PDF/ocr_dataset_1026.jsonl

{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 40, "total_pages": 62, "image_filename": "19930082485_p40.jpg", "text": "NACA TN NO. 1810\n\n78\nNACA\n74\nGas-discharge angle, $\\alpha$, deg\n70\n66\nInner shroud\nOuter shroud\n62\n-O- Experimental\n-- Design\n58\n9.0 9.8 10.6 11.4\nRadius, r, in.\n\nV\n$\\alpha$\n0.1 chord\n\n(a) Radial variation.\nFigure 9. - Variation of gas-discharge angle at 0.1 chord downstream of cascade.\n\n39", "timestamp": "2026-07-22T05:02:10.985210+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 30, "total_pages": 37, "image_filename": "19930082450_p30.jpg", "text": "NACA TN No. 1778\n29\n\n$$\n\\frac{H}{t_W} = 26\n$$\n$$\n\\left( \\frac{b_W}{t_W} = 25 \\right)\n$$\n\n$$\n\\bar{\\sigma}_f, ksi\n$$\n\n$$\n\\sigma_{cr}, ksi\n$$\n11.4\n7.3\n\n$$\n\\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S}\n$$\n\n$$\n\\frac{P_i}{L\\sqrt{E}}, ksi\n$$\n\n$$\n\\frac{P_i}{t_S}, ksi\n$$\n\n$$\n\\frac{H}{t_W} = 36\n$$\n(35)\n\n$$\n\\bar{\\sigma}_f, ksi\n$$\n\n$$\n\\sigma_{cr}, ksi\n$$\n11.4\n7.3\n\n$$\n\\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S}\n$$\n\n$$\n\\frac{P_i}{L\\sqrt{E}}, ksi\n$$\n\n$$\n\\frac{P_i}{t_S}, ksi\n$$\n\n$$\n\\frac{H}{t_W} = 46\n$$\n(45)\n\n$$\n\\bar{\\sigma}_f, ksi\n$$\n\n$$\n\\sigma_{cr}, ksi\n$$\n15.9\n11.4\n7.3\n\n$$\n\\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S}\n$$\n\n$$\n\\frac{P_i}{L\\sqrt{E}}, ksi\n$$\n\n24S-T\n$$\n\\sigma_{cy} = 44 \\text{ ksi}\n$$\n\n$$\n\\frac{P_i}{t_S}, ksi\n$$\n\nFigure 6.-Concluded. $\\frac{t_W}{t_S} = 0.51$.\n\nNACA", "timestamp": "2026-07-22T05:02:11.460065+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 22, "total_pages": 53, "image_filename": "19930082542_p22.jpg", "text": "NACA TN No. 1867\n21\n\nTABLE I.- ROOM-TEMPERATURE PHYSICAL PROPERTIES OF LOW-CARBON B-175 BAR STOCK - Continued\n\n<!-- Table (101, 115, 875, 890) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|}\n\\hline\n\\multicolumn{3}{|c|}{Heat treatment} & \\multicolumn{3}{|c|}{Aging treatment (a)} & \\multicolumn{2}{|c|}{Sub-cold-rolling (b)} & \\multirow{3}{*}{Brinell hardness} & \\multirow{3}{*}{Tensile strength (psi)} & \\multicolumn{3}{|c|}{Offset yield strength (psi)} & \\multirow{3}{*}{Proportional limit (psi)} & \\multirow{3}{*}{Elongation in 2 in. (percent)} & \\multirow{3}{*}{Reduction of area (percent)} \\\\\n\\cline{1-8}\\cline{11-13}\n\\multicolumn{2}{|c|}{Solution treatment} & \\multirow{2}{*}{Method of quench (c)} & \\multirow{2}{*}{Temperature ($^\\circ$F)} & \\multirow{2}{*}{Time (hr)} & \\multirow{2}{*}{Temperature ($^\\circ$F)} & \\multirow{2}{*}{Time (hr)} & \\multirow{2}{*}{Percent reduction} & & & \\multirow{2}{*}{0.02 percent} & \\multirow{2}{*}{0.1 percent} & \\multirow{2}{*}{0.2 percent} & & & \\\\\n\\cline{1-2}\nTemperature ($^\\circ$F) & Time (hr) & & & & & & & & & & & & & & \\\\\n\\hline\n\\multicolumn{17}{|c|}{Solution-treated at 2050$^\\circ$ F} \\\\\n\\hline\n2050 & 1 & W.Q. & --- & --- & --- & --- & --- & 152 & 113,770 & 44,500 & 54,500 & 58,500 & 39,500 & 70 & 63.9 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & --- & 158 & 113,800 & 44,500 & 56,500 & 56,500 & 39,500 & 70 & 65.7 \\\\\n2050 & 5 & A.C. & --- & --- & --- & --- & --- & 189 & 118,150 & 48,500 & 57,000 & 59,000 & 37,500 & 52 & 64.8 \\\\\n2050 & 5 & W.Q. & --- & --- & --- & --- & --- & & & & & & & & \\\\\n\\hline\n\\multicolumn{17}{|c|}{Aging time and temperature:} \\\\\n\\hline\n2050 & 2 & W.Q. & 1400 & 2 & --- & --- & --- & 190 & 119,790 & 49,500 & 57,500 & 61,000 & 37,000 & 42.5 & 51.0 \\\\\n2050 & 2 & W.Q. & 1400 & 16 & --- & --- & --- & 179 & --- & --- & --- & --- & --- & --- & --- \\\\\n2050 & 2 & W.Q. & 1400 & 24 & --- & --- & --- & 177 & 118,500 & 38,000 & 51,000 & 58,500 & 17,500 & 36 & 41.6 \\\\\n2050 & 2 & W.Q. & 1350 & 24 & --- & --- & --- & 220 & --- & --- & --- & --- & --- & --- & --- \\\\\n2050 & 2 & W.Q. & 1350 & 24 & --- & --- & --- & 190 & 119,625 & 48,500 & 56,400 & 60,000 & 37,500 & 42 & 50.0 \\\\\n2050 & 2 & W.Q. & 1600 & 24 & --- & --- & --- & 213 & 122,150 & 39,000 & 52,500 & 54,000 & 22,500 & 38 & 40.3 \\\\\n2050 & 2 & W.Q. & 1600 & 24 & --- & --- & --- & 203 & 119,150 & 39,000 & 49,500 & 50,000 & 22,500 & 38 & 39.9 \\\\\n2050 & 2 & W.Q. & 1750 & 24 & --- & --- & --- & 185 & 114,900 & 42,000 & 49,500 & 51,000 & 35,500 & 29.5 & 43.9 \\\\\n\\hline\n\\multicolumn{17}{|c|}{Cold-working amount and temperature:} \\\\\n\\hline\n2050 & 2 & W.Q. & --- & --- & --- & --- & 7 & 206 & 136,100 & 97,250 & 112,800 & 117,000 & 62,500 & 36 & 60.7 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & 1300 & 10 & 349 & 135,150 & 67,000 & 81,000 & 81,000 & 45,500 & 38.5 & 66.7 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & 1200 & 10 & 287 & 134,100 & 90,000 & 109,000 & 109,000 & 65,000 & 23.5 & 37.9 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & 1200 & 10 & 287 & 135,150 & 91,000 & 109,000 & 109,000 & 65,000 & 23.5 & 37.9 \\\\\n2050 & 2 & W.Q. & 1400 & 24 & --- & --- & 1200 & 10 & 268 & 138,250 & 94,500 & 106,500 & 111,000 & 75,000 & 28.5 & 37.5 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & 1200 & 20 & 340 & 134,100 & 91,000 & 114,500 & 131,000 & 65,000 & 22.5 & 49.5 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & 1200 & 25 & 318 & 136,000 & 104,000 & 130,000 & 137,500 & 70,000 & 22.5 & 48.1 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & 1400 & 15 & 260 & 125,000 & 104,000 & 110,500 & 110,500 & 77,500 & 22.5 & 45.7 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & 1400 & 15 & 260 & 125,000 & 104,000 & 110,500 & 110,500 & 77,500 & 22.5 & 45.7 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & 1700 & 15 & 235 & 128,475 & 81,000 & 95,000 & 94,700 & 57,500 & 38.5 & 48.1 \\\\\n2050 & 2 & W.Q. & --- & --- & --- & --- & 1800 & 15 & 220 & 129,100 & 78,000 & 87,400 & 91,000 & 57,500 & 38.5 & 46.8 \\\\\n\\hline\n\\multicolumn{17}{|c|}{Solution-treated at 2100$^\\circ$ F} \\\\\n\\hline\n2100 & 1 & W.Q. & --- & --- & --- & --- & --- & 197 & 117,250 & 38,500 & 52,000 & 57,000 & 17,500 & 47 & 63.2 \\\\\n\\hline\n\\end{tabular}\n\nAll aging treatments preceded by cold-rolling above where noted.\nW.Q., water-quenched; A.C., air-cooled.\n$^a$Aged after rolling.\n\nNACA", "timestamp": "2026-07-22T05:02:14.995032+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 24, "total_pages": 49, "image_filename": "19930082498_p24.jpg", "text": "NACA TN NO. 1838\n\n11,12\n\n13,14,15\n\nFigure 1.- Sketches illustrating some details of typical mufflers shown in table II.\n\n23", "timestamp": "2026-07-22T05:02:21.883450+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 3, "total_pages": 33, "image_filename": "19930085544_p3.jpg", "text": "2\nNACA RM No. L8K26\n\nrevolution. The airfoil blade section experiences oscillating air forces that vary with the position of the blade around the periphery. These air forces on the propeller blade section in flight must be related to the proper Mach number, advance ratio, blade-section lift coefficient, inclination of the propeller shaft axis to its forward motion, and the wave length of the oscillation. A knowledge of the air forces on the blade section as a function of the propeller operating conditions is needed in a study of the problem. No existing theory completely describes the operating condition of a pitched or yawed propeller.\n\nIn this report the air forces on the propeller blades are calculated first under the assumption that the existing propeller theory may be used in conjunction with the instantaneous angles of attack and resultant velocities along the blades of the pitched propeller at successive blade positions around the periphery. This method, herein termed the \"steady-state\" method, permits the use of the usual steady-state compressible airfoil characteristics with the Goldstein correction factors for a finite number of blades. Then several aspects of the nature of the forces developed by an oscillating airfoil are considered. Expressions based on linearized theory for calculating the air forces on a two-dimensional thin flat-plate airfoil oscillating in angle of attack in a steady stream in a nonviscous incompressible fluid were developed in reference 1. Some modifications to this theory were presented in reference 2 to permit calculations when the stream velocity as well as the angle of attack varied with time. The expressions of reference 2 are used to estimate the changes of the airfoil characteristics in a compressible oscillating flow field.\n\nThere are very little experimental data with which to compare the results of these calculations. The steady-state compressible characteristics are computed for the propeller tested in reference 3, however, and are compared with the experimental data given therein. The calculations are made for two-blade and three-blade single-rotating propellers and satisfactory agreement with the available experimental data is obtained.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| a | distance to center of rotation from midchord of airfoil, feet (fig. 2) |\n| B | number of blades |\n| c | chord, feet |\n| $c_d$ | profile drag coefficient |\n| $c_l$ | two-dimensional lift coefficient |\n| $C(k) = F + iG$ | C function (reference 1) |", "timestamp": "2026-07-22T05:02:25.485525+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 21, "total_pages": 36, "image_filename": "19930082614_p21.jpg", "text": "NACA TN 1939\n19\n\nbrakes is provided by calculations which show how the speed of an\nairplane in a specified maneuver is altered by the employment of the\nair brakes.\n\nEquations are presented in this report which permit a rapid cal-\nculation of the speed changes with time. Use of the equations results\nin close approximations to the values obtained by more accurate\nmethods. The equations are not general, however, and apply only to\nseveral specific problems. The speed during a maneuver can be accurate-\nly calculated as a function of time by a step-by-step procedure. The\ngraphs presented in this report substantially reduce the time required\nto make such calculations.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif., May 31, 1949.\n\nREFERENCES\n\n1. Hood, Manley J., and Allen, H. Julian: The Problem of Longitudinal\nStability and Control at High Speeds. NACA Rep. 767, 1943.\n\n2. Lowell, Arthur I: Fighter Airbrakes. Introductory Investigation of\nMeans for Improving the Tactical Effectiveness of Combat Aircraft\nby Making Provisions for Rapid Decelerations in Flight. ACTR\n4772 Army Air Corps, 1942.\n\n3. Purser, Paul E., and Turner, Thomas R.: Aerodynamic Characteristics\nand Flap Loads of Perforated Double Split Flaps on a Rectangular\nNACA 23012 Airfoil. NACA ARR, Jan. 1943.\n\n4. Purser, Paul E., and Turner, Thomas R.: Wind-Tunnel Investigation\nof Perforated Split Flaps for Use as Dive Brakes on a Tapered\nNACA 23012 Airfoil. NACA ARR, Nov. 1941.\n\n5. Knowler, A. E., and Pruden, F. W.: The Effect of Brake Flaps on\nan Aerofoil at High Speeds. R. & M. No. 2211, June 1942.\n\n6. Fuchs, D.: Wind-Tunnel Investigations of Diving Brakes. NACA TM\n1033, Nov. 1942.\n\n7. Laitone, Edmund V., and Summers, James L.: An Additional Investi-\ngation of the High-Speed Lateral-Control Characteristics of\nSpoilers. ACR 5D28, June 1945.", "timestamp": "2026-07-22T05:02:25.646083+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 22, "total_pages": 58, "image_filename": "19930082617_p22.jpg", "text": "```markdown\nNACA TN 1962\n21\n\nStringers\no 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X 10³\n2 108.0 X 10³\n3 180.0 X 10³\n4 252.0 X 10³\n\n2.57\"\nA\nBand V\nA-A\n45°\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n-20 -16 -12 -8 -4 0 4 8 12 16 20 X 10⁻⁴\nStrain\n\n1 2 3 4\n\n[Figure: Strain diagram of cylinder 74. Band V.]\n\nFigure 10.- Strain diagram of cylinder 74. Band V.\n```", "timestamp": "2026-07-22T05:02:25.847338+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 38, "total_pages": 41, "image_filename": "19930082476_p38.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:02:26.680242+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 15, "total_pages": 36, "image_filename": "19930085487_p15.jpg", "text": "NACA RM No. E8J22\n\n[Figure: Sand pattern for first torsional-mode vibration on first-stage blade. A ruler marked \"INCHES\" with scale from 0 to 2 is visible near the blade root. The NACA logo and document number C-14379 dated 2-27-46 appear in a box at lower right of image.]\n\nFigure 3. - Sand pattern for first torsional-mode vibration on first-stage blade. Frequency of vibration, 2820 cycles per second.\n\n13", "timestamp": "2026-07-22T05:02:32.309634+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 33, "total_pages": 99, "image_filename": "19930082511_p33.jpg", "text": "NACA TN No. 1326\n\n$$\n\\text{I.P.}w_2(a) = \\int_1^a \\sqrt{\\frac{a^2 - z^2}{1 - z^2}} \\, dz\n$$\n\nwhich by the substitution $z^2 = a^2 - (a^2 - 1)\\gamma^2$ reduces to\n\n$$\n\\text{ia} \\int_0^1 \\frac{\\frac{a^2 - 1}{a^2}\\gamma^2}{\\sqrt{(1 - \\gamma^2)\\left(1 - \\frac{a^2 - 1}{a^2}\\gamma^2\\right)}} \\, d\\gamma\n$$\n\n$$\n= \\text{ia} \\int_0^1 \\frac{d\\gamma}{\\sqrt{(1 - \\gamma^2)\\left(1 - \\frac{a^2 - 1}{a^2}\\gamma^2\\right)}} - \\text{ia} \\int_0^1 \\sqrt{\\frac{1 - \\frac{a^2 - 1}{a^2}\\gamma^2}{1 - \\gamma^2}} \\, d\\gamma\n$$\n\n$$\n= \\text{iaK}' - \\text{iaE}'\n$$\n\nSolution of simultaneous equations for M and N.- With the aid of the four formulas just derived, the two previously mentioned equations in M and N may be written\n\n$$\n\\frac{M}{a}K + NaE = -1\n$$\n\n$$\n\\frac{M}{a}K' + NaK' - NaE' = 0\n$$\n\nwhich are easily solved simultaneously for M and N. By introducing the following relation between the complete elliptic integrals (reference 11, p. 520)\n\n$$\nEK' - KK' + KE' = \\frac{\\pi}{2}\n$$\n\nthe expressions for M and N are finally obtained in the following forms:\n\n$$\nM = \\frac{2a}{\\pi}(K' - E')\n$$\n\n$$\nN = -\\frac{2K'}{a\\pi}\n$$", "timestamp": "2026-07-22T05:02:37.843299+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 13, "total_pages": 46, "image_filename": "19930085519_p13.jpg", "text": "12\nNACA RM No. L3K19\n\nComparison of plug ailerons and the half-span plain aileron.- A comparison of figures 17(b) and 22(a) ($\\delta_f = 0^\\circ$) and figures 19 ($\\delta_f = 30^\\circ$) and 22(b) ($\\delta_f = 50^\\circ$) indicates that the plug aileron has favorable yaw over the usable angle-of-attack range as compared to the adverse yaw present with the plain aileron.\n\nFor the flap-neutral condition, the plain aileron gave a maximum rolling-moment coefficient for a total aileron deflection of $40^\\circ$ approximately 130 percent greater than the maximum rolling-moment coefficient produced by the plug aileron at $\\delta_p = -0.07c$. For the flap-deflected condition (partial-span flap with the plain aileron and full-span flap with the plug aileron), the maximum value of rolling-moment coefficient produced by the plug aileron was about the same as that produced by $\\pm 20^\\circ$ deflection of the plain aileron.\n\nAt angles of attack above the wing-tip stall angle, the rolling-moment coefficients produced by the plain aileron were much larger than those produced by the plug aileron, regardless of the lift-flap condition.\n\nCONCLUSIONS\n\nThe results of an investigation of a $42^\\circ$ sweptback semispan-wing model equipped with several high-lift and lateral-control devices lead to the following conclusions:\n\n1. Of the various high-lift flaps investigated (full-span slotted flap at various positions and deflections, a half-span slotted flap at $50^\\circ$ deflection, a half-span split flap and a half-span Zap flap both at $60^\\circ$ deflection), the full-span slotted flap deflected to $30^\\circ$ gave the most satisfactory calculated landing characteristics for an airplane with an assumed wing loading of 40 pounds per square foot and a tail length of 3.0 mean aerodynamic chords.\n\n2. The plug-aileron arrangement investigated with the faired plug-slot lower lip gave positive rolling-moment coefficients at all projections throughout the wing angle-of-attack range, although there was a large reduction in rolling-moment coefficient at all projections at angles of attack above the wing-tip stall angle. The maximum values of rolling-moment coefficient produced by the plug aileron with the faired lower lip were about 130 percent larger with the full-span slotted flap deflected than with flap neutral.", "timestamp": "2026-07-22T05:02:38.657844+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 26, "total_pages": 50, "image_filename": "19930082496_p26.jpg", "text": "NACA TN No. 1836\n25\n\n[Figure: A black and white photograph of a complex piece of laboratory equipment, identified as a thermal-shock evaluation unit. The apparatus consists of various interconnected components including a furnace, quenching chamber, observation window, and multiple control units. Labels point to specific parts of the machinery.]\n\nFurnace\nQuenching chamber\nObservation window\nTimer\nFurnace temperature control\nQuenching air duct\nSpecimen - holder handle\nSpecimen temperature potentiometer\n\nNACA\nC-21416\n5-13-48\n\nFigure 3. - Thermal-shock evaluation unit.", "timestamp": "2026-07-22T05:02:41.942596+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 42, "total_pages": 66, "image_filename": "19930082245_p42.jpg", "text": "```markdown\nNACA TN NO. 1596\n\nPressure coefficient, P\n\n-1.2\n-.8\n-.4\n0\n.4\n.8\n1.2\nM = 0.252\n\n-1.2\n-.8\n-.4\n0\n.4\n.8\n1.2\nM = 0.550\n\n-1.2\n-.8\n-.4\n0\n.4\n.8\n1.2\nM = 0.676\n\n-1.2\n-.8\n-.4\n0\n.4\n.8\n1.2\nM = 0.700\n\n-1.2\n-.8\n-.4\n0\n.4\n.8\n1.2\nM = 0.726\n\n-1.2\n-.8\n-.4\n0\n.4\n.8\n1.2\nM = 0.755\n\n--- Upper surface\n--- Lower surface\n\nP_cr\n\nP_cr\n\nP_cr\n\nNACA\n\nx/c\n0   .2   .4   .6   .8   1.0\n\nx/c\n0   .2   .4   .6   .8   1.0\n\nx/c\n0   .2   .4   .6   .8   1.0\n\nFigure 8.- Pressure distribution about an NACA 66,1-115 airfoil section equipped with an unsealed 0.20c plain aileron of beveled-trailing-edge profile. δ_a = 0°; α = 1°.\n\n41\n```", "timestamp": "2026-07-22T05:02:43.764270+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 46, "total_pages": 47, "image_filename": "19930093773_p46.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:02:50.495630+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 11, "total_pages": 20, "image_filename": "19930085536_p11.jpg", "text": "NACA RM No. E53D5\n\nof pressure coefficient with angle of attack over a range of more than a few degrees. Similar results for angle of yaw are illustrated in figure 7, where linearized theory using the complete equation for pressure coefficient (equation (8)) again closely agrees with the experimental data.\n\nThe experimental results presented show excellent agreement with the linearized theory using the complete equation for the pressure coefficient (equation (8)). If, however, a similar procedure is used in comparing the linearized solution for a right circular cone with the exact values of reference 7, the results predicted by the linearized theory using the linearized pressure-coefficient relation (equation (9)) show better agreement with the results of Taylor and Maccoll (reference 7) than do those predicted by the complete relation. Because opposite results are obtained for the two cases, even though the same linearized theory is used for both, the excellent agreement between the experimental values for the elliptic cone and the values predicted by the linearized theory may be fortuitous.\n\nSUMMARY OF RESULTS\n\nThe following results were obtained from an investigation of the pressure distribution on a thin conical body of elliptical cross section at a Mach number of 1.89:\n\n1. At moderate angle of flow deflection, the experimental pressure distribution was in close agreement with the linearized theory using the complete equation for pressure coefficient. As the angle of flow deflection increased, the deviation from experiment of the theoretical pressure coefficient increased slightly although agreement was satisfactory over the entire range of calculations.\n\n2. Comparison of the complete equation for pressure coefficient with the equation usually used in connection with the linearized theory indicated that the terms omitted in obtaining the linearized equation were too large to be neglected. Inasmuch as the exact results of Taylor and Maccoll for a right circular cone show better agreement with the linearized theory when the linearized pressure-coefficient relation is used than when the complete relation is applied, whereas the opposite result was obtained in comparing the experimental results in this report with the linearized theory, the excellent agreement between the linearized theory and the experimental results presented may be fortuitous.\n\nLewis Flight Propulsion Laboratory,  \nNational Advisory Committee for Aeronautics,  \nCleveland, Ohio.", "timestamp": "2026-07-22T05:02:52.607152+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 16, "total_pages": 28, "image_filename": "19930085471_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:02:53.199309+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 25, "total_pages": 30, "image_filename": "19930082585_p25.jpg", "text": "24\nNACA TN 1907\n\nFlapping angle, $\\beta$, deg\nTime after power failure, sec\nNo pitch change\nSlow exponential\nModerate exponential\nInstantaneous\n\nFlapping angle, $\\beta$, deg\nTime after power failure, sec\n$I_1 = 100 \\text{ slug-ft}^2$\n$I_1 = 200 \\text{ slug-ft}^2$\n$I_1 = 400 \\text{ slug-ft}^2$\nNACA\n\nFigure 5.- Effects of rate of pitch reduction and blade moment of inertia on the variation of flapping angle with time after power failure.", "timestamp": "2026-07-22T05:02:54.546979+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 20, "total_pages": 66, "image_filename": "19930082914_p20.jpg", "text": "NACA TN No. 1857\n19\n\nbut\n\n$$\n\\frac{1}{V} = \\frac{n}{V_o}\n$$\n\ntherefore\n\n$$\n\\frac{dy}{dz} = \\sqrt{\\left(\\frac{c_1}{V_o} n\\right)^2 - 1}\n$$\n\nIt is now required to solve this equation in order to find the amount by which a ray of light is deviated in the mixing region. First, it is necessary to express $n$ as a function of $y$. In a subsequent section, the density distribution through the mixing region is obtained. For the present purpose this distribution is approximated by a linear variation that fits the actual variation over a large portion of the mixing region. The density of the air at the outside of the mixing zone is taken as 0.0024 slug per cubic foot and at the inside edge of the mixing region as 1.5 times as great, or 0.0036 slug per cubic foot. The effect of refraction is greatest at the place where the density gradient is the greatest. This occurs at the cross section that is closest to the nozzle. For the present investigation that cross section is 2 inches from the nozzle. The actual width of the mixing zone there is about 0.33 inch. The assumption is made of a linear density gradient equal to the average gradient across the mixing region. The assumed density variation, then, is given by the equation\n\n$$\n\\rho = 0.0024 + 0.0036y\n$$\n\nand the index variation is, by use of equation (1) and the given value for $k$,\n\n$$\nn = 1.00028 + 0.00042y\n$$\n\nor\n\n$$\nn = a + by\n$$", "timestamp": "2026-07-22T05:02:57.722670+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 41, "total_pages": 62, "image_filename": "19930082485_p41.jpg", "text": "40\nNACA TN No. 1810\n\nSuction surface\nPressure surface\n\n72\nDischarge angle, $\\alpha$, deg\n68\n64\n60\n0 .4 .8 1.2\nCircumferential distance, in.\n\nExperimental\nDesign\n\n(b) Circumferential variation at\nthe 10.35-in. radius.\n\nFigure 9.- Concluded. Variation of\ndischarge angle at 0.1 chord down-\nstream of cascade.", "timestamp": "2026-07-22T05:02:59.540721+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 25, "total_pages": 49, "image_filename": "19930082498_p25.jpg", "text": "24\n\n16\n\n18\n\nFigure 1.— Continued.\n\nNACA\n\nNACA TN No. 1858", "timestamp": "2026-07-22T05:02:59.728704+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 23, "total_pages": 58, "image_filename": "19930082617_p23.jpg", "text": "22.\nNACA TN 1962\n\nStringers\no 1 to 9\nx 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 x 10³\n2 72.0 x 10³\n3 108.0 x 10³\n4 144.0 x 10³\n5 216.0 x 10³\n6 288.0 x 10³\n\n2.57\"\nA\nBand B\nA\n45°\nA-A\n\nDistance from horizontal diameter, In.\nStrain\n\nFigure 11.- Strain diagram of cylinder 75. Band B.", "timestamp": "2026-07-22T05:03:12.225233+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 27, "total_pages": 50, "image_filename": "19930082496_p27.jpg", "text": "26\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:03:18.794879+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 4, "total_pages": 33, "image_filename": "19930085544_p4.jpg", "text": "NACA RM No. L8K26\n3\n\n$C_T$\nthrust coefficient $\\left( \\frac{T}{\\rho n^2 D^4} \\right)$\n\n$\\frac{dC_T}{dx}$\nelement thrust coefficient $\\left( \\frac{dT/dx}{\\rho n^2 D^4} \\right)$\n\nD\npropeller diameter, feet\n\nh\nvertical deflection (flapping) of airfoil, feet (fig. 2)\n\nJ\nadvance ratio (V/nD)\n\n$J_{\\omega t_0}$\nlocal advance ratio, steady part $\\left( J \\cos \\alpha_T \\right)$\n\n$J_{\\omega t}$\ninstantaneous local advance ratio $\\left( \\frac{\\pi x \\cos \\alpha_T}{\\frac{\\pi x}{J} + \\sin \\alpha_T \\sin \\omega t} \\right)$\n\n$k_{\\omega t}, k_{\\alpha_p}$\nparameter used in determining the function F + iG\n$\\left( k = \\frac{\\omega c}{2W_0} \\right)$\n\n$k_{\\omega t} + k_{\\alpha_p} = \\frac{\\omega c}{W_0} = k_1$\n\nL\nlift, pounds\n\n$L_c$\nlift coefficient of oscillating airfoil $\\left( \\frac{L}{2\\pi a p_0 \\frac{\\rho W_0^2}{2} c} \\right)$\n\nm\nturning moment, foot-pounds\n\n$m_c$\nmoment coefficient $\\left( \\frac{m}{\\rho V^2 D^3} \\right)$\n\nM\nMach number\n\nn\npropeller rotational speed, revolutions per second\n\nr\nradius to blade section, feet\n\nR\ntip radius\n\nt\ntime, seconds\n\nT\nthrust, pounds\n\nV\nforward velocity of airplane, feet per second", "timestamp": "2026-07-22T05:03:19.450323+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 6, "total_pages": 46, "image_filename": "19930085542_p6.jpg", "text": "4\nNACA RM No. L8L29\n\nq\ndynamic pressure, pounds per square foot $\\left(\\frac{\\rho V^2}{2}\\right)$\n\n$\\alpha$\nangle of attack in plane of symmetry, degrees\n\n$\\psi$\nangle of yaw, degrees\n\n$\\Lambda_{LE}$\nangle of sweepback of leading edge, degrees\n\n$\\Lambda_{c/4}$\nangle of sweepback of quarter-chord line, degrees\n$\\left(\\cot^{-1} \\frac{A}{3} \\text{ for triangular wings}\\right)$\n\n$\\frac{pb}{2V}$\nhelix angle generated by wing tip in roll, radians\n\np\nangular velocity in roll, radians per second\n\n$C_{l_\\alpha} = \\frac{\\partial C_l}{\\partial \\alpha}$\n\n$C_{l_\\psi} = \\frac{\\partial C_l}{\\partial \\psi}$\n\n$C_{n_\\psi} = \\frac{\\partial C_n}{\\partial \\psi}$\n\n$C_{Y_\\psi} = \\frac{\\partial C_Y}{\\partial \\psi}$\n\n$C_{l_p} = \\frac{\\partial C_l}{\\partial \\frac{pb}{2V}}$\n\n$C_{n_p} = \\frac{\\partial C_n}{\\partial \\frac{pb}{2V}}$\n\n$C_{Y_p} = \\frac{\\partial C_Y}{\\partial \\frac{pb}{2V}}$\n\nAPPARATUS, MODELS, AND TEST\n\nThe present investigation was conducted in the 6-foot-diameter rolling-flow test section of the Langley stability tunnel which is described in detail in reference 6.", "timestamp": "2026-07-22T05:03:22.552551+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 43, "total_pages": 66, "image_filename": "19930082245_p43.jpg", "text": "2\nSection angle of attack, $\\alpha$, deg\n$\\delta a$\n(deg)\n-2\n-4\n-6\n-8\n-10\n-12\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\n.12\nSection pitching-moment coefficient, $C_m$\n.08\n.04\n$\\delta a$\n(deg)\n0\n-.04\n-.08\n-.12\n-.16\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nNACA\n\n(a) $c_n=-0.4$.\nFigure 9.- Variation of angle of attack and pitching-moment coefficient with Mach number\nfor an NACA 66,1-115 airfoil section equipped with an unsealed 0.20c plain aileron\nof beveled-trailing-edge profile.\n\nNACA TN NO. 1596", "timestamp": "2026-07-22T05:03:25.510200+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 17, "total_pages": 28, "image_filename": "19930085471_p17.jpg", "text": "```markdown\nMACH RM No. L53J11\n\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n\nMach number\n\nFree wall\n\n0 1 2 3 4 5 6 7 8 9 10\nDistance from free wall, in.\n\nWall with model base block\n\nNACA\n\nFigure 3.- Mach number survey of the test section.\n\n15\n\n~~CONFIDENTIAL~~\n[annotation: UNCLASSIFIED]\n\n~~CONFIDENTIAL~~\n[annotation: UNCLASSIFIED]\n```", "timestamp": "2026-07-22T05:03:28.292429+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 47, "total_pages": 47, "image_filename": "19930093773_p47.jpg", "text": "SECURITY INFORMATION\nCONFIDENTIAL\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:03:30.210183+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 26, "total_pages": 49, "image_filename": "19930082498_p26.jpg", "text": "NACA TN No. 1838\n\nSteel wool\n\n19\n\nAsbestos\n\n30\n\nFigure 1.— Continued.\n\nNACA\n\n63", "timestamp": "2026-07-22T05:03:34.648563+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 31, "total_pages": 37, "image_filename": "19930082450_p31.jpg", "text": "30\nNACA TN No. 1778\n\n$$\n\\frac{H_+}{t_w} = 21\n$$\n$$\n\\left(\\frac{b_w}{t_w} = 20\\right)\n$$\n\n$$\n\\sigma_{cr}, ksi\n$$\n12.1\n7.8\n\n$$\n\\frac{S}{t_s} \\text{ or } \\frac{b_s}{t_s}\n$$\n\n$$\n\\frac{P_l}{L\\sqrt{e}}, ksi\n$$\n\nColors indicate minimum weight\nproportions for $\\frac{t_w}{t_s} = 0.63$.\nRed means some other,\nblue means no other value\nof $\\frac{t_w}{t_s}$ gives less weight.\n\n$$\n\\bar{\\sigma}_f, ksi\n$$\n\n31\n(30)\n\n$$\n\\sigma_{cr}, ksi\n$$\n12.1\n7.8\n\n$$\n\\frac{S}{t_s} \\text{ or } \\frac{b_s}{t_s}\n$$\n\n$$\n\\frac{P_l}{L\\sqrt{e}}, ksi\n$$\n\n41\n(40)\n\n$$\n\\sigma_{cr}, ksi\n$$\n16.7\n12.1\n7.8\n\n$$\n\\frac{S}{t_s} \\text{ or } \\frac{b_s}{t_s}\n$$\n\n$$\n\\frac{P_l}{L\\sqrt{e}}, ksi\n$$\n\nNACA\n\n$$\n\\frac{P_l}{t_s}, ksi\n$$\n\nFigure 7.- Direct-reading design chart (alternate form) for 24S-T aluminum-alloy Z-stiffened panels, $\\frac{t_w}{t_s} = 0.63$.", "timestamp": "2026-07-22T05:03:36.290425+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 39, "total_pages": 41, "image_filename": "19930082476_p39.jpg", "text": "NACA TN No. 1801\n37\n\n126\n138\n150\n162\n174\n\n132\n144\n156\n168\n180\n\nFigure 6.- Concluded.\nNACA", "timestamp": "2026-07-22T05:03:39.583058+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 16, "total_pages": 36, "image_filename": "19930085487_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:03:39.968874+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 26, "total_pages": 30, "image_filename": "19930082585_p26.jpg", "text": "NACA TN 1907\n25\n\n<!-- Image (189, 105, 851, 730) -->\n\nFigure 6.- Effect of blade moment of inertia on the variations of descending velocity and rotor angular velocity with time after power failure. Slow exponential pitch change (fig. 2).", "timestamp": "2026-07-22T05:03:43.163320+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 12, "total_pages": 20, "image_filename": "19930085536_p12.jpg", "text": "10\nNACA RM No. E9K05\n\nREFERENCES\n\n1. von Kármán, Theodor, and Moore, Norton B.: Resistance of Slender Bodies Moving with Supersonic Velocities, with Special Reference to Projectiles. Trans. A.S.M.E., vol. 54, no. 23, Dec. 15, 1932, pp.303-310.\n\n2. Tsien, Hsue-Shen: Supersonic Flow over an Inclined Body of Revolution. Jour. Aero. Sci., vol. 5, no. 12, Oct. 1938, pp.480-483.\n\n3. Brown, Clinton E., and Parker, Hermon M.: A Method for the Calculation of External Lift, Moment, and Pressure Drag of Slender Open-Nose Bodies of Revolution of Supersonic Speeds. NACA Rep. No. 808, 1945.\n\n4. Ferri, Antonio: Supersonic-Tunnel Tests of Projectiles in Germany and Italy. NACA ACR No. L5H08, 1945.\n\n5. Sauer, R.: Method of Characteristics for Three-Dimensional Axially Symmetrical Supersonic Flows. NACA TM No. 1133, 1947.\n\n6. Maslen, Stephen H.: Method for Calculation of Pressure Distributions on Thin Conical Bodies of Arbitrary Cross Section in Supersonic Stream. NACA TN No. 1659, 1948.\n\n7. Taylor, G. I., and Maccoll, J. W.: The Air Pressure on a Cone Moving at High Speeds. - I and II. Proc. Roy. Soc. (London), ser. A, vol 139, no. 838, Feb. 1, 1933, pp.278-311.", "timestamp": "2026-07-22T05:03:44.515281+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 21, "total_pages": 78, "image_filename": "19930082618_p21.jpg", "text": "NACA TN 1945\n19\n\nTABLE V\nORDINATES OF THE\nNACA 641-012 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | -.973 |\n| .75 | 1.179 | .75 | -1.179 |\n| 1.25 | 1.190 | 1.25 | -1.190 |\n| 2.5 | 2.095 | 2.5 | -2.095 |\n| 5 | 2.810 | 5 | -2.810 |\n| 7.5 | 3.581 | 7.5 | -3.581 |\n| 10 | 3.871 | 10 | -3.871 |\n| 15 | 4.620 | 15 | -4.620 |\n| 20 | 5.173 | 20 | -5.173 |\n| 25 | 5.576 | 25 | -5.576 |\n| 30 | 5.844 | 30 | -5.844 |\n| 35 | 5.978 | 35 | -5.978 |\n| 40 | 5.981 | 40 | -5.981 |\n| 45 | 5.798 | 45 | -5.798 |\n| 50 | 5.430 | 50 | -5.430 |\n| 55 | 5.056 | 55 | -5.056 |\n| 60 | 4.548 | 60 | -4.548 |\n| 65 | 3.974 | 65 | -3.974 |\n| 70 | 3.350 | 70 | -3.350 |\n| 75 | 2.695 | 75 | -2.695 |\n| 80 | 2.029 | 80 | -2.029 |\n| 85 | 1.382 | 85 | -1.382 |\n| 90 | .786 | 90 | -.786 |\n| 95 | .288 | 95 | -.288 |\n| 100 | 0 | 100 | 0 |\n\nL.E. radius: 1.040\n\nTABLE VI\nORDINATES OF THE\nNACA 641A212 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper Surface | | Lower Surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | 0 |\n| .409 | 1.013 | .591 | -.901 |\n| .648 | 1.233 | .852 | -1.075 |\n| 1.135 | 1.580 | 1.365 | -1.238 |\n| 2.262 | 2.225 | 2.655 | -1.808 |\n| 4.249 | 3.147 | 3.351 | -2.425 |\n| 7.242 | 3.848 | 7.657 | -2.870 |\n| 9.825 | 4.132 | 10.150 | -3.240 |\n| 14.849 | 4.358 | 15.153 | -3.796 |\n| 19.842 | 4.060 | 20.138 | -4.200 |\n| 24.880 | 4.584 | 25.120 | -4.482 |\n| 29.900 | 4.956 | 30.100 | -4.660 |\n| 34.922 | 5.189 | 35.078 | -4.741 |\n| 39.946 | 5.272 | 40.094 | -4.714 |\n| 44.970 | 5.177 | 45.050 | -4.549 |\n| 49.993 | 4.935 | 50.007 | -4.275 |\n| 55.015 | 4.570 | 54.989 | -3.918 |\n| 60.036 | 4.103 | 59.966 | -3.469 |\n| 65.050 | 3.544 | 64.950 | -2.934 |\n| 70.064 | 2.903 | 69.936 | -2.337 |\n| 75.075 | 2.197 | 74.925 | -2.057 |\n| 80.090 | 1.433 | 79.910 | -1.563 |\n| 85.088 | 2.601 | 84.912 | -1.159 |\n| 90.062 | 1.751 | 89.938 | -.773 |\n| 95.032 | .888 | 94.968 | -.398 |\n| 100.000 | .025 | 100.000 | -.025 |\n\nL.E. radius: 0.994\nT.E. radius: 0.028\nSlope of radius through L.E.: 0.095\n\nTABLE VII\nORDINATES OF THE\nNACA 641-612 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | 0 |\n| .260 | 1.098 | .740 | -.798 |\n| .482 | 1.358 | 1.018 | -.938 |\n| .946 | 1.780 | 1.554 | -1.158 |\n| 2.149 | 2.563 | 2.851 | -1.447 |\n| 4.609 | 3.731 | 5.391 | -1.853 |\n| 7.092 | 4.542 | 7.904 | -2.098 |\n| 9.596 | 5.101 | 10.404 | -2.209 |\n| 14.819 | 5.823 | 15.381 | -2.589 |\n| 19.658 | 6.250 | 20.341 | -2.774 |\n| 24.700 | 6.453 | 25.292 | -2.873 |\n| 29.764 | 6.475 | 30.236 | -2.923 |\n| 34.823 | 6.065 | 35.177 | -2.839 |\n| 39.884 | 5.195 | 40.116 | -2.767 |\n| 44.945 | 4.035 | 45.055 | -2.513 |\n| 50.000 | 2.789 | 50.000 | -2.173 |\n| 55.048 | 1.341 | 54.952 | -1.771 |\n| 60.098 | 2.760 | 59.912 | -1.406 |\n| 65.117 | 4.062 | 64.883 | -.882 |\n| 70.125 | 5.263 | 69.865 | -.431 |\n| 75.141 | 6.376 | 74.853 | -.006 |\n| 80.134 | 7.413 | 79.846 | .363 |\n| 85.114 | 8.396 | 84.836 | .642 |\n| 90.082 | 9.333 | 89.818 | .769 |\n| 95.040 | 10.235 | 94.760 | .663 |\n| 100.000 | 0 | 100.000 | 0 |\n\nL.E. radius: 1.040\nSlope of radius through L.E.: 0.2527\n\nNACA", "timestamp": "2026-07-22T05:03:44.691413+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 42, "total_pages": 62, "image_filename": "19930082485_p42.jpg", "text": "NACA TN No. 1810\n41\n\n<!-- Image (235, 126, 778, 839) -->\n\n(a) Critical velocity ratio.\n\n(b) Tangential component of critical velocity ratio.\nFigure 10.- Radial variation of critical ratio $\\frac{V}{V_{cr}}$ and tangential component of critical velocity ratio $\\frac{V_u}{V_{cr}}$ at 0.1 chord downstream of cascade.", "timestamp": "2026-07-22T05:03:44.692207+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 21, "total_pages": 66, "image_filename": "19930082914_p21.jpg", "text": "20\nNACA TN No. 1857\n\nThe differential equation then becomes\n\n$$\ndz = \\frac{dy}{\\sqrt{\\left(\\frac{c_1}{V_o} a + \\frac{c_1}{V_o} by\\right)^2 - 1}}\n$$\n\nBy substitution of\n\n$$\np = \\frac{ac_1}{V_o}\n$$\n\n$$\nq = \\frac{bc_1}{V_o}\n$$\n\nand\n\n$$\n\\xi = p + qy\n$$\n\nthen\n\n$$\ndz = \\frac{d\\xi}{q\\sqrt{\\xi^2 - 1}}\n$$\n\nOn integration,\n\n$$\nq(z + c_2) = \\log_e \\left(\\xi + \\sqrt{\\xi^2 - 1}\\right)\n$$\n\nFor evaluation of the integration constant $c_2$, at $z = y = 0$\n\n$$\nqc_2 = \\log_e \\left[\\frac{c_1}{V_o} a + \\sqrt{\\left(\\frac{c_1}{V_o} a\\right)^2 - 1}\\right]\n$$", "timestamp": "2026-07-22T05:03:46.126461+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 14, "total_pages": 46, "image_filename": "19930085519_p14.jpg", "text": "NACA RM No. L8K19\n13\n\n3. The total maximum rolling-moment coefficient resulting from\n$40^\\circ$ total deflection of a 49-percent-span by 20-percent-chord aileron was\nabout the same as that produced by the plug aileron with the full-span\nslotted flap deflected. The aileron rolling-moment coefficients with the\npartial-span slotted flap deflected were equal to or only slightly greater\nthan those with the flap neutral.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.", "timestamp": "2026-07-22T05:03:52.401875+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 28, "total_pages": 50, "image_filename": "19930082496_p28.jpg", "text": "NACA TN No. 1836\n27\n\n[Figure: A photograph of a mechanical holder and specimen assembly. Labels point to various parts: \"Specimen retainers\", \"Clamp\", \"Support rod\", \"Support ring\", and \"Specimen\". A ruler marked \"INCHES\" with markings for 0 and 1 is visible on the left. In the bottom right corner of the image, there is a NACA logo with the text \"C-21898\" and \"7-29-48\".]\n\nFigure 4. - Holder and specimen for thermal-shock evaluation. Specimen floats in retainers.", "timestamp": "2026-07-22T05:03:55.540379+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 34, "total_pages": 99, "image_filename": "19930082511_p34.jpg", "text": "32\nNACA TN No. 1926\n\nValue of $Mw_1 + Nw_2$ at infinity.- The constants M and N have been determined so that $I.P.(Mw_1 + Nw_2) = 0$ at infinity; furthermore, $R.P.Mw_1 = 0$ at infinity, as is clear from figure 16. Therefore, the value of $Mw_1 + Nw_2$ at infinity is merely $R.P.Nw_2$ at infinity. It is necessary to investigate this limit before choosing the form of $w_3$, because, as was previously noted, the purpose of $w_3$ is to provide $Q(z) = 0$ at infinity. The limit may be written\n\n$$R.P.Nw_2(\\infty) = N \\int_0^1 \\sqrt{\\frac{a^2 - z^2}{1 - z^2}} dz + N \\int_a^\\infty \\sqrt{\\frac{a^2 - z^2}{1 - z^2}} dz$$\n\nThe first term is simply $NaE$. In order to evaluate the second term, substitute $z = \\frac{a}{l}$\n\n$$N \\int_a^\\infty \\sqrt{\\frac{a^2 - z^2}{1 - z^2}} dz$$\n\n$$= Na^2 \\int_0^1 \\sqrt{\\frac{1 - l^2}{a^2 - l^2}} \\frac{dl}{l^2}$$\n\n$$= Na^2 \\int_0^1 \\frac{(1 - l^2)}{\\sqrt{(1 - l^2)(a^2 - l^2)}} \\frac{dl}{l^2}$$\n\n$$= - Na^2 \\int_0^1 \\frac{dl}{\\sqrt{(1 - l^2)(a^2 - l^2)}} + Na^2 \\int_0^1 \\frac{dl}{l^2 \\sqrt{(1 - l^2)(a^2 - l^2)}}$$\n\nThe first term is $-NaK$. In order to evaluate the second term, it is noted (reference 12) that\n\n$$\\frac{d}{dl} \\frac{\\sqrt{(1 - l^2)(a^2 - l^2)}}{l}$$\n\n$$= \\frac{l^4 - a^2}{l^2 \\sqrt{(1 - l^2)(a^2 - l^2)}}$$\n\n$$= \\frac{-a^2}{l^2 \\sqrt{(1 - l^2)(a^2 - l^2)}} + \\frac{l^2 - a^2}{\\sqrt{(1 - l^2)(a^2 - l^2)}} + \\frac{a^2}{\\sqrt{(1 - l^2)(a^2 - l^2)}}$$", "timestamp": "2026-07-22T05:04:03.184814+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 18, "total_pages": 28, "image_filename": "19930085471_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:04:03.394602+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 44, "total_pages": 66, "image_filename": "19930082245_p44.jpg", "text": "```markdown\nNACA TN No. 1596\n\nSection angle of attack, $\\alpha$, deg\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n\nMach number, M\n.1 .2 .3 .4 .5 .6 .7 .8 .9\n\nSection pitching-moment coefficient, $c_m$\n$\\delta_a$ (deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n\nMach number, M\n.1 .2 .3 .4 .5 .6 .7 .8 .9\n\n(b) $c_n = -0.2$.\nFigure 9.—Continued.\n\n[Figure: NACA logo]\n\n43\n```", "timestamp": "2026-07-22T05:04:03.889743+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 5, "total_pages": 33, "image_filename": "19930085544_p5.jpg", "text": "4\nNACA RM No. L8K26\n\n$W_o$\ngeometric resultant velocity, steady part, feet per second (fig. 1)\n\n$W_{wt}$\ninstantaneous geometric resultant velocity, feet per second\n\n$x$\nfractional radius to propeller blade section $\\left(\\frac{r}{R}\\right)$\n\n$\\alpha$\nangle of attack, degrees\n\n$\\alpha_i$\nangle of inflow, degrees\n\n$\\alpha_{P_o}$\namplitude in oscillation of angle of attack, radians or degrees (fig. 1)\n\n$\\alpha_P$\ninstantaneous incremental angle of attack of blade section, radians or degrees\n\n$\\alpha_T$\nangle of inclination of propeller thrust axis, degrees (fig. 1)\n\n$\\beta$\nblade-angle setting at 0.75 radius, degrees\n\n$\\gamma = \\tan^{-1} \\frac{c_d}{c_l}$\n\n$\\epsilon$\nfractional amplitude of stream pulsation $\\left(\\frac{W_{wt_{max}}}{W_o} - 1\\right)$\n\n$\\kappa$\nGoldstein correction factor for finite number of blades\n\n$\\rho$\nmass density of air, slugs per cubic foot\n\n$\\sigma$\nsection solidity $\\left(\\frac{Bc}{2\\pi r}\\right)$\n\n$\\phi_o$\nlocal geometric helix angle, steady part, degrees\n$\\left(\\tan^{-1} \\frac{J}{\\pi x} \\cos \\alpha_T\\right)$\n\n$\\phi_{wt}$\ninstantaneous geometric helix angle, degrees\n$\\left(\\tan^{-1} \\frac{\\cos \\omega_T}{\\frac{\\pi x}{J} + \\sin \\alpha_T \\sin \\omega t}\\right)$\n\n$\\phi$\naerodynamic helix angle, degrees (equation (4))", "timestamp": "2026-07-22T05:04:05.498223+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 27, "total_pages": 49, "image_filename": "19930082498_p27.jpg", "text": "26\n\n40\n\n43\n\n[Figure: Two cylindrical cross-sectional diagrams labeled 40 and 43, showing internal components with dotted patterns and wavy lines; NACA logo present near bottom diagram]\n\nFigure 1.— Continued.\n\nNACA TN No. 1838", "timestamp": "2026-07-22T05:04:07.646286+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 40, "total_pages": 41, "image_filename": "19930082476_p40.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:04:09.684183+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 1, "total_pages": 24, "image_filename": "19930085626_p1.jpg", "text": "Copy No. 370\nRM No. L8K23\n\nNACA RM No. L8K23\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nFREE-FLIGHT INVESTIGATION AT TRANSONIC AND SUPERSONIC\nSPEEDS OF THE ROLLING EFFECTIVENESS OF SEVERAL\nAILERON CONFIGURATIONS ON A TAPERED\nWING HAVING 42.7° SWEEPBACK\n\nBy\n\nCarl A. Sandahl\n\nLangley Aeronautical Laboratory\nLangley Field, Va.\n\nCLASSIFIED DOCUMENT\nThis document contains information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. The transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law.\nAPPROVED BY: W. CROWLEY DATE: 10-14-55\nCHANGE NO. 5118\nWHL\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nJanuary 11, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:04:09.993703+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 7, "total_pages": 46, "image_filename": "19930085542_p7.jpg", "text": "NACA RM No. L8I29\n\nThe relevant dimensions of the models and the test conditions are presented in table I; hereinafter, each model will be referred to by the number designated in the table. All profiles referred to are parallel to the plane of symmetry.\n\nAll the tests were made on a six-component strain-gage balance strut with the models mounted at a point two-thirds of the root chord from the apex of the triangles.\n\nFigure 2 presents the profiles of the series of models having $60^\\circ$ sweepback of the leading edge (models 1, 2, and 3). The models were constructed of laminated mahogany and were given highly polished surfaces. Flat-plate fins of aspect ratio 0.77 and 1.15 were constructed of laminated mahogany and were tested on model 2. Various portions of the tips of the triangular wing of aspect ratio 4 (model 7) were cut off (parallel to the plane of symmetry) to give aspect ratios 3 (model 8), 2 (model 9), and 1 (model 10), including tips of revolution.\n\nAll the models were tested with a small canopy covering the strut head and the cut-out to prevent leakage of air through the wing.\n\nPhotographs of some of the models are presented in figures 3 to 7.\n\nThree series of tests were made. In the first series the lift, longitudinal force, and pitching moment were measured at $\\psi = 0^\\circ$ through an angle-of-attack range from about $\\alpha = -4^\\circ$ to an angle of attack beyond the stall. In the second series of tests the static derivatives were determined by measuring the lateral force, rolling moment, and yawing moment at $\\psi = \\pm 5^\\circ$ through the same angle-of-attack range. In the third series the models were tested through the angle-of-attack range at the values of $pb/2V$ listed in table I to obtain the rolling derivatives $C_{l_p}$, $C_{n_p}$, and $C_{Y_p}$.\n\nAll the tests were made at a dynamic pressure of 24.9 pounds per square foot which, when based on the mean aerodynamic chords of the models, corresponds to the Reynolds numbers in table I. The test Mach number was 0.13.\n\nCORRECTIONS AND ACCURACY\n\nThe test data were transferred from the model-mounting position (a point at two-thirds of the root chord from the apex of the triangles) to the quarter-chord point of the mean aerodynamic chord.", "timestamp": "2026-07-22T05:04:11.005756+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 27, "total_pages": 30, "image_filename": "19930082585_p27.jpg", "text": "```markdown\n26\nNACA TN 1907\n\n<!-- Image (112, 143, 804, 666) -->\n\nFigure 7.- Effect of rate of pitch reduction on $\\lambda$ against $\\theta$ throughout the transition maneuver. $I_1 = 200$ slug-feet$^2$.\n```", "timestamp": "2026-07-22T05:04:13.548347+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 32, "total_pages": 37, "image_filename": "19930082450_p32.jpg", "text": "```markdown\nNACA TN No. 1778\n31\n\n$\\bar{\\sigma}_f, ksi$\n\n40\n35\n30\n25\n20\n15\n10\n40\n35\n30\n25\n20\n15\n40\n35\n30\n25\n20\n15\n\n15 20 25 30 35 40 45 50 55 60 65 70\n\n$\\frac{P_t}{t_s}, ksi$\n\nFigure 7.- Concluded. $\\frac{t_w}{t_s} = 0.63$.\n\nNACA\n\n<!-- Image (150, 110, 874, 934) -->\n```", "timestamp": "2026-07-22T05:04:15.065479+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 17, "total_pages": 36, "image_filename": "19930085487_p17.jpg", "text": "NACA RM No. E8J22\n\n[Figure: Sand pattern for third bending-mode vibration on first-stage blade showing location of two nodes other than one at blade root. Frequency of vibration, 5200 cycles per second.]\n\nFigure 4. - Sand pattern for third bending-mode vibration on first-stage blade showing location of two nodes other than one at blade root. Frequency of vibration, 5200 cycles per second.\n\nNACA\nC-14378\n2-27-45\n\n15", "timestamp": "2026-07-22T05:04:16.236346+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 13, "total_pages": 20, "image_filename": "19930085536_p13.jpg", "text": "NACA RM No. E8K05\n\n[Figure: A black-and-white photograph showing a conical object mounted on a support structure. The cone is attached via multiple thin rods or struts to a larger, angular metallic body that appears to be part of a test rig or wind tunnel mount. The background is plain and light-colored.]\n\nFigure 1. - Cone mounted on support body.\n\nNACA  \nC-21730  \n6-23-48", "timestamp": "2026-07-22T05:04:16.386998+00:00"}