PDF/ocr_dataset_1021.jsonl

{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 114, "total_pages": 118, "image_filename": "19930085838_p114.jpg", "text": "112\nNACA RM No. L9B23\n\nAileron section hinge-moment coefficient, $c_{h_a}$\n\n| $\\alpha_f$ | $\\alpha_b$ |\n| :--- | :--- |\n| $\\circ$ $0^\\circ$ | $.351c_a$ |\n| $\\circ$ $0^\\circ$ | $.408c_a$ |\n| $\\square$ $25^\\circ$ | $.351c_a$ |\n| $\\square$ $25^\\circ$ | $.408c_a$ |\n| $\\diamond$ $40^\\circ$ | $.351c_a$ |\n| $\\diamond$ $40^\\circ$ | $.408c_a$ |\n\nAileron deflection, $\\delta_a$\n\nNACA\n\n(a) 0.177c thick airfoil.\n\nFigure 20.- Aileron hinge-moment characteristics of a straight-sided Frise aileron on two NACA 7-series-type airfoils with double slotted flap and flip. $\\alpha_o = 0^\\circ$; $\\delta_f = 0^\\circ$; $R = 6.0 \\times 10^6$ (approx.).", "timestamp": "2026-07-22T04:51:30.439243+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 31, "total_pages": 47, "image_filename": "19930093773_p31.jpg", "text": "30\nNACA RM E9G09\n\n<!-- Image (139, 139, 860, 829) -->\n\n(b) Air flow.\nFigure 5. - Continued. Effect of flight Mach number on variation of\nengine performance with engine speed at altitude of 25,000 feet.", "timestamp": "2026-07-22T04:51:31.910065+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 27, "total_pages": 33, "image_filename": "19930082487_p27.jpg", "text": "NACA TN No. 1813\n25\n\n<!-- Image (120, 135, 898, 859) -->\n\n(e) NACA 65,-206.\n(f) NACA 65,-208.\n(g) NACA 65,-210\n(h) NACA 65,-212.\nFigure 7.- Continued.", "timestamp": "2026-07-22T04:51:39.345831+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 9, "total_pages": 58, "image_filename": "19930082617_p9.jpg", "text": "8\nNACA TN 1962\n\nTABLE I\nDATA OF CYLINDERS\n\n| Cylinder | Number stringers | Number rings | Stringer size | Ring size | Sheet thickness (in.) | Radius of cylinder (in.) | Angle of cutout (deg) | Length of cylinder (in.) | Length of cutout (in.) | Percent length cutout | Ring spacing |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 72 | 16 | 14 | 3/8 x 3/8 | 1/8 x 1/2 | 0.012 | 10 | 45 (bottom) | 57.9 | 34.74 | 60 | 3.86 |\n| 73 | 16 | 14 | 3/8 x 3/8 | 1/4 x 1/2 | .012 | 10 | 45 (bottom) | 57.9 | 34.74 | 60 | 3.86 |\n| 74 | 16 | 22 | 3/8 x 3/8 | 1/4 x 1/2 | .012 | 10 | 45 (bottom) | 59.11 | 33.41 | 56.5 | 2.57 |\n| 75 | 16 | 22 | 3/8 x 3/8 | 3/8 x 3/8 | .012 | 10 | 45 (bottom) | 59.11 | 33.41 | 56.5 | 2.57 |\n| 76 | 16 | 15 | 3/8 x 3/8 | 1/8 x 1/2 | .012 | 10 | 2-45 (two sides) | 61.78 | 47.72 | ---- | 3.86 |\n| 77 | 16 | 29 | 3/8 x 3/8 | 3/8 x 3/8 | .012 | 10 | 45 (bottom) | 77.10 | 51.40 | 66.7 | 2.57 |\n| 78 | 16 | 22 | 3/8 x 3/8 | 1/4 x 1/2 | .012 | 10 | 2-45 (two sides) | 59.11 | 47.71 | ---- | 2.57 |\n| 79 | 16 | 22 | 3/8 x 3/8 | 1/8 x 1/2 | .012 | 10 | 45 (two sides) | 59.11 | 47.71 | ---- | 2.57 |\n\n$^a$Length of each of five symmetrical cutouts.\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T04:51:41.660258+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 8, "total_pages": 78, "image_filename": "19930082618_p8.jpg", "text": "6\nNACA TN 1945\n\nthan with the balance. Hence, all drag measurements were taken by the\nwake-survey method (reference 4) with the gaps between the model and\ntunnel walls sealed with felt packing.\n\nTests.- The tests of each smooth, plain airfoil consisted of\nmeasurements of the section lift, drag, and quarter-chord pitching\nmoment at Reynolds numbers of $2.0 \\times 10^6$, $1.5 \\times 10^6$, $1.0 \\times 10^6$,\nand $0.7 \\times 10^6$. In none of these tests did the Mach number exceed 0.15.\nWith the exception of the NACA 641A212 airfoil section, lift and pitching-\nmoment measurements at each of the four Reynolds numbers were also made\nfor the smooth airfoils equipped with 0.20c simulated split flaps\ndeflected 60°. In addition, all of the measurements except those of the\npitching moment were repeated with standard roughness applied to the\nleading edges of the airfoils. The standard roughness employed was the\nsame as that used in previous investigations (references 1 to 3) and\nconsisted of 0.011-inch-diameter carborundum grains spread over a\nsurface length of 8 percent of the chord measured from the leading edge\non the upper and lower surfaces of the airfoils. The grains were thinly\nspread to cover from 5 to 10 percent of this area.\n\nIn order that comparative data should be available for all the\nairfoils in the range of Reynolds number from $9.0 \\times 10^6$ to $0.7 \\times 10^6$, it\nwas necessary to make standard tests (reference 1) in the Langley two-\ndimensional low-turbulence pressure tunnel of the NACA 64-409\nand NACA 641-612 airfoils at Reynolds numbers of $3.0 \\times 10^6$, $6.0 \\times 10^6$,\nand $9.0 \\times 10^6$ since these data had not previously been obtained. In\naddition, supplementary tests were made in the Langley two-dimensional\nlow-turbulence pressure tunnel at a Reynolds number of $6.0 \\times 10^6$ of the\nNACA 23012 and NACA 23015 sections equipped with split flaps. Such data\nare available in references 1 to 3 for the other airfoils tested in the\npresent investigation (with exceptions as already noted) and were\nconsidered necessary for the NACA 23012 and NACA 23015 sections in order\nto compare adequately the type of scale effect shown by those sections\nwith that of the other sections tested.\n\nRESULTS\n\nThe results are presented (figs. 1 to 15) in the form of standard\naerodynamic coefficients representing the lift, drag, and quarter-chord\npitching moment. Each figure is in three parts. The lift data for the\nplain airfoils and the airfoils with split flaps are contained in parts (a)\nand (b), respectively, together with the appropriate quarter-chord\npitching-moment data; the drag results and data on the aerodynamic\ncenter and the moment coefficient about this point are contained in part (c).", "timestamp": "2026-07-22T04:51:42.197614+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 10, "total_pages": 36, "image_filename": "19930082614_p10.jpg", "text": "8\nNACA TN 1939\n\nConstant flight-path inclination.- When the flight path is inclined, the component of weight W parallel to the direction of motion must be included, and the equation is\n\n$$\n\\frac{dV}{dt} = \\frac{g}{W} (-D_n - W \\sin \\gamma)\n$$\n\nwhere $\\gamma$ is the flight-path angle, positive for climbing flight. Substituting K as in the case of level flight,\n\n$$\n\\frac{dV}{dt} = -KV^2 - g \\sin \\gamma \\tag{3}\n$$\n\nThe assumption can again be made that K does not vary during the period of time being considered. Then (as shown in reference 2) equation (3) integrates as follows to give the velocity variation for an airplane in a dive or climb at a constant angle:\n\nIf $\\gamma$ is negative and $\\sqrt{L/K} < V_0$, where $L = -g \\sin \\gamma$\n\n$$\nV = \\sqrt{L/K} \\coth \\left[ \\sqrt{L/K} (Kt + C_1) \\right] \\tag{4}\n$$\n\nwhere\n\n$$\nC_1 = \\frac{1}{2} \\sqrt{K/L} \\log_e \\frac{V_0 \\sqrt{K} + \\sqrt{L}}{V_0 \\sqrt{K} - \\sqrt{L}}\n$$\n\nIf $\\sqrt{L/K} > V_0$,\n\n$$\nV = \\sqrt{L/K} \\tanh \\left[ \\sqrt{L/K} (Kt + C_2) \\right] \\tag{5}\n$$\n\nwhere\n\n$$\nC_2 = \\frac{1}{2} \\sqrt{K/L} \\log_e \\frac{\\sqrt{L} + V_0 \\sqrt{K}}{\\sqrt{L} - V_0 \\sqrt{K}}\n$$\n\nIf $\\gamma$ is positive,\n\n$$\nV = N \\cot \\left[ N (Kt + C_3) \\right] \\tag{6}\n$$", "timestamp": "2026-07-22T04:51:43.887547+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 18, "total_pages": 50, "image_filename": "19930082496_p18.jpg", "text": "NACA TN No. 1836\n17\n\nTABLE I - SHORT-TIME TENSILE EVALUATION\n\n| 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 | 11 | Remarks |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Specimen | Soaking time (hr) | Soaking temperature (°F) | Evaluation temperature (°F) | Breaking stress (psi) | Specimen diameter at fracture point before test (in.) | Specimen area at fracture point (sq in.) | Fracturing stress (lb/sq in.) | Specimen diameter at fracture point after oxide scale removal (in.) | Specimen area at fracture point (sq in.) | Corrected fracturing stress (lb/sq in.) | |\n| 301 | $12\\frac{1}{2}$ | 1800 | 1800 | 6240 | 0.5000 | 0.1963 | 31,800 | 0.5000 | 0.1963 | 31,800 | Fluorescent-oil surface inspection, satisfactory; Radiograph, fine chemical segregation |\n| 302 | $13\\frac{1}{2}$ | 2300 | 2200 | 1525 | .4990 | .1955 | 7,800 | .4670 | .1712 | 8,900 | Fluorescent-oil surface inspection, satisfactory; Radiograph, fine chemical segregation |\n| 303 | -- | -- | -- | -- | .5000 | -- | -- | -- | -- | -- | Broke during soaking at 2300° F. Load, 200 pounds. Fluorescent-oil surface inspection, satisfactory; Radiograph, fine chemical segregation |\n| 304 | 4 | 2300 | 2200 | 2350 | .4998 | .1958 | 12,000 | .4753 | .1772 | 13,200 | Fluorescent-oil surface inspection, satisfactory; Radiograph, fine chemical segregation |\n| 305 | 12 | 1900 | 1800 | 6820 | .5016 | .1975 | 34,600 | .5010 | .1970 | 34,600 | Fluorescent-oil surface inspection, satisfactory; Radiograph, fine chemical segregation |\n\n[Figure: NACA logo]\n\n$^a$Calculated from diameter of column 6.\n$^b$Calculated from diameter of column 9.", "timestamp": "2026-07-22T04:51:46.693510+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 10, "total_pages": 30, "image_filename": "19930082585_p10.jpg", "text": "NACA TN 1907\n\n$$\nB_{7} = \\left( \\frac{\\rho a c R^{3} \\Omega_{a}}{6 I_{1}} \\right) \\left( \\frac{ \\frac{3}{4} \\Omega_{a} R A_{2} a_{2} - \\frac{\\rho a b c R^{3} \\Omega_{a}^{2}}{4 \\theta M / g} A_{2} }{ a_{2}^{3} - b_{1} a_{2}^{2} + b_{2} a_{2} - b_{3} } \\right)\n\\tag{8d}\n$$\n\nThe initial conditions, which must be satisfied by the complete solution for $\\beta(t)$, are that, at $t = 0$,\n\n$$\n\\beta = \\beta_{0}\n\\tag{9}\n$$\n\n$$\n\\dot{\\beta} = \\dot{\\beta}_{0} = 0\n\\tag{9a}\n$$\n\n$$\n\\ddot{\\beta} = \\ddot{\\beta}_{0} = \\frac{\\rho a c R^{4} \\Omega_{0}^{2}}{2 I_{1}} \\left( \\frac{V_{0} - v_{0}}{3 \\Omega_{0} R} + \\frac{\\theta_{0}}{4} \\right) - \\beta_{0} \\Omega_{0}^{2}\n\\tag{9b}\n$$\n\nTherefore, from equations (9),\n\n$$\nB_{1} + B_{2} + B_{4} + B_{5} + B_{6} + B_{7} = \\beta_{0}\n\\tag{10}\n$$\n\n$$\nm_{1} B_{1} + \\alpha B_{2} + \\omega B_{3} - k B_{5} - a_{1} B_{6} - a_{2} B_{7} = 0\n\\tag{10a}\n$$\n\n$$\nm_{1}^{2} B_{1} + (\\alpha^{2} - \\omega^{2}) B_{2} + 2 \\alpha \\omega B_{3} + k^{2} B_{5} + a_{1}^{2} B_{6} + a_{2}^{2} B_{7} = \\ddot{\\beta}_{0}\n\\tag{10b}\n$$\n\nEquations (10), (10a), and (10b) can be solved for $B_{1}$, $B_{2}$, and $B_{3}$, when $B_{4}$ through $B_{7}$ have been found by equations (8). The complete solution, then, is\n\n$$\n\\beta(t) = B_{1} e^{m_{1} t} + e^{\\alpha t} \\left( B_{2} \\cos \\omega t + B_{3} \\sin \\omega t \\right) +\n$$\n\n$$\nB_{4} + B_{5} e^{-k t} + B_{6} e^{-a_{1} t} + B_{7} e^{-a_{2} t}\n\\tag{11}\n$$", "timestamp": "2026-07-22T04:51:47.826464+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 24, "total_pages": 41, "image_filename": "19930082476_p24.jpg", "text": "CHART 6.- EFFECT OF INDIVIDUAL AILERON DEFLECTIONS - Concluded\n\nD. Loading 3 $\\left(\\frac{I_x - I_y}{mb^2} = 165 \\times 10^{-4}; \\mu = 3.38; \\text{loading 3 in table II and point 3 in fig. 4}\\right)$\n\nRight aileron up setting, degrees\nLeft aileron down setting, degrees\nRight aileron up setting, degrees\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:51:50.972173+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 19, "total_pages": 99, "image_filename": "19930082511_p19.jpg", "text": "NACA TN No. 1826\n17\n\n7. Electrical analogies of either the velocity-potential or the acceleration-potential type may be devised to correspond to most of the problems discussed.\n\n8. In electrical analogies that represent velocity potential by electrical potential, the condition of continuity at the entrance lip appears troublesome, especially for three-dimensional tunnels; however, the exit conditions are easily represented.\n\n9. In electrical analogies that represent acceleration potential by electrical potential, the entrance-lip condition is automatically satisfied, but fulfillment of exit conditions is troublesome. Rough approximation of the exit conditions may, however, be adequate for many purposes.\n\n10. Acceleration-potential analogies are experimentally simpler than velocity-potential analogies.\n\nII - TWO-DIMENSIONAL TUNNELS\n\nIn part II, boundary-induced velocities in two-dimensional open tunnels are derived with special reference to the effects of the closed entrance and exit regions. The cases treated are:\n\n(1) Tunnel with a closed entrance (upstream) region but without a closed exit region\n\n(2) Tunnel with a closed entrance region but with only one exit lip (corresponding to a condition in which the downward deflection of the flow is so large that the flow makes contact only with the lower exit lip)\n\n(3) Tunnel with closed entrance and exit regions\n\n(4) Same as case 3, but with different pressures on the two free surfaces\n\nNumerical results are given for all cases.\n\nSYMBOLS AND DIMENSIONS\n\nEach tunnel is idealized as a strip of uniform height $h$, having a stream velocity $V$, and containing a point vortex of strength $\\Gamma$. For simplification of the present development, lengths and velocities will be made nondimensional by dividing by $h$ and $V$, respectively, and the vortex strength will be made nondimensional by dividing by $hV$.", "timestamp": "2026-07-22T04:51:53.138739+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 5, "total_pages": 66, "image_filename": "19930082914_p5.jpg", "text": "4\nNACA TN No. 1857\n\ncoordinates of the nozzle blocks were calculated by the method of characteristics. No allowance was made for the thickness of the boundary layer, and the nozzle gave the Mach number for which it was designed. A photograph of similar nozzles is shown as figure 1, together with the section for the transition from a 6-inch circular pipe to the rectangular entrance to the nozzle.\n\nInterferometer\n\nIntroduction.- The Mach-Zehnder type of interferometer was originated by L. Mach and Zehnder for use in studying phenomena of gas dynamics. The instrument is useful in the study of gas-flow problems because it gives an instantaneous record from which can be calculated the variation of gas density throughout a flow field. It is particularly applicable to the study of high-speed gas flows, because, by using light waves, it makes unnecessary the insertion into the flow of probes or other measuring instruments that would disturb the flow. The interferometer gives quantitative results in which a rather high degree of accuracy can be obtained. It has a sufficient range of sensitivity to measure both small density changes, as in a weak Prandtl-Meyer expansion, and large density changes, as across a strong shock wave.\n\nThe interferometer was used, but not intensively, by Mach and Zehnder. It was applied to the study of subsonic aerodynamics by Zobel (reference 9). Since then it has been applied by Ladenburg and his coworkers to the study of phenomena in supersonic flow (reference 10).\n\nThe basic arrangement of the Mach-Zehnder interferometer is shown in figure 2. Light from a source S is made into a beam of parallel rays by a collimating lens system $I_1$. This beam falls on the splitter plate $S_1$ where it is split into two beams. Part of the original beam of light is reflected by the splitter plate $S_1$ and part is transmitted by it. The part that is transmitted goes to the mirror $M_1$ where it is reflected onto the splitter plate $S_2$. A portion of the beam is transmitted by $S_2$ and is not used. The other portion is reflected by $S_2$, passes through the lens $I_2$, and falls on a screen or a photographic plate P.\n\nThe light that was reflected by $S_1$ likewise goes to a totally reflecting mirror $M_2$ from which it is reflected onto the splitter plate $S_2$. At $S_2$ a part of the beam is reflected and is not used, and the remainder is transmitted, passes through the lens $I_2$, and falls on the screen or the photographic plate P. This arrangement fulfills one of the necessary conditions for the interference of the waves in two beams of light; that is, that the two beams originate in the same light source. From a practical standpoint the arrangement also permits the condition to", "timestamp": "2026-07-22T04:51:54.307458+00:00"}
{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 6, "total_pages": 14, "image_filename": "19930082712_p6.jpg", "text": "```markdown\n4\nNACA TN 1998\n\nRESULTS AND DISCUSSION\n\nThe results of the present tests, together with the lower Reynolds number data of reference 2 for the NACA 8-H-12 airfoil, are shown in figure 1 as standard plots of section lift, drag, and pitching-moment coefficients for both the smooth airfoil and the airfoil with roughened leading edge. In addition, some of the more important aerodynamic characteristics of the NACA 8-H-12 airfoil, along with those of the NACA 0012 and NACA 23012 sections for comparison (from reference 4), are shown plotted against Reynolds number in figure 2.\n\nIn connection with the comparison of the data of reference 2 with those of the present investigation, it should be noted that the surface length of roughness employed in the present investigation was different from that employed in the tests of reference 2. Roughness was applied to the leading edge for the tests of reference 2 for a surface length of 0.02c along each surface measured from the leading edge as compared with 0.08c for the present tests.\n\nCorrections for tunnel-wall interference have been made to all data procured from the tunnel by the following equations (developed in reference 5) in which the primed quantities represent those measured in the tunnel:\n\n$$\n\\begin{aligned}\n\\alpha_0 &= 1.015\\alpha_0' \\\\\nc_d &= 0.992c_d' \\\\\nc_l &= 0.977c_l' \\\\\nc_{m_p} &= 0.992c_{m_p}'\n\\end{aligned}\n$$\n\nLift.- The maximum section lift coefficient of the smooth NACA 8-H-12 airfoil increases from 1.25 to 1.48 as the Reynolds number is increased from $1.8 \\times 10^6$ to $11.0 \\times 10^6$ (figs. 1(a) and 2(a)). Between Reynolds numbers of $1.8 \\times 10^6$ and $3.0 \\times 10^6$ the maximum lift remains relatively constant. The largest increment in maximum lift coefficient resulting from increases in the Reynolds number occurs between $3.0 \\times 10^6$ and $6.0 \\times 10^6$, with smaller increases occurring up to a Reynolds number of $11.0 \\times 10^6$. The shape of the lift curve near maximum lift is very desirable for all Reynolds numbers. The lift-curve slope of the smooth airfoil, measured from approximately zero lift to slightly above the experimental design lift, increases from a value of 0.098 to\n```", "timestamp": "2026-07-22T04:52:03.554774+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 34, "total_pages": 66, "image_filename": "19930082245_p34.jpg", "text": "```markdown\nNACA TN No. 1596\n\nAileron section normal-force coefficient, $c_{n\\delta}$\n\nAileron section hinge-moment coefficient, $c_h$\n\n| $\\delta_a$ (deg) | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:52:06.694876+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 28, "total_pages": 33, "image_filename": "19930082487_p28.jpg", "text": "26\nNACA TN No. 1813\n\n.12\n.08 (Mg/calc\n.04\nSection drag coefficient, Cd\n0\n.6 .7 .8\nFree-stream Mach number, Mo\n(i) NACA 66,-206.\n\n.6 .7 .8\n(j) NACA 66,-208.\n\n.12\n.08\n.04\n0\n.6 .7 .8\nFree-stream Mach number, Mo\n(k) NACA 66,-212.\n\n$\\alpha$\n(deg)\n$\\circ$ -4\n$\\square$ -2\n$\\diamond$ 0\n$\\triangle$ 2\n$\\nabla$ 4\n$\\triangleright$ 6\n\n.6 .7 .8\n(l) NACA 66-006.\nNACA\n\nFigure 7.- Continued.", "timestamp": "2026-07-22T04:52:11.536822+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 27, "total_pages": 62, "image_filename": "19930082485_p27.jpg", "text": "26\nNACA TN No. 1810\n\nREFERENCES\n\n1. Lighthill, M. J.: A Mathematical Method of Cascade Design.\nR. & M. No. 2104, British A.R.C., 1945.\n\n2. Weinig, F.: The Flow around Turbine and Compressor Blades.\nCGD 291, reproduced by Code 338, Res. and Standards Branch,\nBuShips, Navy Dept., May 1946. (Abs. Bib. Sci. Ind. Res.,\nvol. 2, no. 13, April 5, 1946, p. 997, PB 28689.)\n\n3. Goldstein, Arthur W., and Jerrison, Meyer: Isolated and Grid\nAirfoils with Prescribed Velocity Distribution. NACA TN\nNo. 1308, 1947.\n\n4. Katzoff, S., Finn, Robert S., and Laurence, James C.: Inter-\nference Method for Obtaining the Potential Flow past an\nArbitrary Cascade of Airfoils. NACA TN No. 1252, 1947.\n\n5. Stodola, A.: Steam and Gas Turbines. Vol. II. McGraw-Hill\nBook Co., Inc., 1927, p. 992. (Reprinted, Peter Smith (New\nYork), 1945.)\n\n6. Goldstein, Arthur W.: Analysis of the Performance of a Jet\nEngine from Characteristics of the Components. I - Aero-\ndynamics and Matching Characteristics of the Turbine\nComponent Determined with Cold Air. NACA TN No. 1459, 1947.\n\n7. Jahnke, Eugen, and Emde, Fritz: Table of Functions. Dover\nPub. (New York), 4th ed., 1945, p. 32.", "timestamp": "2026-07-22T04:52:12.895947+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 115, "total_pages": 118, "image_filename": "19930085838_p115.jpg", "text": "NACA RM No. L9B23\n113\n\nAileron section hinge-moment coefficient, $C_{h_a}$\n\n| $\\delta_f$ | $c_b$ |\n| :--- | :--- |\n| $\\circ$ $0^\\circ$ | $0.351c_a$ |\n| $\\circ$ $0^\\circ$ | $.408c_a$ |\n| $\\square$ $25^\\circ$ | $.351c_a$ |\n| $\\square$ $25^\\circ$ | $.408c_a$ |\n| $\\diamond$ $40^\\circ$ | $.351c_a$ |\n| $\\diamond$ $40^\\circ$ | $.408c_a$ |\n\nAileron deflection, $\\delta_a$\n\n(b) $0.154c$ thick airfoil.\n\nFigure 20.- Concluded", "timestamp": "2026-07-22T04:52:14.702374+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 19, "total_pages": 50, "image_filename": "19930082496_p19.jpg", "text": "18\nNACA TN No. 1836\n\nTABLE II - THERMAL-SHOCK EVALUATION\n[Radiography indicated fine chemical segregation in ceramal\nspecimens, none in ceramic. No external flaws existed.]\n\n| Speci- men | Constituent | Weight (percent) | Cycles completed | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | $1800^\\circ$ F | $2000^\\circ$ F | $2200^\\circ$ F | $2400^\\circ$ F |\n| -3D13 | Titanium carbide (TiC) | 80 | 25 | 25 | 25 | 25 |\n| | Cobalt (Co) | 20 | | | | |\n| 3D14 | Titanium carbide (TiC) | 80 | 25 | 25 | 25 | 25 |\n| | Cobalt (Co) | 20 | | | | |\n| 7A10 | Zircon ($ZrSiO_4$) | 100 | 1 | ----- | ----- | ----- |\n| 3A7 | Titanium carbide (TiC) | 100 | 25 | 25 | 25 | 14 |\n| 3A10 | Titanium carbide (TiC) | 100 | 25 | 25 | 25 | 21 |\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T04:52:20.003919+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 11, "total_pages": 53, "image_filename": "19930082542_p11.jpg", "text": "10\nNACA TN No. 1867\n\nHot-cold-working at 1200° F did not appreciably change the microstructure of the alloy. The photomicrographs in figure 13 show that relatively little grain distortion occurred and that the major effect was some twinning and increase in the ease of etching of grain boundaries in comparison with the plain solution-treated material.\n\nAging prior to hot-cold-work.- The major effect of aging solution-treated material for 24 hours at 1400° F prior to 10-percent hot-cold-work at 1200° F was a slight reduction in rupture strength and a considerable increase in rupture test ductility when compared with plain solution-treated and hot-cold-worked material. (See fig. 12.)\n\nAging after hot-cold-work.- Aging for 24 hours at 1400° F after 15-percent reduction at 1200° F lowered both the room-temperature and rupture strengths of 2050° F solution-treated material over the unaged material. The rupture test ductility was materially improved. (See fig. 10.)\n\nWhen solution-treated at 2200° F, however, the major change produced by aging after hot-cold-work was a substantial increase in rupture strength compared with the unaged material.\n\nEffect of Holding Time in the Rupture Test Unit before Testing\n\nThe possibility exists that the results of the rupture tests might be influenced by aging during the standard test procedure of holding specimens for approximately 24 hours in the units for temperature adjustment before applying the stress. Partial data from a study of the effect of holding time in table III are inconclusive. Materials solution-treated at 2200° F apparently were strengthened somewhat by holding 24 hours while those solution-treated at 2050° F were weakened. In either case the variation was not enough to change the results of the investigation.\n\nDISCUSSION OF RESULTS\n\nThe properties of low-carbon N-155 alloy can be influenced to a pronounced degree by heat treatment and hot-cold-work. The ranges in properties for each type of treatment are summarized in figure 14. For the particular hot-rolled stock tested, strength at room temperature will be reduced by any heat treatment alone and only hot-cold-work will increase it. Only the rupture strength at 1000 hours can be improved by solution and aging treatments. Marked increases in room-temperature strength and in rupture strength can be obtained by hot-cold-work. For any type of treatment, however, a range in properties will result depending on the conditions of treatment.", "timestamp": "2026-07-22T04:52:20.592025+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 9, "total_pages": 78, "image_filename": "19930082618_p9.jpg", "text": "NACA TN 1945\n\nThe Reynolds number range for which the plain, smooth airfoil characteristics are presented extends from $9.0 \\times 10^6$ to $0.7 \\times 10^6$. Data are presented for each of the plain airfoils with roughened leading edges and for most of the airfoils with split flaps in both the smooth and rough surface conditions at five Reynolds numbers from $6.0 \\times 10^6$ to $0.7 \\times 10^6$. The characteristics at a Reynolds number of $6.0 \\times 10^6$ of the NACA 4412 and NACA 4415 sections with split flaps are not available in reference 1 for the rough surface condition, nor were these data obtained in the present investigation. From the quarter-chord pitching-moment data, the position of the aerodynamic center and the variation of the moment about this point were calculated and are presented for each of the plain, smooth airfoils (figs. 1 to 15, part (c)).\n\nThe influence of the tunnel boundaries has been removed from the aerodynamic data for all the airfoil sections. The following equations (developed in reference 4) which contain the correction factors for the NACA $6h_2$-415 airfoil show the order of magnitude of the boundary effect:\n\n$$\n\\begin{aligned}\nc_a &= 0.991 c_a' \\\\\nc_l &= 0.976 c_l' \\\\\nc_{m_c/4} &= 0.991 \\, c_{m_c/4}' \\\\\n\\alpha_0 &= 1.015 \\alpha_0'\n\\end{aligned}\n$$\n\nwhere the primed quantities represent the coefficients measured in the tunnel.\n\nDISCUSSION\n\nA detailed evaluation of the comparative merits of a large number of airfoils is given in reference 1 for a Reynolds number of $6.0 \\times 10^6$. In the present paper such a detailed evaluation is not attempted for each of the seven Reynolds numbers investigated, but, rather, the data are analyzed to show the effects of several airfoil design parameters upon the manner in which the more important aerodynamic characteristics of the airfoils vary with Reynolds number. As an aid to this study, cross plots (figs. 16 to 22) are used to show some of the important aerodynamic characteristics of the airfoils as functions of Reynolds number. The aerodynamic characteristics discussed concern the drag, the lift, and the pitching moment.", "timestamp": "2026-07-22T04:52:20.693311+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 10, "total_pages": 58, "image_filename": "19930082617_p10.jpg", "text": "```markdown\nNACA TN 1962\n9\n\nTABLE II\nEXPERIMENTAL MAXIMUM STRAINS AND MOMENTS\n\n<!-- Table (175, 130, 808, 888) -->\n\\begin{tabular}{|c|c|c|c|c|c|c|}\n\\hline\nCylinder & Experimental maximum moment & Experimental maximum strain & Type of failure & \\multicolumn{2}{c|}{Description of cylinder after buckling} & Jack load \\\\\n\\cline{5-6}\n& & & & Near side or stringer 8 & Far side or stringer 11 & \\\\\n\\hline\n72 & 212,544 & -29.72 & General instability & Out \\newline $\\frac{1}{2}$ sine, 9 fields & In \\newline $\\frac{1}{2}$ sine, 9 fields & 2992 \\\\\n\\hline\n73 & 272,664 & -21.18 & General instability & In \\newline $\\frac{1}{2}$ sine, 8 fields & Out \\newline $\\frac{1}{2}$ sine, 8 fields & 3787 \\\\\n\\hline\n74 & 323,208 & -21.29 & General instability$^1$ & Out \\newline $\\frac{1}{2}$ sine, 10 fields & Out \\newline $\\frac{1}{2}$ sine, 13 fields & 4489 \\\\\n\\hline\n75 & 393,600 & -36.32 & Local failure$^2$ & Out \\newline $\\frac{1}{2}$ sine, 9 fields & Out \\newline S-shape, 13 fields & 5189 \\\\\n\\hline\n76 & 324,000 & -20.35 & General instability & Side cutout & Side cutout & 4900 \\\\\n\\hline\n77 & 306,720 & -28.50 & General instability & Out \\newline $\\frac{1}{2}$ sine, 11 fields & Out \\newline $\\frac{1}{2}$ sine, 15 fields & 4510 \\\\\n\\hline\n78 & 451,224 & -32.71 & Tension failure$^3$ & Side cutout & Side cutout & 6267 \\\\\n\\hline\n79 & 370,800 & -23.58 & General instability & Side cutout & Side cutout & 5190 \\\\\n\\hline\n\\end{tabular}\n\n$^1$After a certain point the load gradually decreased (568 lb). Skin began to rupture and there was no collapse of the structure.\n$^2$Bolt at intersection of stringer and ring sheared, which precipitated failure that resembled general type of instability.\n$^3$Stringers on tension side of cylinder failed in tension.\n\n[Figure: NACA logo]\n```", "timestamp": "2026-07-22T04:52:21.437489+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 11, "total_pages": 30, "image_filename": "19930082585_p11.jpg", "text": "10\nNACA TN 1907\n\nSolution for $\\Omega$ and V\n\nThe remaining two variables are $V(t)$ and $\\Omega(t)$. The equations to be solved are the thrust and torque equations:\n\n$$\n\\dot{V} = g - \\frac{\\rho a b c R^3}{2W/g} \\left( \\frac{V - v}{2R} \\Omega - \\frac{\\dot{\\beta}}{3} \\Omega + \\frac{\\theta}{3} \\Omega^2 \\right) \\quad (12)\n$$\n\nand\n\n$$\n\\dot{\\Omega} = \\frac{\\rho c R^4 \\Omega^2}{2I_1} \\left[ \\frac{\\lambda^2 f_1}{2} - \\frac{\\lambda}{3} (f_3 - f_1 f_2) - \\frac{f_2 f_3 + \\delta_0'}{4} \\right] + \\frac{2\\beta \\dot{\\beta} \\Omega}{I_1} \\quad (13)\n$$\n\nwhere\n\n$$\nf_1 = a - \\delta_2\n$$\n\n$$\nf_2 = \\theta - \\dot{\\beta}/\\Omega\n$$\n\n$$\nf_3 = a \\dot{\\beta}/\\Omega + \\delta_1 + f_2 \\delta_2\n$$\n\nand $\\beta(t)$ is known approximately from equation (11).\n\nEquations (12) and (13) can be solved by a tabular, step-by-step process as outlined below.\n\nA sample calculation has been performed (fig. 1), by the step-by-step process, comparing the results obtained by including and neglecting the $\\dot{\\beta}$ terms in equations (12) and (13). Reference to the figure will show that the effect of flapping is negligible on the variations of $\\Omega(t)$", "timestamp": "2026-07-22T04:52:23.672926+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 1, "total_pages": 36, "image_filename": "19930085487_p1.jpg", "text": "NACA RM No. E8J22\nRM No. E8J22\n\n[Figure: NACA logo]\n\nRESEARCH MEMORANDUM\n\nVIBRATION SURVEY OF BLADES IN 10-STAGE\nAXIAL-FLOW COMPRESSOR\n\nI - STATIC INVESTIGATION\n\nBy André J. Meyer, Jr. and Howard F. Calvert\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\n\nJanuary 31, 1949\nDeclassified December 14, 1953", "timestamp": "2026-07-22T04:52:24.651207+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 104, "total_pages": 104, "image_filename": "19930085842_p104.jpg", "text": "100\nNACA RM L9C29\n\n<!-- Image (107, 109, 843, 855) -->\n\nFigure 48.- Variation of $HP_{req}$, $C_L$, $V/nD$, and $\\alpha$ with $V$ for level-flight conditions at $\\beta = 11.5^\\circ$. All control surfaces neutral; normal gross weight.", "timestamp": "2026-07-22T04:52:24.825528+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 20, "total_pages": 99, "image_filename": "19930082511_p20.jpg", "text": "18\nNACA TN No. 1826\n\nEssentially, then, the solutions will be developed for a vortex of strength $\\Gamma' = \\frac{\\Gamma}{hV}$ in a tunnel of unit height; and in the following list of symbols the lengths in the complex planes are in terms of $h$, and the complex velocities are in terms of $V$:\n\n| | |\n| :--- | :--- |\n| $h$ | tunnel height |\n| $V$ | tunnel velocity |\n| $\\zeta$ | complex variable of physical plane $(\\xi + i\\eta)$ |\n| $\\zeta_1$ | location of vortex in $\\zeta$-plane |\n| $z$ | complex variable of transformed plane $(x + iy)$ |\n| $z_1$ | location of vortex in $z$-plane |\n| $q$ | complex velocity in physical plane $(u - iv)$ |\n| $Q$ | complex velocity in transformed plane $(u - iv)$ |\n| $A, B, C, M, N$ | real constants |\n| $\\Gamma$ | vortex strength |\n| $\\Gamma^*$ | nondimensional vortex strength $\\left(\\frac{\\Gamma}{hV}\\right)$ |\n| $q_1(\\zeta, \\zeta_1)$ | induced complex velocity at $\\zeta$ when vortex is at $\\zeta_1$ |\n| $q_1(\\zeta_1)$ | induced complex velocity at $\\zeta_1$ |\n| $a$ | abscissa of exit lip in transformed space |\n| $w_1$ | a complex velocity in the form of an elliptic integral of the first kind |\n| $w_2$ | a complex velocity in the form of an elliptic integral of the second kind |\n| $w_3$ | a complex velocity |\n| $l$ | variable of integration |\n| $G$ | function defined by equation (2) |", "timestamp": "2026-07-22T04:52:26.920605+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 16, "total_pages": 49, "image_filename": "19930082498_p16.jpg", "text": "```markdown\nTABLE II.- RESULTS OF MUFFLER INVESTIGATION MADE WITH PROPELLER REMOVED - Continued\n\n| Muffler configuration | Engine speed (rpm) | Fundamental firing frequency, F (cps) | Sound-pressure level (db) | | | | | | | | | | | | | | | Other sounds | | | | Back pressure |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| (a) | | | Over-all | 0.2F | 1.0F | 1.5F | 2.0F | 2.5F | 3.0F | 3.5F | 4.0F | 4.5F | 5.0F | 5.5F | 6.0F | 6.5F | 7.0F | cps | db | cps | db | |\n| 13 [Figure: Diagram of muffler configuration 13 with dimensions 51, 34, 17 1/4, 1/8, 2 1/2, 6 1/2. (see fig. 1)] | 1650 | 82.5 | 83.0 | 48 | 83 | 60 | 81 | 59 | 62 | 68 | 67 | 55 | 55 | 53 | 50 | 50 | 50 | 684 | 47 | 707 | 47 | High |\n| | 2000 | 100.0 | 85.5 | 50 | 83 | 60 | 79 | 70 | 82 | 63 | 62 | 60 | 56 | 55 | 53 | 50 | 48 | --- | --- | --- | --- | |\n| | 2790 | 139.5 | 89.5 | 55 | 86 | 75 | 87 | 78 | 88 | 73 | 69 | 66 | 58 | 60 | (b) | (b) | (b) | --- | --- | --- | --- | |\n| 14 Same as 13 | 2000 | 100.0 | 88.5 | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | High |\n| | 2790 | 139.5 | 94.0 | 57 | 94 | 70 | 93 | 73 | 80 | 70 | 65 | 57 | 58 | 58 | (b) | (b) | (b) | --- | --- | --- | --- | |\n| 15 13 reversed | 1650 | 82.5 | 86.5 | 50 | 89 | 66 | 83 | 62 | 66 | 65 | 70 | 58 | 55 | 55 | 55 | 53 | 52 | --- | --- | --- | --- | High |\n| | 2000 | 100.0 | 84.5 | 50 | 84 | 58 | 73 | 66 | 65 | 67 | 55 | 55 | 55 | 53 | 55 | 55 | 48 | --- | --- | --- | --- | |\n| | 2790 | 139.5 | 92.0 | 58 | 90 | 62 | 78 | 62 | 65 | 63 | 60 | 60 | 55 | 55 | (b) | (b) | (b) | --- | --- | --- | --- | |\n| 16 [Figure: Diagram of muffler configuration 16 with dimensions 16 1/2, 35 1/2, 21, 15, 20, 18, 18, 3 3/4. (see fig. 1)] | * | | | | | | | | | | | | | | | | | | | | | |\n| | 2000 | 100.0 | 83.0 | 55 | 81 | 50 | 71 | 60 | 70 | 55 | 65 | 55 | 50 | 53 | 50 | (b) | (b) | 435 | 62 | --- | --- | Low |\n| 17 [Figure: Diagram of muffler configuration 17 with dimensions 21, 15, 20, 8, 18, 3 3/4.] | 1650 | 82.5 | 96.0 | 66 | 96 | 76 | 86 | 70 | 77 | 76 | 86 | 75 | 76 | 70 | 65 | (b) | (b) | --- | --- | --- | --- | Low |\n| | 2000 | 100.0 | 92.0 | 63 | 91 | 65 | 80 | 68 | 72 | 72 | 80 | 81 | 75 | 63 | 60 | (b) | (b) | --- | --- | --- | --- | |\n| | 2790 | 139.5 | 98.5 | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | |\n| 18 [Figure: Diagram of muffler configuration 18 with dimensions 10 1/2, 56, 10 1/2, 2 3/4. (see fig. 1)] | 1650 | 82.5 | 96.0 | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | Medium |\n| | 2000 | 100.0 | 94.0 | (b) | 94 | (b) | 86 | 75 | 78 | (b) | (b) | (b) | (b) | (b) | (b) | (b) | (b) | --- | --- | --- | --- | |\n| | 2790 | 139.5 | 100.5 | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | --- | |\n| 19 Steel wool [Figure: Diagram of muffler configuration 19 with dimensions 10, 12, 48, 8, 2 3/4. (see fig. 1)] | 2000 | 100.0 | 89.0 | 55 | 89 | (b) | 83 | (b) | 69 | (b) | 81 | (b) | 65 | (b) | (b) | (b) | (b) | --- | --- | --- | --- | Medium |\n| 20 Steel wool [Figure: Diagram of muffler configuration 20 with dimensions 16, 12, 48, 8, 2 3/4.] | 1650 | 82.5 | 84.0 | 62 | 84 | 60 | 68 | (b) | 68 | (b) | (b) | (b) | 60 | (b) | (b) | (b) | (b) | --- | --- | --- | --- | Medium |\n| | 2000 | 100.0 | 86.0 | 55 | 84 | 56 | 85 | 58 | 67 | 53 | 75 | 56 | 60 | 55 | 57 | (b) | (b) | --- | --- | --- | --- | |\n\n$^a$In the sketches of the configurations, all dimensions are in inches and all cross sections are circular, except where otherwise indicated.\n$^b$The sound-pressure level was below the range of the analyzer.\n\nNACA\nNACA TN No. 1838\n15\n```", "timestamp": "2026-07-22T04:52:28.566016+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 3, "total_pages": 28, "image_filename": "19930085471_p3.jpg", "text": "NACA RM No. L8J11\n\nUNCLASSIFIED\nCONFIDENTIAL\nRESTRICTED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nINITIAL EXPERIMENTS ON FLUTTER OF UNSWEPT\nCANTILEVER WINGS AT MACH NUMBER 1.3\n\nBy W. J. Tuovila, John E. Baker, and Arthur A. Regier\n\nSUMMARY\n\nA supersonic tunnel designed to operate at Mach number 1.3 was used for a preliminary experimental flutter investigation of widely different unswept cantilever wings. Data for 12 wings with mass-density parameters $1/k$ ranging from 52 to 268, center-of-gravity positions ranging from 46 to 63 percent chord from the leading edge, and elastic-axis positions ranging from 34 to 52 percent chord from the leading edge are considered.\n\nA comparison is made of the test results with calculations of bending-torsion flutter obtained by the theory of flutter in supersonic two-dimensional flow and it is concluded that the test data are in reasonable agreement with the calculated results. In general, the theoretical values are conservative. As shown by the theory, the flutter results are quite sensitive to the location of the center of gravity. Thick and thin, blunt and sharp airfoil-section shapes were used, but no very pronounced effect of the section shape on flutter characteristics was found. The data suggest that for cantilever wings the bending degree of freedom may suppress the one-degree-of-freedom torsional flutter and that coupled bending-torsion flutter effects occur. The experiments include a study of the effect of the addition of tip moments of inertia. With the center of gravity of the tip weights coincident with the center of gravity of the wing section no detrimental effect on the flutter speed was found.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| b | semichord, feet |\n| $c_w$ | chord, feet |\n| GJ | torsional stiffness, inches$^2$-pounds |\n\n[Stamp: Classification Changed to RESTRICTED, NACA Res. Div. #46d-712/13, 46d-312, 8.3.11-31, d-112153, Date 1/4/54, By J.E. Newlan, NACA]\n\n[Stamp: Classification Changed to UNCLASSIFIED, NACA, Authority Jan. 13, 1954, Research Abstract '56, Date FEB 24 1954, I.E. Newlan]\n\nUNCLASSIFIED\nRESTRICTED\nCONFIDENTIAL\nUNCLASSIFIED", "timestamp": "2026-07-22T04:52:34.704957+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 32, "total_pages": 47, "image_filename": "19930093773_p32.jpg", "text": "NACA RM E9G09\n31\n\n[Figure: A line graph plotting Fuel flow against Engine speed. The graph contains a legend box and a NACA logo in the bottom right corner.]\n\n| | Flight Mach number |\n| :--- | :--- |\n| O | 0.21 |\n| $\\square$ | .53 |\n| $\\diamond$ | .72 |\n| $\\triangle$ | .85 |\n| $\\nabla$ | .97 |\n\nFuel flow, $W_f$, lb/hr\n\nEngine speed, N, rpm\n\n(c) Fuel flow.\n\nFigure 5. - Continued. Effect of flight Mach number on variation of engine performance with engine speed at altitude of 25,000 feet.", "timestamp": "2026-07-22T04:52:34.899972+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 35, "total_pages": 66, "image_filename": "19930082245_p35.jpg", "text": "```markdown\n34\n\n1.8\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nAileron section normal-force coefficient, $C_{n\\alpha}$\n\n$\\delta_a$\n(deg)\n18\n12\n4\n2\n0\n-2\n-4\n-6\n-12\n\n.20\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.28\n-.32\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nAileron section hinge-moment coefficient, $c_h$\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n\n(b) $c_n = -0.2$.\nFigure 7 - Continued.\n\nNACA\nNACA TN No. 1596\n```", "timestamp": "2026-07-22T04:52:39.946234+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 6, "total_pages": 66, "image_filename": "19930082914_p6.jpg", "text": "NACA TN No. 1857\n\nbe met that the two beams be widely enough separated in space that the disturbance to be studied can be introduced into one of the beams without disturbing the other. The disturbance, or the \"test section,\" can, of course, be located anywhere in either of the two beams. In the apparatus described in the present paper the test section was situated midway between the mirror $M_2$ and the splitter plate $S_2$.\n\nTheory of ideal fringe formation.- The two beams of light that reach the photographic plate P appear to come from separate sources that are situated somewhere to the left of the mirror $M_2$. By proper orientation of the two splitter plates and the two mirrors, the two beams of light can be made to appear to cross each other, as is shown in figure 3. If, for the moment, it is assumed that each beam is composed of strictly parallel and monochromatic rays, then the wave fronts can be represented by equally spaced straight lines perpendicular to the direction of propagation, as is also shown in figure 3. Here the straight lines represent the \"crests\" of the waves. Midway between two successive \"crests\" are the \"troughs\" of the waves. Where two lines intersect, two crests occupy the same position in space, reinforce each other, and cause an increase in the amplitude of vibration and an increase in the intensity of the light. Where a line intersects a point that is midway between two adjacent lines of the other beam, a crest and a trough interfere destructively, so that a decrease in the amplitude of motion and a decrease in the intensity of the light results. The plate P will therefore be crossed by parallel, horizontal (in this case) lines of alternately weak and strong intensity. These lines are the interference fringes. They can be oriented in any direction by proper rotation of the two splitter plates and the two mirrors about two axes, one in a plane parallel to that of the paper and one perpendicular to the plane of the paper.\n\nIf now a disturbance is produced in the air that is in the test section, and this disturbance changes the speed of the light in the beam that actually traverses the test section, then the wave fronts in that beam will be advanced or retarded, and the positions of the fringes will be shifted down or up. From the shift in the fringes the average speed of the light and, from that, the average density of the air can be calculated. The only gas-flow cases for which the density field can be evaluated are the one-dimensional case of uniform density throughout the field, the two-dimensional case of uniform densities along lines (or planes) that lie parallel to the light-beam direction of propagation, and the three-dimensional case of axially symmetric densities, such as the flow field about bodies of revolution at zero angle of attack.\n\nTheory of practical fringe formation.- So far it has been assumed that the light is strictly monochromatic and is a beam of parallel rays. But parallel light can be obtained only from a truly point source. Because in practice neither monochromatic light nor a point source is used (because, with available means of approximating these to a very high degree, the intensity of the light would be too small), neither condition", "timestamp": "2026-07-22T04:52:42.303576+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 116, "total_pages": 118, "image_filename": "19930085838_p116.jpg", "text": "114\nNACA RM No. L9B23\n\n<!-- Image (132, 110, 743, 894) -->\n\nFigure 21.- Variation of straight-sided Frise aileron hinge-moment characteristics with flap deflection on two NACA 7-series-type airfoils with flap and double slotted flap. $c_a = 0^\\circ$; $\\delta_a = 0^\\circ$; R = 6.0 x $10^6$ (approx.).", "timestamp": "2026-07-22T04:52:43.391195+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 18, "total_pages": 37, "image_filename": "19930082450_p18.jpg", "text": "NACA TN No. 1778\n17\n\nTABLE 9.- Z-FAMILY PROPERTIES - Concluded $\\left[\\frac{W_{\\mathrm{e}}}{W_{\\mathrm{a}}}=1.00 ; \\frac{S_{\\mathrm{e}}}{S_{\\mathrm{a}}}=8.6 ; \\frac{Z_{\\mathrm{e}}}{Z_{\\mathrm{a}}}=0.4 ; \\frac{I_{\\mathrm{e}}}{I_{\\mathrm{a}}}=3 ; \\frac{t_{\\mathrm{e}}}{t_{\\mathrm{a}}}=\\frac{1}{4} ; \\frac{d_{\\mathrm{e}}}{d_{\\mathrm{a}}}=1.95 ; \\frac{h_{\\mathrm{e}}}{h_{\\mathrm{a}}}=11.7\\right]$\n\n| $\\frac{h_{\\mathrm{e}}}{h_{\\mathrm{a}}}$ | $\\frac{d_{\\mathrm{e}}}{d_{\\mathrm{a}}}$ | 33 | 34 | 35 | 36 | 37 | 38 | 39 | 40 | 41 | 42 | 43 | 44 | 45 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 25 | | 2.955 | 3.111 | 3.167 | 3.223 | 3.279 | 3.335 | 3.391 | 3.447 | 3.503 | 3.559 | 3.615 | 3.671 | 3.727 |\n| 26 | | 2.976 | 3.030 | 3.084 | 3.137 | 3.191 | 3.245 | 3.299 | 3.353 | 3.407 | 3.460 | 3.514 | 3.568 | 3.622 |\n| 27 | | 2.993 | 3.044 | 3.094 | 3.144 | 3.194 | 3.244 | 3.294 | 3.344 | 3.394 | 3.444 | 3.494 | 3.544 | 3.594 |\n| 28 | | 2.895 | 2.985 | 3.035 | 3.085 | 3.135 | 3.185 | 3.235 | 3.285 | 3.335 | 3.385 | 3.435 | 3.485 | 3.535 |\n| 29 | | 2.773 | 2.850 | 2.892 | 2.935 | 2.977 | 3.019 | 3.061 | 3.103 | 3.145 | 3.187 | 3.229 | 3.271 | 3.313 |\n| 30 | | 2.713 | 2.780 | 2.820 | 2.860 | 2.900 | 2.940 | 2.980 | 3.020 | 3.060 | 3.100 | 3.140 | 3.180 | 3.220 |\n| 31 | | 2.656 | 2.702 | 2.747 | 2.792 | 2.838 | 2.883 | 2.928 | 2.973 | 3.018 | 3.063 | 3.109 | 3.154 | 3.199 |\n| 32 | | 2.600 | 2.646 | 2.690 | 2.734 | 2.778 | 2.822 | 2.866 | 2.910 | 2.954 | 2.998 | 3.042 | 3.086 | 3.130 |\n| 33 | | 2.557 | 2.599 | 2.642 | 2.684 | 2.727 | 2.769 | 2.811 | 2.854 | 2.896 | 2.938 | 2.981 | 3.023 | 3.066 |\n| 34 | | 2.514 | 2.556 | 2.597 | 2.639 | 2.680 | 2.721 | 2.763 | 2.804 | 2.846 | 2.887 | 2.928 | 2.970 | 3.011 |\n| 35 | | 2.478 | 2.508 | 2.548 | 2.588 | 2.628 | 2.668 | 2.708 | 2.748 | 2.788 | 2.828 | 2.868 | 2.908 | 2.948 |\n| 36 | | 2.422 | 2.462 | 2.502 | 2.542 | 2.582 | 2.622 | 2.662 | 2.702 | 2.742 | 2.782 | 2.822 | 2.862 | 2.902 |\n| 37 | | 2.388 | 2.426 | 2.464 | 2.502 | 2.540 | 2.578 | 2.616 | 2.654 | 2.692 | 2.730 | 2.768 | 2.806 | 2.844 |\n| 38 | | 2.352 | 2.389 | 2.426 | 2.463 | 2.500 | 2.537 | 2.574 | 2.611 | 2.648 | 2.685 | 2.722 | 2.759 | 2.796 |\n| 39 | | 2.317 | 2.353 | 2.389 | 2.425 | 2.461 | 2.497 | 2.533 | 2.569 | 2.605 | 2.641 | 2.677 | 2.713 | 2.749 |\n| 40 | | 2.282 | 2.317 | 2.352 | 2.387 | 2.422 | 2.457 | 2.492 | 2.527 | 2.562 | 2.597 | 2.632 | 2.667 | 2.702 |\n| 41 | | 2.228 | 2.262 | 2.296 | 2.330 | 2.364 | 2.398 | 2.432 | 2.466 | 2.500 | 2.534 | 2.568 | 2.602 | 2.636 |\n| 42 | | 2.217 | 2.250 | 2.283 | 2.316 | 2.349 | 2.382 | 2.415 | 2.448 | 2.481 | 2.514 | 2.547 | 2.580 | 2.613 |\n| 43 | | 2.168 | 2.199 | 2.231 | 2.263 | 2.295 | 2.327 | 2.359 | 2.391 | 2.423 | 2.455 | 2.487 | 2.519 | 2.551 |\n| 44 | | 2.117 | 2.147 | 2.178 | 2.208 | 2.239 | 2.269 | 2.300 | 2.330 | 2.361 | 2.391 | 2.421 | 2.452 | 2.482 |\n| 45 | | 2.070 | 2.099 | 2.128 | 2.157 | 2.186 | 2.215 | 2.244 | 2.273 | 2.302 | 2.331 | 2.360 | 2.389 | 2.418 |\n| 46 | | 2.027 | 2.055 | 2.083 | 2.111 | 2.139 | 2.167 | 2.195 | 2.223 | 2.251 | 2.279 | 2.307 | 2.335 | 2.363 |\n| 47 | | 1.988 | 2.015 | 2.042 | 2.069 | 2.096 | 2.123 | 2.150 | 2.177 | 2.204 | 2.231 | 2.258 | 2.285 | 2.312 |\n| 48 | | 1.951 | 1.977 | 2.003 | 2.029 | 2.055 | 2.081 | 2.107 | 2.133 | 2.159 | 2.185 | 2.211 | 2.237 | 2.263 |\n| 49 | | 1.917 | 1.942 | 1.967 | 1.992 | 2.017 | 2.042 | 2.067 | 2.092 | 2.117 | 2.142 | 2.167 | 2.192 | 2.217 |\n| 50 | | 1.885 | 1.909 | 1.933 | 1.957 | 1.981 | 2.005 | 2.029 | 2.053 | 2.077 | 2.101 | 2.125 | 2.149 | 2.173 |\n| 51 | | 1.857 | 1.879 | 1.902 | 1.925 | 1.948 | 1.971 | 1.994 | 2.017 | 2.040 | 2.063 | 2.086 | 2.109 | 2.132 |\n| 52 | | 1.791 | 1.812 | 1.834 | 1.855 | 1.877 | 1.898 | 1.920 | 1.941 | 1.963 | 1.984 | 2.006 | 2.027 | 2.049 |\n| 53 | | 1.758 | 1.778 | 1.798 | 1.818 | 1.838 | 1.858 | 1.878 | 1.898 | 1.918 | 1.938 | 1.958 | 1.978 | 1.998 |\n| 54 | | 1.685 | 1.704 | 1.723 | 1.742 | 1.761 | 1.780 | 1.799 | 1.818 | 1.837 | 1.856 | 1.875 | 1.894 | 1.913 |\n| 55 | | 13.16 | 13.70 | 14.24 | 14.78 | 15.32 | 15.86 | 16.40 | 16.94 | 17.48 | 18.02 | 18.56 | 19.10 | 19.64 |\n| 56 | | 12.93 | 13.45 | 13.97 | 14.49 | 15.01 | 15.53 | 16.05 | 16.57 | 17.09 | 17.61 | 18.13 | 18.65 | 19.17 |\n| 57 | | 12.81 | 13.32 | 13.83 | 14.34 | 14.85 | 15.36 | 15.87 | 16.38 | 16.89 | 17.40 | 17.91 | 18.42 | 18.93 |\n| 58 | | 12.68 | 13.17 | 13.66 | 14.15 | 14.64 | 15.13 | 15.62 | 16.11 | 16.60 | 17.09 | 17.58 | 18.07 | 18.56 |\n| 59 | | 12.58 | 13.05 | 13.52 | 13.99 | 14.46 | 14.93 | 15.40 | 15.87 | 16.34 | 16.81 | 17.28 | 17.75 | 18.22 |\n| 60 | | 12.48 | 12.93 | 13.38 | 13.83 | 14.28 | 14.73 | 15.18 | 15.63 | 16.08 | 16.53 | 16.98 | 17.43 | 17.88 |\n| 61 | | 12.20 | 12.63 | 13.06 | 13.49 | 13.92 | 14.35 | 14.78 | 15.21 | 15.64 | 16.07 | 16.50 | 16.93 | 17.36 |\n| 62 | | 12.10 | 12.51 | 12.92 | 13.33 | 13.74 | 14.15 | 14.56 | 14.97 | 15.38 | 15.79 | 16.20 | 16.61 | 17.02 |\n| 63 | | 12.00 | 12.40 | 12.80 | 13.20 |", "timestamp": "2026-07-22T04:52:51.090061+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 2, "total_pages": 36, "image_filename": "19930085487_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:52:53.990115+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 12, "total_pages": 53, "image_filename": "19930082542_p12.jpg", "text": "NACA-TN No. 1867\n\nThe total range in properties produced by the treatments was very wide. Yield strengths at room temperature ranged from 30,000 to 134,000 psi at 0.02-percent offset. The stress for rupture in 100 hours at $1200^\\circ$ F varied from 40,000 to 66,000 psi and in 1000 hours from 35,000 to 56,000 psi. In terms of time for fracture these results mean that in its weakest condition the alloy would fracture at $1200^\\circ$ F under a stress of 40,000 psi in 100 hours. In the strongest condition the alloy could carry the 40,000 psi for an estimated 600,000 hours. Under a stress of 56,000 psi the strongest condition would fracture in 1000 hours at $1200^\\circ$ F, whereas the weakest condition would fail in less than 0.1 hour. Ductility in the rupture test also varied over wide limits. The percentage elongation for fracture in 100 hours ranged from 1 to 40 percent.\n\nThese wide variations are significant for several reasons. The most important is that processing and heat treatment must be controlled to obtain the highest strengths possible from the alloy. Rather wide variations in properties for several heats have been due to a lack of suitable control of the processing conditions. The absence of systematic correlations between chemical composition and properties for a wide range of alloys is believed due to uncontrolled comparative processing and to the possibility that optimum properties may require different treatments for each alloy. The wide range in properties which has been observed for any one alloy in the hot-worked condition quite certainly is due to the effects shown for heat treatment and hot-cold-work. During hot-working an alloy is heat-treated and hot-cold-worked simultaneously. Therefore a range in properties is to be expected depending on the conditions of hot-working.\n\n### Limitations of Data\n\nThere are several limitations to the general applicability of the data. One of the most important is that only one heat and lot of bar stock has been studied. Information is not available regarding the influence of melting and hot-working variables on the results obtained. The data shown for hot-worked material are considered to have the least general applicability because hot-worked stock could conceivably vary from a quite thoroughly solution-treated condition to a severely cold-worked and agglomerated material. Those experiments which involved a fairly thorough solution treatment should have given typical data for any heat. Further work on other heats is needed, however, to demonstrate the degree of reproducibility between heats.\n\nAll comparisons have been based on room-temperature properties and rupture test characteristics at $1200^\\circ$ F. Caution is needed in applying the results of the rupture tests to applications involving limited permissible deformation. There is reason to believe that the relative effects of the different treatments would vary considerably", "timestamp": "2026-07-22T04:52:55.034638+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 20, "total_pages": 50, "image_filename": "19930082496_p20.jpg", "text": "NACA TN No. 1836\n19\n\nTABLE III - PHASE 1 OF QUASI-SERVICE TURBINE-BLADE EVALUATION\n\n| Run | Operating time | | Cumulative operating time | | Shaft speed (rpm) | Tip speed (ft/sec) | Indicated inlet gas temperature (°F) | Metal blades cracked and replaced | Metal blades completely fractured | Total metal blades replaced | Cumulative total metal blades replaced | Ceramic blades fractured | Cumulative total ceramic blades fractured |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | (hr) | (min) | (hr) | (min) | | | | | | | | | |\n| 1 | 1 | 30 | 1 | 30 | 10,000 | 478 | 1750 | 0 | 0 | 0 | 0 | 0 | 0 |\n| | | 30 | 1 | 1 | 10,000 | 478 | 1800 | 0 | 0 | 0 | 0 | 0 | 0 |\n| | | 30 | 2 | 30 | 10,000 | 478 | 1850 | 0 | 0 | 0 | 0 | 0 | 0 |\n| | | 30 | 2 | 2 | 10,000 | 478 | 1900 | 0 | 0 | 0 | 0 | 0 | 0 |\n| | | 30 | 3 | 30 | 10,000 | 478 | 1950 | 0 | 0 | 0 | 0 | 0 | 0 |\n| | | 30 | 3 | 4 | 10,000 | 478 | 2000 | 0 | 0 | 0 | 0 | 0 | 0 |\n| | | 30 | 4 | 30 | 12,500 | 596 | 2000 | 0 | 0 | 0 | 0 | 0 | 0 |\n| | | 30 | 4 | 4 | 15,000 | 716 | 2000 | 0 | 0 | 0 | 0 | 0 | 0 |\n| | | 10 | 4 | 40 | 17,500 | 835 | 2000 | 0 | 0 | 0 | 0 | 1 | 1 |\n| 2 | | 5 | 4 | 45 | 17,500 | 835 | 2000 | 0 | 0 | 0 | 0 | 2 | 3 |\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T04:52:56.730715+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 10, "total_pages": 78, "image_filename": "19930082618_p10.jpg", "text": "8\nNACA TN 1945\n\nDrag\n\nThe general form of the drag polars corresponding to the various Reynolds numbers may be seen in figures 1 to 15. The principal effects on the drag of decreasing the Reynolds number from $9.0 \\times 10^6$ to $0.7 \\times 10^6$ appear to be a variation in width of the flat portion of the polars, an increase in value of the minimum drag coefficient, and a steepening of the drag curves beyond the flat portion of the polars.\n\nLow-drag range.- The extent of the lift-coefficient range over which the 10 NACA 6-series airfoils in the smooth condition have low drag, which corresponds to extensive laminar flow, generally increases as the Reynolds number is lowered, with the greatest increase usually occurring as the Reynolds number is lowered below $3.0 \\times 10^6$ (figs. 1 to 15). The magnitude of the effect is greatest for the airfoils of greatest thickness, highest design lift coefficient, and farthest rearward position of minimum pressure. It is of interest to note that the actual low-drag range is considerably greater than the theoretical low-drag range for all the airfoils at the lower Reynolds numbers. Hence, for these Reynolds numbers, the first small pressure peaks which form near the leading edge as the lift coefficient is increased do not cause transition from laminar to turbulent flow. For most Reynolds numbers, the data show that thick airfoils with the position of minimum pressure far forward tend to have the widest low-drag ranges.\n\nThe increase in drag with increasing lift coefficient within the low-drag region shown for some of the airfoils at the lower Reynolds numbers (particularly pronounced for the NACA 642-415 section, fig. 3(o)) is believed to be associated with the formation of a laminar separation bubble behind the position of minimum pressure. The exact behavior of this bubble as the lift coefficient and Reynolds number are varied, however, is not entirely clear at the present time.\n\nAlthough the comparatively high values of the minimum drag coefficient shown by the five NACA 4- and 5-digit-series airfoil sections (figs. 11 to 15) preclude the possibility of a low-drag range corresponding to extensive laminar layers on the airfoil surfaces, there is a range of lift coefficient through which the drag of these airfoils changes very little. Although the manner in which this range varies with Reynolds number for the different airfoils is not very well defined, there does seem to be a tendency, which is especially marked in the cases of the NACA 4412 and NACA 23012 airfoil sections, toward a decrease in the extent of this range as the Reynolds number is decreased. This effect is believed to be associated with the formation and behavior of a laminar separation bubble a short distance behind the leading edge on the suction side of the airfoil. As previously stated, however, the details of the mechanics of the laminar separation bubble are not completely understood.", "timestamp": "2026-07-22T04:52:57.988639+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 1, "total_pages": 46, "image_filename": "19930085519_p1.jpg", "text": "NACA RM No. L8K19\n\nRESTRICTED\n\nCopy No. 187\nRM No. L8K19\n\nNACA\nCASE FILE\nCOPY\n\nRESEARCH MEMORANDUM\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nAUTHORITY CROWLEY CHANGE #2031\nDATE 12-14-53\nT.C.F.\n\nWIND-TUNNEL INVESTIGATION AT LOW SPEEDS OF VARIOUS\nPLUG-AILERON AND LIFT-FLAP CONFIGURATIONS\nON A 42° SWEPTBACK SEMISPAN WING\n\nBy\nLeslie E. Schneiter and James M. Watson\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nJanuary 26, 1949\n\nRESTRICTED", "timestamp": "2026-07-22T04:52:59.151139+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 28, "total_pages": 62, "image_filename": "19930082485_p28.jpg", "text": "NACA TN No. 1810\n27\n\nTABLE I - BLADE COORDINATES\n\n[Figure: Diagram showing Section, Outer shroud, Radius (in.), Tip, Pitch, Root, h, chord, Axial depth, Circumferential depth, $X_1$, $X_2$]\n\nShaded areas are boundary layer\n\n| h (in.) | Root section | | Pitch section | | Tip section | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | $X_1$ (in.) | $X_2$ (in.) | $X_1$ (in.) | $X_2$ (in.) | $X_1$ (in.) | $X_2$ (in.) |\n| 0 | 0.063 | -0.053 | 0.063 | -0.053 | 0.063 | -0.053 |\n| .100 | .114 | - .068 | .114 | - .068 | .114 | - .068 |\n| .200 | .148 | - .083 | .148 | - .083 | .148 | - .083 |\n| .300 | .168 | - .110 | .167 | - .105 | .165 | - .100 |\n| .400 | .178 | - .138 | .173 | - .130 | .167 | - .122 |\n| .500 | .172 | - .171 | .164 | - .161 | .156 | - .151 |\n| .600 | .152 | - .211 | .143 | - .199 | .134 | - .187 |\n| .700 | .117 | - .258 | .107 | - .234 | .096 | - .229 |\n| .800 | .064 | - .316 | .053 | - .299 | .041 | - .281 |\n| .900 | - .012 | - .388 | - .016 | - .365 | - .029 | - .342 |\n| 1.000 | - .116 | - .471 | - .115 | - .442 | - .114 | - .412 |\n| 1.100 | - .245 | - .569 | - .229 | - .531 | - .212 | - .493 |\n| 1.200 | - .407 | - .688 | - .369 | - .638 | - .331 | - .589 |\n| 1.300 | - .612 | - .835 | - .542 | - .769 | - .471 | - .703 |\n| 1.400 | - .869 | -1.022 | - .754 | - .930 | - .638 | - .838 |\n| 1.500 | -1.176 | -1.244 | -1.008 | -1.126 | - .840 | -1.001 |\n| 1.600 | | | | | -1.101 | -1.178 |\n\n| Section | Pitch (in.) | Chord (in.) | Axial depth (in.) | Circumferential depth (in.) | Leading-edge radius (in.) | Trailing-edge radius (in.) |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Root | 1.193 | 2.052 | 1.535 | 1.275 | 0.050 | 0.010 |\n| Mean | 1.334 | 2.074 | 1.585 | 1.245 | .050 | .010 |\n| Tip | 1.475 | 2.100 | 1.638 | 1.216 | .050 | .010 |\n\nNACA", "timestamp": "2026-07-22T04:52:59.368713+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 11, "total_pages": 58, "image_filename": "19930082617_p11.jpg", "text": "10\nNACA TN 1962\n\nTABLE III\nBEHAVIOR AFTER COLLAPSE\n\n| Cylinder | Failed at | | After collapse | | Second failure after overnight rest | | After collapse | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | Load on jack | Moment | Load on jack | Moment | Load on jack | Moment | Load on jack | Moment |\n| 72 | 2992 | 212,544 | 2730 | 196,560 | 2800 | 201,600 | 2935 | 182,920 |\n| 73 | 3787 | 272,664 | 3579 | 257,688 | 3996 | 257,712 | 3352 | 241,344 |\n| 74 | 4489 | 323,208 | 3921 | 282,312 | 3571 | 257,112 | 3401 | 244,872 |\n| 75 | 4300 | 309,600 | 3997 | 287,784 | 5189 | 373,608 | 5019 | 361,368 |\n| 76 | 4500 | 324,000 | 2607 | 187,704 | 2916 | 209,952 | 2644 | 190,368 |\n| 77 | 4260 | 306,720 | 3560 | 256,320 | 3693 | 265,896 | 3091 | 219,672 |\n| a78 | 6267 | 451,224 | | | | | | |\n| 79 | 5150 | 370,800 | 3404 | 173,088 | 2840 | 204,480 | 2462 | 177,264 |\n\naCylinder 78 failed in tension.", "timestamp": "2026-07-22T04:52:59.763896+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 12, "total_pages": 30, "image_filename": "19930082585_p12.jpg", "text": "NACA TN 1907\n\nand $V(t)$. The flapping terms are therefore neglected in the following development of the step-by-step method of solving equations (12) and (13). In effect, then, only rigid blades are considered, although the effects of blade flapping would be negligible.\n\nEquations (12) and (13) then become\n\n$$\n\\dot{V} = g - \\frac{\\rho a b c R^3}{2W/g} \\left( \\frac{V - v}{2R} \\Omega + \\frac{\\theta}{3} \\Omega^2 \\right)\n$$\n\nand\n\n$$\n\\dot{\\Omega} = \\frac{\\rho c R^4}{2I_1} \\left[ \\frac{a - \\delta_2}{2R^2} (V - v)^2 - \\frac{\\delta_1 - \\theta(a - 2\\delta_2)}{3R} \\Omega(V - v) - \\frac{\\delta_0' + \\delta_1 \\theta + \\delta_2 \\theta^2}{4} \\Omega^2 \\right]\n$$\n\nBy differentiating equation (13a),\n\n$$\n\\ddot{\\Omega} = \\frac{\\rho c R^4}{2I_1} \\left\\{ \\frac{a - \\delta_2}{R^2} (V - v)(\\dot{V} - \\dot{v}) - \\frac{\\delta_1 - \\theta(a - \\delta_2)}{3R} \\left[ \\Omega(\\dot{V} - \\dot{v}) + \\dot{\\Omega}(V - v) \\right] + \\frac{\\dot{\\theta}(a - 2\\delta_2)}{3R} \\Omega(V - v) - \\frac{\\delta_0' + \\delta_1 \\theta + \\delta_2 \\theta^2}{2} \\Omega \\dot{\\Omega} - \\frac{\\dot{\\theta}}{4} (\\delta_1 + 2\\theta \\delta_2) \\Omega^2 \\right\\}\n$$", "timestamp": "2026-07-22T04:53:00.687590+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 21, "total_pages": 99, "image_filename": "19930082511_p21.jpg", "text": "NACA TN No. 1826\n19\n\nK, K'\ncomplete elliptic integrals of the first kind, with modulus 1/a\n\nE, E'\nwhen not followed by parenthesis, complete elliptic integrals of the second kind, with modulus 1/a; with upper limit indicated in parentheses, incomplete elliptic integrals of the second kind, with modulus 1/a\n\nF, F'\nincomplete elliptic integrals of the first kind, with modulus 1/a, and with upper limit indicated in parentheses\n\nR.P.\nreal part\n\nI.P.\nimaginary part\n\nc\nairfoil chord\n\n$c_l$\nairfoil lift coefficient\n\n$\\epsilon$\ntunnel-induced angle, radians\n\n$u_b$\nhorizontal perturbation velocity at free boundary\n\nSubscript\n\ni\ninduced\n\nBOUNDARY CONDITIONS\n\nThe two-dimensional tunnels discussed are considered to have their fixed and free boundaries parallel to the real axis, with the main tunnel flow from left to right. The physical plane (in which lengths and velocities have been made nondimensional as just described) will be designated the $\\zeta$-plane, with the complex perturbation velocity $u - iv$, or $q(\\zeta)$, subject to the following conditions:\n\n(1) on each fixed (or closed) boundary, $I.P. q(\\zeta) \\equiv -v = 0$\n\n(2) On each free (or open) boundary, $R.P. q(\\zeta) \\equiv u = 0$ or a constant\n\n(3) At each lip of the closed entrance section, $q(\\zeta)$ is continuous", "timestamp": "2026-07-22T04:53:02.439812+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 36, "total_pages": 66, "image_filename": "19930082245_p36.jpg", "text": "```markdown\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nAileron section normal-force coefficient, $C_{n_a}$\n\n$\\delta_a$\n(deg)\n30\n18\n12\n4\n2\n0\n-2\n-4\n-6\n-12\n\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n-.24\n-.56\n-.60\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\nAileron section hinge-moment coefficient, $C_h$\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n12\n18\n30\n\nNACA\n\n(c) $c_n = 0$.\nFigure 7 - Continued.\n\nNACA TN No. 1596\n35\n```", "timestamp": "2026-07-22T04:53:04.924420+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 33, "total_pages": 47, "image_filename": "19930093773_p33.jpg", "text": "32\nNACA RM E9G09\n\n[Figure: A line graph plotting specific fuel consumption against engine speed. The graph contains multiple curves representing different flight Mach numbers, indicated by symbols in a legend box. The NACA logo is present in the bottom right corner of the plot area.]\n\nSpecific fuel consumption based on net thrust, $W_f/F_n$\nlb/(hr)/(lb thrust)\n\n| Flight Mach number | |\n| :--- | :--- |\n| $\\circ$ | 0.21 |\n| $\\square$ | .53 |\n| $\\diamond$ | .72 |\n| $\\triangle$ | .85 |\n| $\\nabla$ | .97 |\n\nEngine speed, N, rpm\n\n(d) Specific fuel consumption.\nFigure 5. - Continued. Effect of flight Mach number on variation of engine performance with engine speed at altitude of 25,000 feet.", "timestamp": "2026-07-22T04:53:05.557448+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 117, "total_pages": 118, "image_filename": "19930085838_p117.jpg", "text": "NACA RM No. L9B23\n115\n\n<!-- Image (166, 108, 874, 433) -->\n\n(a) 0.177c thick airfoil.\n\n$\\delta_f$ $\\delta_b$\n$\\square$ $0^\\circ$ $0.408c_a$\n$\\square$ $25^\\circ$ $.351c_a$\n$\\square$ $25^\\circ$ $.408c_a$\n$\\diamond$ $40^\\circ$ $.351c_a$\n$\\diamond$ $40^\\circ$ $.408c_a$\n\n<!-- Image (166, 571, 874, 878) -->\n\n(b) 0.154c thick airfoil.\n\nFigure 22.- Effect of aileron deflection on flap hinge-moment characteristics on two NACA 7-series-type airfoils with double slotted flap and straight-sided Frise aileron. $\\delta_o = 0^\\circ$; $\\delta_f = 0^\\circ$; $R = 6.0 \\times 10^6$ (approx.).", "timestamp": "2026-07-22T04:53:11.117584+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 19, "total_pages": 37, "image_filename": "19930082450_p19.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:53:14.097099+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 7, "total_pages": 66, "image_filename": "19930082914_p7.jpg", "text": "6\nNACA TN No. 1857\n\nis actually met. As has been pointed out before (reference 10), the complete theory of fringes formation for the actual case has not yet been given. Schardin, however, has shown (reference 11) that the following is the case. With a point source and a strictly parallel beam, fringes are formed at all points along the two axes $I_1 - I_1'$ and $I_2 - I_2'$, where the two beams overlap. (See fig. 4.) But in an actual case, in which a light source of finite extent is used, the lines $I_1 - I_1'$ and $I_2 - I_2'$ become very numerous. The most distinct fringes are then formed where the various lines between $I_1$ and $I_1'$ and between $I_2$ and $I_2'$ intersect. The interferometer should be so adjusted, by rotation of the splitter plates, that this intersection is in the center of the test section.\n\nFor obtaining the greatest contrast between fringes, it is also necessary that the optical-path length through the interferometer be the same for the two beams. In other words, it is necessary that the optical path for the two beams be nearly the same from the time the original beam is split into two, at the splitter plate $S_1$, until the two are reunited at the splitter plate $S_2$. This condition becomes the more important the more the light departs from being monochromatic. For a given setting of the splitter plates and the mirrors, the fringes have a given spacing for each wavelength. The greater the wavelength, the greater the spacing between wavefronts and consequently between fringes. Thus, fringes produced by a large number of wavelengths coincide only at one point, the center of the band of fringes. On either side of this center the fringes get out of step with each other and the contrast between light and dark becomes less and less. With the optical paths equal, the center of the band of fringes, where the contrast is greatest, lies at the center of the test section.\n\nDescription of interferometer.- The interferometer with which the results reported in the present paper were obtained was designed and constructed at the Langley Laboratory and is installed in the boundary layer Laboratory of the Physical Research Division. The base on which the splitter plates and mirrors are mounted is a one-piece iron casting in the form of a four-leaf clover. (See figs. 5 and 6.) The assembly is supported in a vertical plane at its center by a single mount, which is attached to a framework of structural steel that is welded to a steel table. The table rests on steel plates that are bolted to the concrete floor of the second story of the building. The building happens to house numerous motors, compressors, and other sources of vibration. It is within 50 feet of a projectile gallery, a 500-horsepower wind tunnel, and a 1000-horsepower wind tunnel. All of these pieces of equipment cause vibrations of considerable amplitude in the second-story floor, which is supported by columns only every 20 feet or so.", "timestamp": "2026-07-22T04:53:16.904112+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 4, "total_pages": 28, "image_filename": "19930085471_p4.jpg", "text": "2\nUNCLASSIFIED\nRE-CONFIDENTIAL\nNACA RM No. L8J11\n\n$\\rho$ mass density of air in test section\nM mass of wing per unit span\n$1/\\kappa$ mass-density parameter $(M/\\pi\\rho b^2)$\n$I_\\alpha$ mass moment of inertia of wing about elastic axis per unit span\n$r_\\alpha^2$ radius-of-gyration parameter $(I_\\alpha/Mb^2)$\nV flutter velocity, feet per second\n$f_f$ flutter frequency, cycles per second\n$f_h$ first bending frequency, cycles per second\n$f_\\alpha$ first torsion frequency, cycles per second\n$\\omega_h = 2\\pi f_h$\n$\\omega_\\alpha = 2\\pi f_\\alpha$\n$g_h$ first bending damping coefficient\n$g_\\alpha$ first torsion damping coefficient\n\nINTRODUCTION\n\nThe background and theory for the flutter of an airfoil in a two-dimensional flow at supersonic speeds is given in reference 1. The present investigation is a preliminary survey to determine the possibility of using the theory of reference 1 for flutter at supersonic speeds to predict the coupled bending-torsion flutter of widely different unswept cantilever wings at a low supersonic Mach number. This preliminary investigation is not intended as a critical test of the theory since the analysis does not consider the effect of mode shape, aspect ratio, section shape, tip Mach cone, or viscous effects.\n\nA single-degree-of-freedom torsional instability which may occur in the Mach number range 1.0 to 1.58 is discussed in reference 1. In order to also investigate the possible occurrence of such single-degree flutter on cantilever wings, the test apparatus was designed to operate at a Mach number of 1.3.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:53:17.011137+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 11, "total_pages": 36, "image_filename": "19930082614_p11.jpg", "text": "NACA TN 1939\n\nwhere\n\n$$\nN = \\sqrt{-L/K}\n$$\n\nand\n\n$$\nC_3 = \\frac{1}{N} \\cot^{-1} (v_o/N)\n$$\n\nIn assuming that $K$ is constant the same approximations are made as in the case of level flight, together with the additional assumption that the density does not vary as a result of the inclined flight. It is necessary to estimate an average value of density, and, for large changes in altitude, some improvement in accuracy can be gained by dividing the time under consideration into shorter intervals and making separate calculations for each interval.\n\nConstant-speed dive.— If there is adequate control over the drag, an airplane can be dived at any angle at constant speed. Such control requires a net drag coefficient given by the relation\n\n$$\nC_{D_n} = - \\frac{2(W/S)\\sin\\gamma}{\\rho V^2}\n$$\n\nFor most airplanes having even moderately high wing loadings and low drag, however, it is not feasible to provide air brakes effective enough to prevent increases in speed under all conditions. The combination of factors which govern the maximum angle for a constant-speed dive can be determined from figure 3. This figure shows the relation between forward speed, angle of dive, and rate of descent. The other variables which affect the dive, such as altitude, net drag coefficient, and wing loading, are again included in the single quantity $K$. Figure 3, used in conjunction with figures 1(a) and (b), permits graphical solution for any one of these quantities when equilibrium exists.\n\nAn indication of the separate effects of altitude and speed upon the net drag required for equilibrium is given in figure 4. Wing loadings of 30 and 50 pounds per square foot were assumed. On the left in figure 4 the drag coefficient is plotted against altitude for a vertical dive. It is apparent that unless the airplane speed is very high, the equilibrium drag coefficient is large even at low altitudes and becomes extremely large at high altitudes. On the right in figure 4, curves of constant drag coefficient are plotted against the dive angle. The drag coefficient as a function of altitude and airspeed may be found for any dive angle by reading $C_{D_n}$ corresponding to the ordinate from the left part of the figure at the desired value of dive angle.", "timestamp": "2026-07-22T04:53:21.002112+00:00"}
{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 7, "total_pages": 14, "image_filename": "19930082712_p7.jpg", "text": "NACA TN 1998\n\napproximately 0.112 per degree as the Reynolds number is increased from $1.8 \\times 10^6$ to $11.0 \\times 10^6$ (fig. 2(a)). The measured angle of zero lift for the airfoil in the smooth condition varies only about $\\frac{1}{2}^\\circ$ over the range of Reynolds number covered in this investigation (fig. 2(a)).\n\nFor the NACA 8-H-12 airfoil with roughened leading edge, there appears to be a relatively insignificant variation of the maximum lift with Reynolds number; the decrement in maximum lift due to surface roughness therefore increases with Reynolds number (fig. 1(a)). The amount of variation of the lift-curve slope with Reynolds number is small when the leading edge is roughened (fig. 2(b)). In comparison with the data for the smooth condition, the addition of roughness causes the angle of zero lift to become slightly more negative at the lower Reynolds number and approximately $0.7^\\circ$ more negative at the higher Reynolds numbers so that there is substantially no variation of the angle of zero lift with Reynolds number for the rough condition.\n\nFor the smooth condition, the maximum section lift coefficient of the NACA 8-H-12 section is somewhat less than those of the NACA 0012 and NACA 23012 sections at corresponding Reynolds numbers (fig. 2(a)). The difference between the maximum lift coefficients of the NACA 8-H-12 and the NACA 0012 section is smallest at the lowest Reynolds number, becomes a maximum at a Reynolds number of $3.0 \\times 10^6$, then diminishes as the Reynolds number is increased to $9.0 \\times 10^6$. In comparison with the NACA 23012 airfoil, the difference is again smallest at the lowest Reynolds number, increases to a maximum at a Reynolds number of $3.0 \\times 10^6$, but remains relatively fixed up to a Reynolds number of $9.0 \\times 10^6$.\n\nAt corresponding Reynolds numbers, the decrement in maximum lift coefficient due to roughness is not as great for the NACA 8-H-12 section as for either the NACA 0012 or NACA 23012 airfoils, with the result that the maximum lift coefficient for the NACA 8-H-12 section at corresponding Reynolds numbers exceeds that for the NACA 0012 section and is only slightly less than that for the NACA 23012 airfoil (fig. 2(b)). The NACA 8-H-12 airfoil section, moreover, stalls in a manner which is much less abrupt than that of the two other sections mentioned, for both the smooth and rough conditions and for all corresponding Reynolds numbers within the range for which data are given (fig. 1(a) and reference 4).\n\nDrag.— For the smooth airfoil there appears to be, in most cases, some reduction in the extent of the low-drag range of lift coefficient with increasing Reynolds number (fig. 1(b)). This trend is characteristic of NACA 6-series airfoils (references 4 and 6). For most of the lift coefficients shown, the drag coefficient outside the low-drag range becomes lower in magnitude as the Reynolds number is increased. For the", "timestamp": "2026-07-22T04:53:23.083724+00:00"}