PDF/ocr_dataset_1051.jsonl

{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 17, "total_pages": 21, "image_filename": "19930082474_p17.jpg", "text": "NACA TN No. 1799\n15\n\n[Figure: A black and white photograph of a helicopter in flight against a cloudy sky. The helicopter has a skeletal frame, a single main rotor, and a tail rotor. It is marked with white crosses on its side and tail.]\n\nFigure 1.- General view of test helicopter.", "timestamp": "2026-07-22T05:55:32.603628+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 26, "total_pages": 114, "image_filename": "19930086061_p26.jpg", "text": "22\nNACA RM L9J07\n\nREFERENCES\n\n1. Wilson, Herbert A., Jr., and Lovell, J. Calvin: Full-Scale Investigation of the Maximum Lift and Flow Characteristics of an Airplane Having Approximately Triangular Plan Form. NACA RM L6K20, 1947.\n\n2. Whittle, Edward F., Jr., and Lovell, J. Calvin: Full-Scale Investigation of an Equilateral Triangular Wing Having 10-Percent-Thick Biconvex Airfoil Sections. NACA RM L8G05, 1948.\n\n3. Anderson, Adrien E.: An Investigation at Low Speed of a Large-Scale Triangular Wing of Aspect Ratio Two.- I. Characteristics of a Wing Having a Double-Wedge Airfoil Section with Maximum Thickness at 20-Percent Chord. NACA RM A7F06, 1947.\n\n4. Wick, Bradford H.: Chordwise and Spanwise Loadings Measured at Low Speed on a Triangular Wing Having an Aspect Ratio of Two and an NACA 0012 Airfoil Section. NACA TN 1650, 1948.\n\n5. Anderson, Adrien E.: Chordwise and Spanwise Loadings Measured at Low Speed on Large Triangular Wings. NACA RM A9B17, 1949.\n\n6. Orlik-Rückemann, K.: Experimental Determination of Pressure Distributions and Transition Lines of Plane Delta Wings at Low Speeds and Zero Yaw. KTH-Aero TN 3, Roy. Inst. of Technology, Div. of Aero., Stockholm, Sweden, 1948.\n\n7. Lange, Roy H., Whittle, Edward F., Jr., and Fink, Marvin P.: Investigation at Large Scale of the Pressure Distribution and Flow Phenomena over a Wing with the Leading Edge Swept Back $47.5^\\circ$ Having Circular-Arc Airfoil Sections and Equipped with Drooped-Nose and Plain Flaps. NACA RM L9G15, 1949.\n\n8. Underwood, William J., and Nuber, Robert J.: Aerodynamic Load Measurements over Leading-Edge and Trailing-Edge Plain Flaps on a 6-Percent-Thick Symmetrical Circular-Arc Airfoil Section. NACA RM L7H04, 1947.\n\n9. Theodorsen, Theodore, and Silverstein, Abe: Experimental Verification of the Theory of Wind-Tunnel Boundary Interference. NACA Rep. 478, 1934.\n\n10. Theodorsen, T., and Garrick, I. E.: General Potential Theory of Arbitrary Wing Sections. NACA Rep. 452, 1933.", "timestamp": "2026-07-22T05:55:34.001056+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 17, "total_pages": 29, "image_filename": "19930093789_p17.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:55:34.191780+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 24, "total_pages": 28, "image_filename": "19930086092_p24.jpg", "text": "22\nCONFIDENTIAL\nNACA RM A9F14\n\n<!-- Image (220, 112, 736, 814) -->\n\n(a) $C_m, C_D, C_L$ vs $\\beta$.\nFigure 5. - Aerodynamic characteristics at several angles of attack of the 63° swept-back wing with fuselage. Vertical tail on.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:55:34.690988+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 30, "total_pages": 50, "image_filename": "19930086076_p30.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:55:37.041410+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 3, "total_pages": 65, "image_filename": "19930082546_p3.jpg", "text": "2\nNACA TN No. 1870\n\nA section of the present paper is devoted to the fuselage response to these oscillating pressures and indicates some of the factors to be considered in solving the problems of fuselage vibration and noise.\n\nINTRODUCTION\n\nLarge-amplitude fuselage-wall vibrations in the region near the propeller plane have been experienced recently in several experimental airplanes. Fuselage-panel failures have occurred and great discomfort to the crew has resulted from the noise and vibration inside the airplane. These vibrations are known to result from the oscillating pressures associated with the rotating propeller. Up to the present time, however, very little information has been published that would enable a designer to predict these pressures in the critical region near the propeller tips.\n\nIn reference 1 Gutin has developed a theory by means of which the sound of a propeller may be predicted. By making several simplifying assumptions Gutin simplified the final equations, which were then useful only at a large distance from the propeller. The analysis presented herein is based on Gutin's fundamental equations without some of the simplifying assumptions of the original paper. The solution obtained then makes possible the prediction of oscillating pressures at any point in space. Its practical usefulness, however, is limited to the area close to the propeller tips, where Gutin's simplified solution is not valid. At a larger distance away the Gutin solution is much more convenient to use.\n\nStatic tests were made in which several different propeller models were used for comparison with analytical results. These tests evaluated the effects on the free-space oscillating-pressure distributions of such parameters as propeller diameter, blade plan form, number of blades, blade loading, tip clearance, and tip Mach number. Charts based on experimental data were calculated to enable a designer to estimate the average maximum free-space oscillating pressures in the critical region near the plane of rotation. Comparative data were obtained at the surface of two different simulated fuselage wall shapes to determine their effects on the free-space pressures. The fuselage response to these pressures is treated herein and indicates some of the factors to be considered in solving fuselage vibration and noise problems.\n\nSYMBOLS\n\n$R_e$ effective propeller radius\n\n$S$ distance between doublet and observer", "timestamp": "2026-07-22T05:55:37.806216+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 44, "total_pages": 51, "image_filename": "19930085962_p44.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 43\n\nLift coefficient, $C_L$\n\nMach number, M\n\n(d) $\\delta_e$, 20°.\n\nFigure 13.— Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:55:38.145206+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 39, "total_pages": 46, "image_filename": "19930085972_p39.jpg", "text": "NACA RM L9B18\n37\n\nDownwash angle, $\\epsilon$, deg\nAngle of attack, $\\alpha$, deg\n\nTail-off lift-curve slope, $(\\frac{dC_L}{d\\alpha})_o$\nHorizontal-tail effectiveness, $\\frac{dC_{H\\alpha}}{d\\alpha}$\n\nNeutral-point and tail-off aerodynamic-center location, $\\eta_p$ and $\\eta_o$, percent $\\bar{c}'(C_L=0)$\nLift coefficient, $C_L$\n\nDownwash gradient, $\\frac{d\\epsilon}{d\\alpha}$\nTail-off lift coefficient, $C_{L_o}$\n\n(c) External airfoil flaps.\nFigure 12.- Continued.", "timestamp": "2026-07-22T05:55:45.530961+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 63, "total_pages": 96, "image_filename": "19930085880_p63.jpg", "text": "NACA RM No. L9C03\n61\n\nResistance, lb\nSpeed\n(fps)\n30\n25\n20\n15\n10\n0\n.05\n.10\n.15\n.20\n.25\n.30\n.35\nWetted area, sq ft\n(b) $\\tau = 80$.\nFigure 19.- Continued.\nNACA", "timestamp": "2026-07-22T05:55:48.322793+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 14, "total_pages": 78, "image_filename": "19930082483_p14.jpg", "text": "```markdown\n12\nNACA TN No. 1807\n\nvelocity components of the driving gases through the turbine rotor\nand represent unavailable power in the driving fluid.\n\nAn additional loss may be introduced in a reaction-type tur-\nbine in which there is a tendency for some of the driving fluid to\ndiffuse at the nozzle discharge and to spread throughout the clear-\nance space between the nozzles and the rotor blading. The sub-\nsequent loss results chiefly from the unfavorable rotor-inlet\nvelocity component of these diffused gases.\n\nEach of these losses has a common characteristic, which per-\nmits them to be grouped into a single aggregate loss. The relating\ncharacteristic is that the magnitude of each loss is a function of\nthe product of the gas density and the rotor speed to the first\npower.\n\n$$\n\\left.\n\\begin{array}{l}\n\\text{scavenging loss} \\\\\n\\text{eddy loss} \\\\\n\\text{diffusion loss}\n\\end{array}\n\\right\\} \\sim f(K_{III} \\rho_1 N)\n$$\n\nThese losses in the driving fluid itself, which have been\ndesignated driving-fluid losses, are thus distinguished from the\nshaft and tip leakage losses in the text. Driving-fluid losses are\nnot reliably or readily evaluated from gas-state measurements\nbecause of the mixing action of the gases that produces them.\nAccordingly, they must be experimentally determined and compared\nwith a general expression developed to describe them, as follows:\n\nGross power and blade power are seen to be related by the\nequation\n\n$$\n\\text{gross power} = \\text{blade power} + \\text{driving-fluid losses} \\quad (25)\n$$\n\nBecause at $360^\\circ$ admission there are no significant driving-\nfluid losses, the driving-fluid losses can be determined from\nequations (19) and (25) as\n\n$$\n(\\text{driving-fluid losses})_F = F (\\text{blade power})_{360^\\circ} - (\\text{blade power})_F \\quad (26)\n$$\n\nUse of power expressions corrected to sea-level conditions\nresults in evaluation of driving-fluid losses corrected to sea-\nlevel conditions.\n\n1032\n```", "timestamp": "2026-07-22T05:55:51.900109+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 19, "total_pages": 34, "image_filename": "19930082472_p19.jpg", "text": "NACA TN NO. 1797\n\n[Figure: A large swept-forward wing mounted on two vertical supports inside a wind tunnel. A person stands near the base of one support for scale. The NACA logo and identifier \"A-11370\" are visible in the lower right corner of the image.]\n\nFigure 2.— The $45^\\circ$ swept-forward wing mounted in the 40- by 80-foot wind tunnel.\n\n17", "timestamp": "2026-07-22T05:55:56.219270+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 38, "total_pages": 42, "image_filename": "19930085938_p38.jpg", "text": "NACA RM No. L9B04\n37\n\nc.g. location\nper cent M.A.C.\n$\\circ$ 20\n$\\square$ 30\n$\\diamond$ 50\n\nAmplitude of vertical motion at c.g.\nper cent beam\nContact trim, deg\n(a) Amplitude of vertical motion.\n\nAmplitude of oscillillation in trim, deg\nContact trim, deg\n(b) Amplitude of trim oscillation.\n\nFigure 13.- Landing stability of twin-boom model.", "timestamp": "2026-07-22T05:55:56.654811+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 42, "total_pages": 42, "image_filename": "19930085914_p42.jpg", "text": "```markdown\nNACA-Langley\n\nNACA RM A9D25\n\n— Cambered and twisted wing ($l_c=.05$), $R=20 \\times 10^6$\n-- Plane wing of ref. 9 ($l_c=.06$), $R=2.35 \\times 10^6$\n\nLift coefficient, $C_L$\n\nPitching-moment coefficient, $C_m$\n\nfor M=.20\n\n(c) $C_L$ vs $C_m$.\n\nFigure 16.- Concluded.\n\nNACA\n\n41\n```", "timestamp": "2026-07-22T05:56:02.473020+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 31, "total_pages": 98, "image_filename": "19930086073_p31.jpg", "text": "```markdown\nNACA RM A59E04\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n\nLift coefficient, $C_L$\n\n0 0 0 0 4 8 12 16 20 24 28 32 36\nAngle of attack, $\\alpha$, deg\n\n$\\triangle$ $\\diamond$ $\\circ$ $\\square$\n44.0 21.2 0 -22.0\nFlap deflection, $\\delta_f$, deg\n\n(a) $C_L$ vs $\\alpha$.\n\nFigure 5.- Wing alone at 12.1° angle of sideslip with various flap deflections.\n\n29\n\n[Figure: NACA logo]\n```", "timestamp": "2026-07-22T05:56:03.733370+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 18, "total_pages": 21, "image_filename": "19930082474_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:56:08.395119+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 71, "total_pages": 78, "image_filename": "19930082618_p71.jpg", "text": "NACA TN 1945\n69\n\nSection lift-curve slope per degree, $d c_l / d \\alpha_0$\n\n.12\n.10\n.08\n$\\circ$ NACA 641-409\n$\\square$ NACA 641-412\n$\\diamond$ NACA 642-415\n$\\triangle$ NACA 643-418\n\n.12\n.10\n.08\n$\\circ$ NACA 641-012\n$\\square$ NACA 641-4212\n$\\diamond$ NACA 641-412\n$\\triangle$ NACA 641-612\n\n.12\n.10\n.08\n$\\circ$ NACA 652-415\n$\\square$ NACA 642-415\n$\\diamond$ NACA 652-415\n$\\triangle$ NACA 662-415\n\n.12\n.10\n.08\n.06\n$\\circ$ NACA 0012\n$\\square$ NACA 4412\n$\\diamond$ NACA 4415\n$\\triangle$ NACA 23012\n$\\nabla$ NACA 23015\n\n.5 1.0 2.0 3.0 4.0 5.0 10.0 x $10^6$\nReynolds number, R\n\n(b) Airfoils with standard leading-edge roughness.\n\nFigure 17.— Concluded.", "timestamp": "2026-07-22T05:56:10.115302+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 89, "total_pages": 99, "image_filename": "19930082511_p89.jpg", "text": "$$\\frac { \\Delta V } { V } = \\frac { 2 \\varepsilon h } { \\sigma b _ { 2 } }$$\n\nNACA TN NO. 1826\n\n<!-- Image (79, 133, 808, 822) -->\n\nDistance from entrance, $\\xi$, units of tunnel height\n\nFigure 17.- Tunnel-induced angle on axis of two-dimensional closed-open tunnel, with vortex at several locations along axis.\n\n87", "timestamp": "2026-07-22T05:56:11.794989+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 23, "total_pages": 36, "image_filename": "19930086097_p23.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 21\n\non a swept-back wing have indicated that the control effectiveness in the transonic range may be improved considerably by employing a blunt-trailing-edge aileron. (See, e.g., references 13 and 14.) Still another possible advantage that can be listed concerns the undesirable \"flat spot\" which has been observed in the hinge-moment curves of a conventional flap arrangement tested at low flap deflections and low angles of attack. (See reference 15.) This flat spot presumably would not occur for a control surface with a sufficiently blunt trailing edge, since the source of the trouble arises from separation of the flow forward of the trailing edge.\n\nThe various miscellaneous advantages such as those just mentioned, taken together with the general structural advantages and the improvement in certain aerodynamic characteristics, leave little doubt as to the practical usefulness of blunt-trailing-edge airfoils. Like many other examples of departure from conventional design, however, care must be exercised in designing airfoils with thick trailing edges. In this regard it is to be remembered that the highest Reynolds number in the present investigation is 1.2 million, and that additional experiments are needed before conclusions can be drawn about conditions at much higher Reynolds numbers.\n\nCONCLUSIONS\n\nThe following conclusions have been obtained from a preliminary theoretical study and from an experimental investigation conducted with airfoils of approximately 10-percent-thickness ratio at Reynolds numbers between 0.2 and 1.2 million, and at Mach numbers of 1.5 and 2.0:\n\n1. At supersonic velocities a properly designed airfoil having a blunt trailing edge produces a lower drag and a greater lift-curve slope than a conventional sharp-trailing-edge airfoil.\n\n2. Further theoretical and experimental study of blunt-trailing-edge airfoils is needed before it is possible to specify the airfoil shape that is nearly optimum for a given structural requirement.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:56:11.995308+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 27, "total_pages": 114, "image_filename": "19930086061_p27.jpg", "text": "NACA RM L9J07\n23\n\n11. Jones, Robert T.: Subsonic Flow over Thin Oblique Airfoils at Zero Lift. NACA Rep. 902, 1948.\n\n12. Jaquet, Byron M., and Brewer, Jack D.: Effects of Various Outboard and Central Fins on Low-Speed Static-Stability and Rolling Characteristics of a Triangular-Wing Model. NACA RM L9E18, 1949.\n\n13. DeYoung, John: Theoretical Additional Span Loading Characteristics of Wings with Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA TN 1491, 1947.", "timestamp": "2026-07-22T05:56:12.414470+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 31, "total_pages": 50, "image_filename": "19930086076_p31.jpg", "text": "NACA RM E9F09\n29\n\n<!-- Image (345, 124, 719, 736) -->\n\nFigure 10. - Schematic diagram of flame holder 9.", "timestamp": "2026-07-22T05:56:13.112583+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 33, "total_pages": 33, "image_filename": "19930085977_p33.jpg", "text": "```markdown\nWing alone\nWing-fuselage\n\nCONFIDENTIAL\n\n32\n\n$(C_D)_{L=0}$\n.08\n.04\n0\n\n$(\\frac{\\partial C_L}{\\partial \\alpha})_M$\n.12\n$C_L = 0$\n.08\n.04\n\n$(\\frac{\\partial C_m}{\\partial C_L})_M$\n.2\n$C_L = 0$\n0\n-.2\n\n$y_{cp}$\n60\n$C_L = 0.4$\n.40\n\n.6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\n\n$(\\frac{\\partial \\epsilon}{\\partial \\alpha})_M$\n.8\n$h_t$\n.4\n$C_L = 0$\n0\n0\n-30\n-30\n\n$(\\frac{\\partial \\epsilon}{\\partial \\alpha})_M$\n.8\n$h_t$\n.4\n$C_L = 0$\n0\n0\n$\\pm 30$\n\n$\\frac{q_{wake}}{q}$\n1.2\n$h_t$\n.8\n$C_L = 0$\n.4\n.6 .7 .8 .9 1.0 1.1 1.2\nMach number, M\n+30\n0\nNACA\n\nCONFIDENTIAL\n\nNACA-Langley - 10-21-49 - 400\n\nFigure 15.-- Summary of aerodynamic characteristics for a model with $0^\\circ$ sweptback wing, aspect ratio 4,\ntaper ratio 0.6, and NACA 65A006 airfoil section.\n\nNACA RM L9E22\n```", "timestamp": "2026-07-22T05:56:15.650871+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 45, "total_pages": 51, "image_filename": "19930085962_p45.jpg", "text": "44\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (125, 197, 875, 792) -->\n\nFigure 13. — Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:56:18.432092+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 20, "total_pages": 34, "image_filename": "19930082472_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:56:19.498675+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 64, "total_pages": 96, "image_filename": "19930085880_p64.jpg", "text": "62\nNACA RM No. L9C03\n\n<!-- Image (146, 198, 874, 878) -->\n\nWetted area, sq ft\n(c) $\\tau = 12^\\circ$.\nFigure 19.- Continued.\nNACA", "timestamp": "2026-07-22T05:56:20.548751+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 4, "total_pages": 65, "image_filename": "19930082546_p4.jpg", "text": "NACA TN No. 1870\n3\n\n$S_o$ distance from observer to doublets at effective propeller radius\n\n$x,y,z$ Cartesian system of coordinates, propeller axis along x-axis\n\n$x',y',z'$ axes with origin at doublet and parallel to x-, y-, and z-axes\n\nd tip clearance\n\nD propeller diameter\n\nr station radius\n\nb blade width\n\nh maximum thickness of blade section\n\nB number of blades\n\n$\\rho$ density of air\n\nc speed of sound\n\n$k = \\frac{mBv}{c}$\n\n$M_t$ tip Mach number (rotation only)\n\nR tip radius of propeller\n\nQ torque\n\nT thrust\n\nP power\n\n$p_i$ instantaneous pressure for a given harmonic $\\left(p_i = \\rho \\frac{\\partial \\phi}{\\partial t}\\right)$\n\np free-space oscillating pressure for a given harmonic, root mean square\n\n$\\bar{p}$ total free-space oscillating pressure, root mean square\n$$ \\left( \\bar{p} = \\sqrt{\\sum_{m=1}^{m=\\infty} p_{mB}^2} \\right) $$\n\n$p_{mB}$ p for any mB value", "timestamp": "2026-07-22T05:56:22.644792+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 39, "total_pages": 42, "image_filename": "19930085938_p39.jpg", "text": "```markdown\n38\nNACA RM No. L9B04\n\nc.g. location\nper cent M.A.C.\no 20\n□ 30\n◇ 40\n\n<!-- Image (130, 111, 781, 417) -->\n\n(a) Amplitude of vertical motion.\n\n<!-- Image (155, 494, 781, 858) -->\n\n(b) Amplitude of trim oscillation.\n\nFigure 14.- Landing stability of single-boom model.\n```", "timestamp": "2026-07-22T05:56:23.088781+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 1, "total_pages": 44, "image_filename": "19930082566_p1.jpg", "text": "629.1309\n1639\nY3.N21/5:6/1889\nGOVT. DOC.\nNACA TN No. 1889\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\nNo. 1889\n\nBIAXIAL FATIGUE STRENGTH OF 24S-T ALUMINUM ALLOY\nBy Joseph Marin and William Shelson\nThe Pennsylvania State College\n\n[Figure: NACA logo]\n\nWashington\nMay 1949\n[annotation: COMMERCE LIBRARY]\nMAY 5 1949\n\nBUSINESS, SCIENCE\n& TECHNOLOGY DEPT.", "timestamp": "2026-07-22T05:56:25.960531+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 19, "total_pages": 21, "image_filename": "19930082474_p19.jpg", "text": "NACA TN No. 1799\n17\n\nStick position,\nin. forward\n0\n4\nAngle of pitch, deg\n10\n-10\nNormal\nacceleration, g\n1.2\n.8\n0\n10\n20\n30\nTime, sec\nNACA\n\nFigure 2.- Longitudinal oscillation at 40 miles per hour.\n\nStick position,\nin. forward\n0\n5\nNormal\nacceleration, g\n1.5\n1.0\n.5\n0\n5\n10\n15\n20\nTime, sec\nNACA\n\nFigure 3.- Longitudinal oscillation at 65 miles per hour.", "timestamp": "2026-07-22T05:56:26.785488+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 14, "total_pages": 21, "image_filename": "19930091993_p14.jpg", "text": "```markdown\n10\nREPORT 928—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nIf the turbine air flow is accurately known as a function of the pressure ratio (or equivalent isentropic enthalpy drop), it is clear that on this figure the turbine curves will shift vertically. Because of the large slope of these curves, the intersections with the compressor curves will be very little affected by such an efficiency error and no effect would be observed for choking of the turbine, inasmuch as this condition results in vertical turbine curves. Turbine pressure ratio and therefore exhaust pressure ratio would be affected by such an error. Experimental results of the turbine performance obtained with the cold-air investigation of the turbine component show excellent correlation of air-flow characteristics and slight variations of efficiency. Hence, the results of cold-air turbine-component experiments may be used to predict accurately compressor operation in the engine but slight errors in jet-power estimates would result from the use of such data.\n\nThe lines of constant corrected engine temperature ratio obtained from equation (4) are shown in figure 14. At low speeds, the highest temperature ratio that may be used without surging the compressor is 3.0. The ratio of 4.0 has only one useable compressor speed of those plotted, whereas the ratio of 4.5 is entirely in the surge region for all compressor speeds.\n\nratio. Surge limitation for temperature ratios of 3.5 and 4.0 are clearly shown.\n\n<!-- Image (503, 111, 942, 327) -->\nFIGURE 15.—Compressor operating conditions at various temperature ratios in engine. Compressor-inlet stagnation temperature and pressure, 480° R and 1414 pounds per square foot, respectively.\n\nThe operation of the engine at air flows lower than for peak compressor pressure ratio in what is normally the surge region was not accompanied by the usual pressure fluctuations. In varying the operation of the engine from this surge region to the normal operating range, the gas-flow parameter used in the matching chart changes by the usual small increments. This parameter, however, uses the combustion-chamber pressure to correct the weight flow. The discontinuous jump in pressure ratio therefore corresponds to a jump in the weight flow correlated on the basis of the inlet pressure. For fixed engine-inlet conditions, the actual (uncorrected) air flow through the engine will also change discontinuously. This phenomenon is illustrated in figure 16, which shows actual data for compressor speeds of 214 and 245 rps. The discontinuous increase in air flow at 214 rps is 40 percent whereas at 245 rps the increase is 50 percent. In both cases the combustion-chamber volume flow changed very little. This jump will be reflected in engine thrust when passing through the surge region.\n\n<!-- Image (318, 407, 938, 660) -->\nFIGURE 14.—Turbine operating conditions at various temperature ratios in engine. Compressor-inlet stagnation temperature and pressure, 480° R and 1414 pounds per square foot, respectively.\n\nOperation of the compressor under these various conditions is shown more clearly in figure 15 where the pressure ratio is plotted against the air-flow parameter $W_1 n / (\\sigma_1 \\sqrt{\\theta_1})$ used in matching the compressor and turbine characteristics. Lines of constant compressor speed and compressor efficiency are also shown. Engine operating conditions beyond surge are not shown although the engine was operated in this condition as shown by the extension of several of the compressor speed lines beyond the region of peak pressure\n\n<!-- Image (503, 613, 942, 856) -->\nFIGURE 16.—Variation of compressor pressure ratio with inlet flow.\n```", "timestamp": "2026-07-22T05:56:31.410158+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 25, "total_pages": 28, "image_filename": "19930086092_p25.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 23\n\nSide-force coefficient, $C_Y$\nYawing-moment coefficient, $C_n$\nRolling-moment coefficient, $C_l$\nAngle of sideslip, $\\beta$, deg\n\n| Angle of attack, $\\alpha$ | |\n| :--- | :--- |\n| $\\circ$ | $0^\\circ$ |\n| $\\square$ | $6^\\circ$ |\n| $\\diamond$ | $12^\\circ$ |\n| $\\triangle$ | $21^\\circ$ |\n\n(b) $C_Y, C_n, C_l$ vs $\\beta$.\nFigure 5. — Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:56:38.705341+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 18, "total_pages": 29, "image_filename": "19930093789_p18.jpg", "text": "NACA RM No. E8121\nCONFIDENTIAL\n17\n\n1031\n\n[Figure: Diagram showing rotor-blade profile (mean section).]\n\nNACA\n\nFigure 4. - Diagram showing rotor-blade profile (mean section).\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:56:44.100003+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 72, "total_pages": 78, "image_filename": "19930082618_p72.jpg", "text": "```markdown\n70\nNACA TN 1945\n\n<!-- Image (143, 110, 806, 865) -->\n\n(a) Airfoils with smooth surfaces.\nFigure 18.- Variation of section angle of zero lift with Reynolds number\nfor the plain airfoils.\n```", "timestamp": "2026-07-22T05:56:45.068328+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 1, "total_pages": 50, "image_filename": "19930082592_p1.jpg", "text": "NATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE 1914\n\nOXIDATION OF TITANIUM CARBIDE BASE CERMALS\nCONTAINING MOLYBDENUM, TUNGSTEN, AND COBALT\n\nBy M. J. Whitman and A. J. Repko\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\n[Figure: NACA logo]\n\nWashington\nJuly 1949", "timestamp": "2026-07-22T05:56:47.407915+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 32, "total_pages": 98, "image_filename": "19930086073_p32.jpg", "text": "```markdown\n30\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| $\\triangle$ | $\\diamond$ | $\\circ$ | $\\square$ |\n| 44.0 | 21.2 | 0 | -22.0 |\n| Flap deflection, $\\delta_f$, deg | | | |\n\n(b) $C_L$ vs $C_D$.\n\nFigure 5.- Continued.\n\n[Figure: NACA logo]\n\nNACA RM A9H04\n```", "timestamp": "2026-07-22T05:56:48.394055+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 2, "total_pages": 44, "image_filename": "19930082566_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:56:49.832155+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 46, "total_pages": 51, "image_filename": "19930085962_p46.jpg", "text": "NACA RM A9E05\nCONFIDENTIAL\n\n.8\nLift coefficient, $C_L$\n.6\n.4\n.2\n0\n-.2\n\n$\\delta_e$\n$30^\\circ$\n$20^\\circ$\n$10^\\circ$\n$6^\\circ$\n$4^\\circ$\n$2^\\circ$\n$0^\\circ$\n\n0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0\nMach number, M\n(a) $\\alpha_{approx.}$, $0^\\circ$.\n\n[Figure: Graph showing the variation of lift coefficient with Mach number for various elevator deflections. The graph contains multiple data series represented by different symbols (triangles, diamonds, squares, circles) corresponding to different elevator deflection angles ($\\delta_e$). The NACA logo is present in the lower right corner of the plot area.]\n\nFigure 14.—The variation of lift coefficient with Mach number for various elevator deflections.\n\nCONFIDENTIAL\n45", "timestamp": "2026-07-22T05:56:50.407277+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 5, "total_pages": 65, "image_filename": "19930082546_p5.jpg", "text": "4\nNACA TN No. 1870\n\n$P_s$\npressure at panel surface\n\n$\\omega$\nrotational speed, radians per second\n\n$\\omega_n$\nundamped natural angular frequency of vibration of panel,\nradians per second\n\n$\\omega_1$\nangular frequency of sound or vibration, radians per second\n\n$t$\ntime, seconds\n\n$n$\npropeller rotational speed, revolutions per second\n\n$C_T$\nthrust coefficient $\\left(\\frac{T}{\\rho n^2 D^4}\\right)$\n\n$W = (C + 2K) + 1 \\left(M_{D1} - \\frac{s}{\\omega_1}\\right)$\n\n$C_Q$\ntorque coefficient $\\left(\\frac{Q}{\\rho n^2 D^5}\\right)$\n\n$C_P$\npower coefficient $\\left(\\frac{P}{\\rho n^3 D^5}\\right)$\n\n$C_{\\bar{p}}$\ntotal free-space oscillating pressure coefficient $\\left(\\frac{\\bar{p}}{\\rho n^2 D^2}\\right)$\n\n$C_p$\nfree-space oscillating pressure coefficient $\\left(\\frac{p}{\\rho n^2 D^2}\\right)$\n\n$m$\norder of the harmonic\n\n$A(r) = \\frac{1}{B} \\frac{dt}{dr}$\n\n$F(r) = \\frac{1}{Br} \\frac{dQ}{dr}$\n\n$\\epsilon_m$\nphase angle between Fourier harmonic of impulse and torque\ncomponent of impulse\n\n$\\eta_m$\nphase angle between Fourier harmonic of impulse and thrust\ncomponent of impulse\n\n$\\beta$\nblade angle, degrees\n\n$\\delta$\nangle of doublet from observer with respect to $x'$ axis\n\n$\\chi$\nangle of doublet from observer with respect to $y'$ axis\n\n$\\nu$\nangle of doublet from observer with respect to $z'$ axis", "timestamp": "2026-07-22T05:56:50.582345+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 16, "total_pages": 39, "image_filename": "19930093769_p16.jpg", "text": "```markdown\nNACA RM No. E8L08a\nCONFIDENTIAL\n\n| | | | | | | | | | | | | |\n|---|---|---|---|---|---|---|---|---|---|---|---|---|\n| 30,300 | 0.7148 | 11,504 | 13,120 | 899 | 3142 | 1.175 | 1.340 | 1282 | 1666 | 0.293 | 1.169 | 48-206 |\n| 30,200 | .7130 | 11,920 | 13,563 | 1024 | 3538 | 1.151 | 1.319 | 1340 | 1761 | .328 | 1.297 | 48-206 |\n| 30,250 | .7092 | 12,480 | 14,342 | 1124 | 3839 | 1.167 | 1.341 | 1499 | 1970 | .364 | 1.430 | 48-206 |\n| 30,050 | .8520 | 10,536 | 11,854 | 414 | 1404 | 2.027 | 2.280 | 1086 | 1375 | .233 | .890 | 48-206 |\n| 30,200 | .8560 | 11,505 | 12,978 | 845 | 2900 | 1.314 | 1.482 | 1254 | 1595 | .308 | 1.193 | 48-206 |\n| 30,200 | .8560 | 11,970 | 13,484 | 953 | 3270 | 1.336 | 1.505 | 1350 | 1713 | .354 | 1.367 | 48-206 |\n| 30,200 | .8560 | 12,420 | 13,991 | 1027 | 3524 | 1.394 | 1.559 | 1427 | 1811 | .395 | 1.526 | 48-206 |\n| 45,000 | .2227 | 10,536 | 11,928 | 617 | 4322 | 1.328 | 1.503 | 1418 | 1817 | .227 | 1.805 | 48-206 |\n| 45,000 | .2227 | 11,540 | 13,066 | 663 | 4649 | 1.338 | 1.516 | 1650 | 2120 | .247 | 1.959 | 48-206 |\n| 45,000 | .2227 | 11,720 | 13,300 | 640 | 4484 | 1.388 | 1.575 | 1697 | 2185 | .247 | 1.962 | 48-206 |\n| 45,250 | .5022 | 10,404 | 12,049 | 479 | 3364 | 1.770 | 2.050 | 1256 | 1684 | .236 | 1.915 | 48-206 |\n| 45,250 | .6022 | 11,376 | 13,191 | 687 | 4824 | 1.322 | 1.533 | 1428 | 1920 | .252 | 2.054 | 48-206 |\n| 44,650 | .5595 | 12,060 | 13,984 | 703 | 4812 | 1.388 | 1.609 | 1639 | 2204 | .271 | 2.152 | 48-206 |\n| 43,950 | .5895 | 12,280 | 14,258 | 741 | 4960 | 1.325 | 1.538 | 1713 | 2309 | .273 | 2.120 | 48-206 |\n| 45,200 | .8603 | 10,460 | 11,890 | 252 | 1778 | 2.349 | 2.693 | 1145 | 1504 | .164 | 1.330 | 48-206 |\n| 45,200 | .8603 | 11,424 | 13,095 | 531 | 3747 | 1.273 | 1.459 | 1346 | 1767 | .188 | 1.519 | 48-206 |\n| 45,200 | .8639 | 11,972 | 13,740 | 640 | 4516 | 1.217 | 1.397 | 1457 | 1919 | .216 | 1.752 | 48-206 |\n| 44,500 | .8513 | 12,436 | 14,237 | 635 | 4367 | 1.313 | 1.504 | 1589 | 2083 | .232 | 1.824 | 48-206 |\n| 50,500 | .2314 | 10,504 | 11,920 | 817 | 5460 | .959 | 1.089 | 1506 | 1939 | .164 | 1.651 | 48-210 |\n| 49,800 | .2145 | 11,258 | 12,775 | 585 | 4900 | ----- | ----- | 1720 | 2215 | ----- | 1.723 | 48-210 |\n| 50,120 | .7338 | 10,452 | 12,312 | 469 | 4128 | 1.275 | 1.502 | 1251 | 1736 | .166 | 1.723 | 48-210 |\n| 50,120 | .7406 | 11,560 | 13,527 | 637 | 5607 | 1.162 | 1.360 | 1475 | 2020 | .206 | 2.116 | 48-210 |\n| 50,120 | .7359 | 11,968 | 14,004 | 635 | 5590 | 1.240 | 1.451 | 1636 | 2240 | .219 | 2.252 | 48-210 |\n| 50,100 | .8357 | 10,498 | 12,068 | 480 | 4210 | 1.278 | 1.470 | 1225 | 1622 | .170 | 1.719 | 48-210 |\n| 50,740 | .8424 | 11,540 | 13,313 | 615 | 5588 | 1.213 | 1.399 | 1469 | 1954 | .208 | 2.171 | 48-210 |\n| 49,500 | .8269 | 12,012 | 13,821 | 657 | 5601 | 1.231 | 1.416 | 1571 | 2080 | .225 | 2.205 | 48-210 |\n| 50,120 | .8359 | 12,310 | 14,182 | 675 | 5942 | 1.245 | 1.434 | 1676 | 2225 | .234 | 2.368 | 48-210 |\n\naMaximum rotor speed limited by capacity of exhauster system.\n\n[Figure: NACA logo]\n\nCONFIDENTIAL\n15\n```", "timestamp": "2026-07-22T05:56:54.283470+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 20, "total_pages": 21, "image_filename": "19930082474_p20.jpg", "text": "18\nNACA TN No. 1799\n\n$\\Delta g$/Unit time\n\n[Figure: A graph plotting $\\Delta g$/Unit time against V, mph. The x-axis ranges from 0 to 90. The curve starts near zero, rises slightly, dips around 45 mph, then rises sharply after 60 mph. A NACA logo is present near the x-axis label.]\n\nV, mph\n\nFigure 4.- Rate of deviation from steady trimmed flight.", "timestamp": "2026-07-22T05:56:54.622787+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 32, "total_pages": 50, "image_filename": "19930086076_p32.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:56:55.530989+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 15, "total_pages": 78, "image_filename": "19930082483_p15.jpg", "text": "NACA TN No. 1807\n13\n\nThis equation is used to evaluate the driving-fluid losses due to partial admission. The driving-fluid losses represent the amount by which the blade power output at partial admission fails to equal the fraction of full-admission power represented by the admission ratio.\n\nA more general expression that defines the driving-fluid losses may be developed. From equations (24) and (26),\n\n$$\n\\begin{aligned}\n(\\text{driving-fluid losses})_F &= FK_{III}\\bar{\\rho}_i N \\left\\{ \\sum_{B=0}^{B=b} \\sum_{r_h}^{r_T} (c_{x,1})_{360^\\circ} \\left[ (c_{u,1})_{360^\\circ} - (c_{u,2})_{360^\\circ} \\right] \\right\\} \\\\\n&- K_{III}\\bar{\\rho}_i N \\left\\{ \\sum_{B=0}^{B=bF} \\sum_{r_h}^{r_T} (c_{x,1})_F \\left[ (c_{u,1})_F - (c_{u,2})_F \\right] \\right\\}\n\\end{aligned}\n\\tag{27}\n$$\n\nwhere the number of active blades at any instant B equals Fb and $(c_{x,1})_F$, $(c_{u,1})_F$, and $(c_{u,2})_F$ are the gas velocities that are present at that degree of admission.\n\nEquation (27) expresses the differential between the power output with nozzle-arc reduction as would be indicated on the basis of reduced weight flow and the actual power output as would be observed for that same amount of nozzle-arc reduction. For fixed conditions of gas state and rotor speed, good correlation is obtained if this power difference is considered to increase linearly with the amount of active nozzle reduction over the range of practical reductions.\n\n$$\n\\frac{(\\text{driving-fluid losses})_F}{(\\text{driving-fluid losses})_G} = \\frac{1-F}{1-G}\n\\tag{28}\n$$\n\nwhere F and G are the fractions of active nozzle arc for any particular degrees of admission.\n\nThe net observed dynamometer power at partial admission must now be written", "timestamp": "2026-07-22T05:56:57.340502+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 65, "total_pages": 96, "image_filename": "19930085880_p65.jpg", "text": "NACA RM No. L9C03\n63\n\n[Figure: Graph plotting Resistance, lb (y-axis) against Wetted area, sq ft (x-axis). The y-axis ranges from 0 to 7. The x-axis ranges from 0 to .35. There are five curves representing different speeds (fps): 10, 15, 20, 25, and 30. The curves show increasing resistance with increasing wetted area and speed. Data points are marked with various symbols (circles, squares, diamonds, triangles).]\n\nWetted area, sq ft\n(d) $\\tau = 16^\\circ$.\nFigure 19.- Continued.\nNACA", "timestamp": "2026-07-22T05:57:01.348467+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 26, "total_pages": 28, "image_filename": "19930086092_p26.jpg", "text": "24\nCONFIDENTIAL\nNACA RM A9F14\n\n<!-- Image (159, 103, 813, 788) -->\n\nFigure 6.—Effect of the vertical tail on the values of the directional-stability parameter, $C_{n\\beta}$, and the effective-dihedral parameter, $C_{l\\beta}$, of the 63° swept-back wing with fuselage. $\\beta=0$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:57:11.065368+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 40, "total_pages": 42, "image_filename": "19930085938_p40.jpg", "text": "```markdown\nNACA RM No. L9B04\n39\n\nGross load for stability\nand resistance tests\n\nGross load coefficient, $C_{\\Delta_0}$\n\n[Figure: Graph showing two curves labeled \"Twin boom\" and \"Single boom\". Arrows point to the curves indicating \"Intermittent spray in propellers\".]\n\nSpeed coefficient, $C_V$\n\nNACA\n\nFigure 15.- Gross load coefficient at which spray enters propellers.\n```", "timestamp": "2026-07-22T05:57:11.269838+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 90, "total_pages": 99, "image_filename": "19930082511_p90.jpg", "text": "```markdown\n88\n\n$$ \\frac{V}{\\Gamma} = \\frac{\\epsilon h V}{\\Gamma} = \\frac{2 \\epsilon h}{c_0^2} $$\n\n[Figure: A graph plotting $\\frac{V}{\\Gamma}$ on the y-axis against \"Distance from entrance, $\\xi$, units of tunnel height\" on the x-axis. The y-axis ranges from -0.2 to 0.8. The x-axis ranges from -0.8 to 2.0. The graph contains multiple curves labeled with values: -3, -5, -7, 0, 1.0, 1.3, 1.5, 2.0, -5. An inset diagram in the top left shows a schematic of a tunnel with labels $h$, $v$, $\\Gamma$, and $\\xi h$. The NACA logo is in the bottom right corner of the plot area.]\n\nDistance from entrance, $\\xi$, units of tunnel height\n\nFigure 18.- Tunnel-induced angle on axis of symmetrical two-dimensional closed-open-closed tunnel, with vortex at several locations along axis. Length of open section is 1.5 times tunnel height.\n\nNACA TN No. 1826\n```", "timestamp": "2026-07-22T05:57:14.050375+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 24, "total_pages": 36, "image_filename": "19930086097_p24.jpg", "text": "22\nCONFIDENTIAL\nNACA RM A9H11\n\nREFERENCES\n\n1. Busemann, A., and Walchner, O.: Aerofoil Characteristics at Supersonic Speeds. British R.T.P. Translation 1786. (Forschung. Vol. 4, no. 2, Mar./Apr. 1933, pp. 87-92.)\n\n2. Eggers, A. J., Jr.: Aerodynamic Characteristics at Subcritical and Supercritical Mach Numbers of Two Airfoil Sections Having Sharp Leading Edges and Extreme Rearward Positions of Maximum Thickness. NACA RM A7C10, 1947.\n\n3. Chapman, Dean R.: Base Pressure at Supersonic Velocities. Thesis submitted to California Institute of Technology, June, 1948.\n\n4. Ivey, H. Reese: Notes on the Theoretical Characteristics of Two-Dimensional Supersonic Airfoils. NACA TN 1179, 1947.\n\n5. Valensi, J., and Pruden, F. W.: Some Observations on Sharp Nosed Profiles at Supersonic Speed. A.R.C., Fluid Motion Sub-Committee, 10607 (FM 1108), May 1947.\n\n6. Sawyer, Richard H., and Daum, Fred L.: Measurement Through the Speed of Sound of Static Pressures on the Rear of Unswept and Sweptback Circular Cylinders and on the Rear and Sides of a Wedge by the NACA Wing-Flow Method. NACA RM L8B13, 1948.\n\n7. Vincenti, Walter G., Van Dyke, Milton D., and Matteson, Frederick H.: Investigation of Wing Characteristics at a Mach Number of 1.53. II - Swept Wings of Taper Ratio 0.5. NACA RM A8E05, 1948.\n\n8. Chapman, Dean R., and Perkins, Edward W.: Experimental Investigation of the Effects of Viscosity on the Drag of Bodies of Revolution at a Mach number of 1.5. NACA RM A7A31a, 1947.\n\n9. Van Dyke, Milton D.: Aerodynamic Characteristics Including Scale Effect of Several Wings and Bodies Alone and in Combination at a Mach Number of 1.53. NACA RM A6K22, 1946.\n\n10. Vincenti, Walter G., Nielsen, Jack N., and Matteson, Frederick H.: Investigation of Wing Characteristics at a Mach Number of 1.53. I - Triangular Wings of Aspect Ratio 2. NACA RM A7I10, 1947.\n\n11. Ferri, Antonio: Experimental Results with Airfoils Tested in the High-Speed Tunnel at Guidonia. NACA TM 946, 1940.\n\n12. Puckett, A. E.: Optimum Shapes for Thin Two-Dimensional Airfoil Sections. Papers Presented at the Symposium on Aerodynamics held Dec. 6 and 7, 1945. John Hopkins Univ., Applied Physics Lab., Silver Spring, Md., pp. 46-51.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:57:15.127230+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 17, "total_pages": 39, "image_filename": "19930093769_p17.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:57:17.568489+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 21, "total_pages": 21, "image_filename": "19930082474_p21.jpg", "text": "NACA TN No. 1799\n19\n\nRotor A\nStick position\nRotor B\n\nStick force\n\n0 1 2\nTime, sec\n\n0 1 2\nTime, sec\nNACA\n\nFigure 5.- Stick forces following abrupt lateral stick deflection.\n\nRotor A\nStick position\nRotor B\n\nStick force\n\n0 1 2\nTime, sec\n\n0 1 2\nTime, sec\nNACA\n\nFigure 6.- Stick forces following abrupt longitudinal stick deflection.", "timestamp": "2026-07-22T05:57:19.435335+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 6, "total_pages": 65, "image_filename": "19930082546_p6.jpg", "text": "NACA TN No. 1870\n\n$\\phi$ velocity potential\n\n$\\theta$ angle between y-axis and radius of doublet circle\n\n$\\xi_{01}$ amplitude of impinging free wave\n\n$\\dot{\\xi}_{01}$ velocity of impinging free wave\n\n$\\xi_{02}$ amplitude of panel vibration\n\n$\\dot{\\xi}_{02}$ velocity of panel vibration\n\nC structural damping of wall\n\n$C_c$ critical structural damping ($2M\\omega_n$)\n\nK acoustical radiation resistance ($\\rho c$)\n\nM mass of panel per unit area\n\ns effective stiffness of panel per unit area ($M\\omega_n^2$)\n\n$T_c$ transmission coefficient $(\\xi_{02}/\\xi_{01})^2$\n\n$A_c$ absorption coefficient\n\n$f_1$ frequency of sound or vibration, cycles per second\n\n$f_o$ natural frequency of panel, cycles per second\n\nA dot over a quantity indicates the first derivative with respect to time of that quantity.\n\nTHEORY\n\nThe theory for the generation of sound by a propeller is given by Gutin in reference 1. His basic assumptions are that the propeller is replaced by concentrated forces or acoustic doublets distributed over the propeller disk, the strength of the doublets being a function of the torque and thrust of the propeller. By considering only the sound at a great distance from the propeller, Gutin could make further simplifying assumptions which permitted a solution in terms of Bessel functions. In the present analysis, which considers the oscillating pressures near the propeller tips, the assumptions of great distance cannot be made. The analysis therefore follows closely that of Gutin, with the exception that no simplifying assumption as to distance is made.", "timestamp": "2026-07-22T05:57:21.223348+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 33, "total_pages": 50, "image_filename": "19930086076_p33.jpg", "text": "NACA RM E9F09\n31\n\n[Figure: Cutaway view of a metallic flame holder assembly showing internal fuel-discharge bars.]\n\nNACA\nC-21885\n7-9-48\n\nFigure 11. - Cutaway view of flame holder 9 showing position of fuel-discharge bars.", "timestamp": "2026-07-22T05:57:21.330585+00:00"}