single-boom configuration, the chine strips extended from 0.5b aft of the nose to the point of the step where they were faired to zero depth. The chine strips on the twin-boom configuration extended from 0.5b aft of the nose to 1.45b forward of the point of the step. The booms of either configuration were simple cones of circular cross section. The twin-boom configuration had two conical booms, one faired out of each engine nacelle. The nacelles of this configuration were moved outboard to reduce the interference between the forebody wake and the booms. Both configurations had slightly shorter booms than those tested in the wind tunnel, but it is believed this difference would have no appreciable effect on the aerodynamic characteristics.\n\nThese models were $\\frac{1}{16}$-size powered dynamic models of a hypothetical flying boat of 65,000 pounds gross load ($C_{\\Delta_0} = 3.87$). The wing and power used for both configurations corresponded to those of the Boeing XPBB-1 which resulted in a wing loading of 35.6 pounds per square foot and a power loading of 14.8 pounds per brake horsepower for the hypothetical design. The wing was located as shown in figures 3 and 6. The wing incidence relative to the base line was $4^\\circ$. The tail surfaces of the single-boom configuration were those of the Boeing XPBB-1 to $\\frac{1}{16}$ scale. The area of the horizontal tail surfaces of the twin-boom configuration was the same as that of the single boom, but the shape and arrangement were altered to facilitate mounting between the vertical fins. The total area of the vertical fins was approximately 1.75 times that used on the single-boom model.\n```", "timestamp": "2026-07-22T05:25:26.664249+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 18, "total_pages": 37, "image_filename": "19930085889_p18.jpg", "text": "NACA RM L9F14 CONFIDENTIAL 17\n\n12. Bamber, M. J., and House, R. O.: Wind-Tunnel Investigation of Effect of Yaw on Lateral-Stability Characteristics. I – Four N.A.C.A. 23012 Wings of Various Plan Forms with and without Dihedral. NACA TN 703, 1939.\n\n13. Recant, Isidore G., and Wallace, Arthur R.: Wind-Tunnel Investigation of Effect of Yaw on Lateral-Stability Characteristics. III – Symmetrically Tapered Wing at Various Positions on Circular Fuselage with and without a Vertical Tail. NACA TN 825, 1941.\n\n14. Salmi, Reino J., Conner, D. William, and Graham, Robert R.: Effects of a Fuselage on the Aerodynamic Characteristics of a $42^\\circ$ Sweptback Wing at Reynolds Numbers to 8,000,000. NACA RM L7E13, 1947.\n\n15. Neely, Robert H., and Conner, D. William: Aerodynamic Characteristics of a $42^\\circ$ Swept-Back Wing with Aspect Ratio 4 and NACA 64$_1$-112 Airfoil Sections at Reynolds Numbers from 1,700,000 to 9,500,000. NACA RM L7D14, 1947.\n\n16. Bird, John D.: Some Theoretical Low-Speed Span Loading Characteristics of Swept Wings in Roll and Sideslip. NACA TN 1839, 1949.\n\n17. Bird, John D., Lichtenstein, Jacob H., and Jaquet, Byron M.: Investigation of the Influence of Fuselage and Tail Surfaces on Low-Speed Static Stability and Rolling Characteristics of a Swept-Wing Model. NACA RM L7H15, 1947.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:25:27.484857+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 59, "total_pages": 99, "image_filename": "19930082511_p59.jpg", "text": "NACA TN No. 1826\n\n$$\n\\int_{0}^{\\infty} \\frac{J_{m}(i p q) \\cos k q \\, d q}{i q J_{m}'(i q)(q^{2} - h^{2})} = \\frac{\\pi i J_{m}(i p h)}{2 h^{2} J_{m}'(i h)} \\sin k h\n$$\n\n$$\n+ \\pi \\sum_{s=0}^{\\infty} \\frac{J_{m}(\\rho y_{sm}) e^{-k y_{sm} y_{sm}}}{\\left(\\sqrt{y_{sm}^{2} + h^{2}}\\right)\\left(m^{2} - y_{sm}^{2}\\right) J_{m}(y_{sm})}\n$$\n\nExpressions for $P_{mn}$, $n \\neq 0$, $m \\neq 0$.— In the preceding development it was assumed that $k \\geq 0$, which was acceptable with regard to equation (A2) in view of the fact that the cosine is an even function of the variable. This essentially nonnegative value of $k$ must be retained, however, in the final expressions for $P_{mn}(\\xi, \\rho)$:\n\n$$\nP_{mn} = -\\frac{1}{2} \\frac{J_{m}\\left(i \\rho \\frac{n \\pi}{b-a}\\right)}{i \\frac{n \\pi}{b-a} J_{m}'\\left(i \\frac{n \\pi}{b-a}\\right)} \\left[ (-1)^{n} \\sin \\frac{|b-\\xi| n \\pi}{b-a} - \\sin \\frac{|\\xi-a| n \\pi}{b-a} \\right]\n$$\n\n$$\n+ \\frac{n \\pi}{b-a} \\sum_{s=0}^{\\infty} \\frac{J_{m}(\\rho y_{sm}) y_{sm}}{\\left[\\left(y_{sm}\\right)^{2} + \\left(\\frac{n \\pi}{b-a}\\right)^{2}\\right]\\left[m^{2} - \\left(y_{sm}\\right)^{2}\\right] J_{m}(y_{sm})} \\left[ (-1)^{n} e^{-|b-\\xi| y_{sm}} - e^{-|\\xi-a| y_{sm}} \\right]\n$$\n\n(A3)\n\nNow for $a \\leq \\xi \\leq b$\n\n$$\n(-1)^{n} \\sin \\frac{|b-\\xi| n \\pi}{b-a} = (-1)^{n} \\sin \\left[ \\frac{b-a-(\\xi-a)}{b-a} n \\pi \\right]\n$$\n\n$$\n= -\\sin n \\pi \\frac{\\xi-a}{b-a}\n$$\n\n$$\n= -\\sin n \\pi \\frac{|\\xi-a|}{b-a}\n$$", "timestamp": "2026-07-22T05:25:31.559407+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 8, "total_pages": 42, "image_filename": "19930085914_p8.jpg", "text": "NACA RM A9D25\n\nCalibration of the strain-gage balance indicated that interactions due to deformation of gage members were negligible. Corrections were made for zero shift of the strain indicating instruments.\n\nAnother possible source of error in the results was friction in the balance. The effect of friction was largest on the drag measurements of the fuselage alone where the drag force imposed by the weight of the fuselage was large compared to the aerodynamic drag of the fuselage. Reasonably good indication that the effect of frictional forces on the other components was small is the fact that, in general, experimental scatter lies within the limits of error given in the preceding table.\n\nThe angle of attack of the model was observed visually by means of a cathetometer. From numerous test readings it was determined that the angle of attack could be set repeatedly within $\\pm 0.15^\\circ$.\n\nRESULTS AND DISCUSSION\n\nEffects of Mach Number\n\nGeneral aerodynamic characteristics.— General aerodynamic characteristics of the wing-fuselage combination are presented in figures 4 and 5 for Mach numbers from 0.20 to 0.93 and Reynolds numbers of 0.8 million and 2.0 million, respectively. The drag variation with lift (figs. 4(a) and 5(a)) shows no pronounced effect of Mach number. The angle of attack for minimum drag was about $0^\\circ$ throughout the Mach number range. The values of drag coefficient were abnormally low at low lift coefficients for 0.40 Mach number at 0.8 million Reynolds number and for 0.20, 0.40, and 0.70 Mach number at 2.0 million Reynolds number. These small magnitudes are attributed to malfunction of the strain-gage balance rather than to a characteristic of the model.\n\nNo pronounced effect of Mach number is noted in the variation of lift coefficient with angle of attack (figs. 4(b) and 5(b)). The angle of attack for zero lift was about $0.5^\\circ$ at a Mach number of 0.20 and increased gradually to about $1.0^\\circ$ at a Mach number of 0.93. A slight decrease in lift-curve slope is noted at a lift coefficient of about 0.2, with subsequent recovery to a value even greater than that at zero lift. Neither the severity of this reduction of slope nor the lift coefficient at which it occurred was affected by Mach number. A corresponding forward movement of the aerodynamic center is discernible from the pitching-moment data (figs. 4(c) and 5(c)) over the range of lift coefficients affected, with subsequent rearward movement to a location generally behind that at zero lift. There was complete loss of static longitudinal stability at the higher lift coefficients. (The lift coefficient at which instability occurred had no consistent variation with Mach number, but was between 0.5 and 0.6 for most of the test Mach numbers.) This trend is typical of the stalling characteristics peculiar to wings with large amounts of sweep (references 6, 7, and 8).", "timestamp": "2026-07-22T05:25:33.114815+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 22, "total_pages": 96, "image_filename": "19930085880_p22.jpg", "text": "20\nNACA RM No. L9C03\n\n[Figure: A line graph plotting Resistance (lb) against Speed (fps). The y-axis ranges from 0 to .70. The x-axis ranges from 0 to 35. There are five lines representing different Draft (ft) values: 0, 0.2, 0.4, 0.6, and 0.8. The NACA logo is present in the bottom right corner of the graph.]\n\nFigure 9.- Aerodynamic drag of towing gear less model.", "timestamp": "2026-07-22T05:25:35.978243+00:00"}
{"citation_id": "19930085964", "source_url": "https://ntrs.nasa.gov/api/citations/19930085964/downloads/19930085964.pdf", "page_number": 2, "total_pages": 18, "image_filename": "19930085964_p2.jpg", "text": "NACA RM E9G25 RESTRICTED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nVIBRATIONAL MODES OF SEVERAL HOLLOW TURBINE BLADES AND OF\nSOLID TURBINE BLADE OF SIMILAR AERODYNAMIC DESIGN\n\nBy R. H. Kemp and J. Shifman\n\nSUMMARY\n\nThe vibrational modes of several hollow turbine blades and a\nsolid turbine blade of similar aerodynamic design were experi-\nmentally determined. In general, several vibrational character-\nistics common to the hollow blades investigated were as follows:\n\n1. Approximately twice as many readily excited modes were\ndetected in the hollow blades than in the solid blade; nearly all\nthe vibrational modes of the hollow blades could be excited if the\nblades were operated in a conventional turbojet engine.\n\n2. An effect termed \"breathing\" was found to be predominant\nin a vibrational mode occurring at approximately 1230 cycles per\nsecond. In this mode, the two sides of the blade move alternately\ntoward and then away from each other causing stress concentrations\nto be produced at the leading and trailing edges and at the rivet\nholes of a rivet-stiffened blade. Failure by fatigue in this mode\nof a hollow blade having rivets along the trailing edge produced\ncracks that started at the rivet holes and extended to the tip of\nthe blade.\n\n3. Nodal patterns of the vibrational modes below approximately\n2600 cycles per second were similar on the two sides of the blades,\nwhereas in the higher frequency modes the nodal patterns were dis-\nsimilar on the two sides. The hollow blades vibrated as an integral\nunit in the lower modes and were characterized by plate vibrations\nin the higher modes.\n\n4. Nodal patterns of the hollow blades of same design and\nmanufacture were similar in the lower modes (those below approxi-\nmately 2600 cps) but were dissimilar in the higher modes.\n\n5. No vibrational modes of the hollow blades were detected\nthat could be designated true fundamental bending modes. The lowest\nfrequency mode occurred at approximately 850 cycles per second.\n\nRESTRICTED", "timestamp": "2026-07-22T05:25:38.415019+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 17, "total_pages": 31, "image_filename": "19930085881_p17.jpg", "text": "NACA RM L9D12 CONFIDENTIAL 15\n\n[Figure: A model rocket with a pointed nose cone, cylindrical body, and three fins at the base. A scale in inches is visible near the base. A label in the bottom right corner reads \"NACA L-57479\".]\n\n(b) $\\Lambda = 45^\\circ$.\n\nFigure 2.— Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:25:38.452707+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 25, "total_pages": 46, "image_filename": "19930085542_p25.jpg", "text": "NACA RM No. L6L29\n\n[Figure: Model 7 mounted in tunnel]\n\nNACA\nL-55508\n\nFigure 6.- Model 7 mounted in tunnel. $A = 4$; $\\Lambda_{c/4} = 36.9^\\circ$; profile, NACA 0012.\n\n63", "timestamp": "2026-07-22T05:25:41.770179+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 43, "total_pages": 78, "image_filename": "19930082618_p43.jpg", "text": "```markdown\nNACA TN 1945\n\nSection drag coefficient, $c_d$\n.024\n.020\n.016\n.012\n.008\n.004\n0\n-.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\nR\n$\\circ$ $0.7 \\times 10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\nFlagged symbols denote\nstandard roughness\n\nSection drag coefficient, $c_d$\n.024\n.020\n.016\n.012\n.008\n.004\n0\n-1.2 -.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\nR\n$\\nabla$ $3.0 \\times 10^6$\n$\\triangle$ 6.0\n$\\blacktriangle$ 9.0\nFlagged symbols denote\nstandard roughness\n\nMoment coefficient, $c_{mac}$\n0\n0\n0\n0\n0\n0\n-.1\n-.2\n-.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\nR a.c. position\nx/c y/c\n$\\circ$ $0.7 \\times 10^6$ .260 -.011\n$\\square$ 1.0 .268 -.072\n$\\diamond$ 1.5 .271 -.048\n$\\triangle$ 2.0 .271 -.045\n$\\nabla$ 3.0 .270 -.028\n$\\blacktriangle$ 6.0 .265 -.039\n$\\blacktriangle$ 9.0 .264 .005\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic center of the plain NACA 64$_1$-612 airfoil section.\n\nFigure 7.- Concluded.\n\n41\n```", "timestamp": "2026-07-22T05:25:42.935840+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 2, "total_pages": 51, "image_filename": "19930085962_p2.jpg", "text": "NACA RM No. A9E05 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nTHE AERODYNAMIC CHARACTERISTICS THROUGHOUT THE SUBSONIC SPEED RANGE OF A THIN, SHARP-EDGED HORIZONTAL TAIL OF ASPECT RATIO 4 EQUIPPED WITH A CONSTANT-CHORD ELEVATOR\n\nBy Angelo Bandettini and Verlin D. Reed\n\nSUMMARY\n\nWind-tunnel tests have been made of a semispan model of a horizontal tail of aspect ratio 4 and taper ratio 0.5 equipped with a full-span constant-chord elevator having an area equal to 20 percent of the total semispan-tail area. The horizontal tail was not swept and the profile was a sharp-edged, faired double wedge with a thickness-chord ratio of 0.042. Lift, drag, and pitching-moment data are presented for a Reynolds number of 2,000,000 at Mach numbers from 0.20 to 0.94.\n\nAt small angles of attack and small elevator deflections, the effect of compressibility on the effectiveness of the elevator in producing lift was small at Mach numbers less than 0.60. There was a gradual increase in effectiveness between Mach numbers of 0.60 and 0.90 to a value equal to 146 percent of the low-speed value. At higher Mach numbers the effectiveness decreased slightly to a value at a Mach number of 0.94 equal to 127 percent of the low-speed value. The variation of lift coefficient with elevator deflection was not linear, the effectiveness being lower for deflections from $0^\\circ$ to $2^\\circ$ than for deflections from $2^\\circ$ to $4^\\circ$. Neither the magnitude nor extent of this nonlinearity in effectiveness at small elevator deflections was aggravated by compressibility.\n\nINTRODUCTION\n\nRecent investigations have indicated several wing plan forms, wing sections, and wing-body-tail combinations suitable for flight at supersonic speeds. One such lifting surface, a thin, sharp-edged wing without sweep of aspect ratio 4 and taper ratio 0.5, has been\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:25:45.472532+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 42, "total_pages": 46, "image_filename": "19930085519_p42.jpg", "text": "NACA RM No. L8K19\n41\n\n$\\delta p$\n(percent local wing chord)\nActuating arms\n$\\circ$ -3 filled-in\n$\\triangle$ -5 filled-in\n$\\square$ -7 filled-in\n$\\diamond$ -3 open\n$\\triangleleft$ -5 open\n$\\triangleright$ -7 open\n\n<!-- Image (168, 283, 894, 764) -->\n\n(a) Flap retracted.\n\nFigure 20.- The effect of plug-aileron actuating-arm configuration on the variation of rolling-moment and yawing-moment coefficients with angle of attack at various plug-aileron projections on the 42° sweptback wing. Faired plug-slot lower lip.", "timestamp": "2026-07-22T05:25:51.340803+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 25, "total_pages": 29, "image_filename": "19930085879_p25.jpg", "text": "NACA RM L9D11\n\n[Figure: Graph plotting Velocity (ft/sec) on the y-axis (ranging from 400 to 1100) against Time from launching, sec on the x-axis (ranging from 4 to 14). The graph shows two curves: one labeled \"Velocity nose\" and another labeled \"Velocity rear body\". Annotations on the graph include \"Separation\" pointing to a point on the curves around time 10 sec, and \"Flaps start to open\" pointing to a point on the curves around time 11 sec. The NACA logo is present in the bottom right corner of the graph area.]\n\nTime from launching, sec\n\nFigure 11.— RM-11B velocity plotted against time from Doppler radar.\n\n23", "timestamp": "2026-07-22T05:25:51.907841+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 38, "total_pages": 53, "image_filename": "19930082542_p38.jpg", "text": "NACA TN No. 1867\n39\n\n[Figure: Micrographs of material microstructure at different magnifications and heat treatments]\n\n100X\n(c) $1950^\\circ$ F 2 hours.\n1000X\n\n100X\n(d) $2050^\\circ$ F 2 hours.\n1000X\n[Figure: NACA logo]\n\nFigure 6.- Continued.", "timestamp": "2026-07-22T05:25:52.906020+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 9, "total_pages": 36, "image_filename": "19930085912_p9.jpg", "text": "NACA RM No. E9C16\n\nRam-pressure recovery. - A ram-pressure recovery $\\eta$ of approximately 0.95 was obtained at an angle of attack of $0^\\circ$ and an inlet-velocity ratio of 0.82. No appreciable change in ram-pressure recovery was observed for an increase in angle of attack from $0^\\circ$ to $8^\\circ$ and a decrease in inlet-velocity ratio from 0.82 to 0.64.\n\nLip-pressure distribution. - The effect of angle of attack on lip-pressure distribution is shown in figure 4. The pressure distribution is presented in terms of a pressure coefficient\n\n$$\nS = 1 - \\left( \\frac{p - p_0}{q_0} \\right)\n$$\n\nAerodynamic Investigation with Cold-Gas Bleedback\n\nMass-flow characteristics. - No measurable decrease in mass flow through the model was observed with increasing bleedback. A decreasing inlet-velocity ratio must therefore occur with increasing bleedback because an increasing part of the total flow through the model is represented by the bleedback gas and as a consequence the flow entering the inlet is reduced.\n\nRam-pressure recovery. - In order to determine the effect of the jets alone, cold gas was bled into the inlet air stream; the effect of bleedback on ram-pressure recovery $\\eta$ is shown in figure 5 for a free-stream total temperature of $0^\\circ$ F and tunnel velocities of 220, 300, 380, and 450 feet per second. Figure 5 shows that the loss in ram-pressure recovery is linearly related to bleedback and no effect of velocity on ram-pressure recovery is evident.\n\nLip-pressure distribution. - A slight movement of the stagnation point to a position farther inside the lip was observed with cold-gas bleedback. This movement increased with increasing bleedback (fig. 6) and was caused by the decrease in the inlet-velocity ratio with increasing bleedback. The decreased inlet velocity was further evidenced by the reduced pressure coefficients in the inlet.\n\nAerodynamic Investigation with Hot-Gas Bleedback\n\nOptimum orifice configuration. - Several orifice configurations were investigated in order to obtain an optimum configuration that would give the most uniform temperature distribution at the simulated engine inlet. Results are presented for the orifice configuration that gave the most uniform temperature distribution at the calculated value of bleedback (4.4 percent) and gas temperature ($1000^\\circ$ F) necessary for adequate ice prevention corresponding to an icing condition", "timestamp": "2026-07-22T05:25:55.645916+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 41, "total_pages": 58, "image_filename": "19930082617_p41.jpg", "text": "40\nNACA TN 1962\n\nMoment at\nfailure\nCylinder (in-lb)\n75 373,600\n77 324,700\n\nCylinder 75, band L\nCylinder 77, band O\n(longer cutout)\n\nDistance from horizontal diameter, in.\nStrain\n\nFigure 29.- Effect of length of cutout on strain distribution. Load, 3000 pounds.", "timestamp": "2026-07-22T05:25:56.712576+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 11, "total_pages": 23, "image_filename": "19930085906_p11.jpg", "text": "```markdown\n10\nCONFIDENTIAL\nNACA RM E9F20\n\nTemperature, °F\nFuel evaporated, percent\n\n[Figure: A line graph plotting Temperature (°F) on the y-axis against Fuel evaporated (percent) on the x-axis. The curve starts near (0, 140) and rises to approximately (100, 340). The NACA logo is in the bottom right corner of the plot area.]\n\nFigure 2. - A.S.T.M. distillation curve for AN-F-28 fuel used in 20-inch-diameter ram jet.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:26:02.443799+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 6, "total_pages": 23, "image_filename": "19930085934_p6.jpg", "text": "NACA RM E9G12\n\nIn order to compute the outlet temperature for which the entropy at the outlet equals that at the inlet, an outlet temperature is first assumed. With the assumed temperature and the water-air ratio at the outlet, the entropy is computed in the same manner as for the inlet. When the isentropic outlet temperature is found, either of two sets of conditions may prevail at the outlet depending on the quantity of water injected: The air may be saturated at the outlet with or without water droplets being present, or the vapor may be in the superheated condition, in which case no droplets will be present. In the case of saturation with or without water droplets, the entropy can be easily found because these properties are the same as those for saturation temperature. With the superheated vapor, the exact temperature at which saturation occurs must be determined. The pressure corresponding to the temperature of saturation is the same that the vapor exerts at any higher temperature, the pressure of the mixture remaining constant. When this pressure is determined, the entropy of the vapor can be obtained from the superheated-vapor tables of reference 1. The enthalpy of the mixture at the outlet for either the isentropic or the actual case is determined by the same procedure.\n\nIf the mixture at the outlet contains droplets in a mixture of superheated vapor and air, the actual outlet temperature cannot be used. The compressor work, or actual enthalpy rise, must then be determined by mechanical means. An estimate of the quantity of droplets present can be made by the method illustrated in the Numerical Example if the dry-bulb temperature at the outlet is known. The humidity due to the superheated vapor at the outlet can be approximated by trial and error by finding a vapor-air ratio at the outlet that gives an enthalpy change equal to the power input. The procedure is the same as that for calculating the actual outlet temperature except that the vapor-air ratio must be varied for succeeding computations. Any water not in the superheat state is considered to be in droplet form.\n\nNUMERICAL EXAMPLE\n\nA numerical example is presented to show the method of computing isentropic enthalpy rise, actual enthalpy rise, and adiabatic efficiency of a compressor for a given set of conditions. In each case where a temperature is assumed for trial-and-error solution, only the calculations resulting in the correct temperature are shown. The following conditions from actual experimental data prevail:", "timestamp": "2026-07-22T05:26:07.181871+00:00"}