PDF/ocr_dataset_1005.jsonl

{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 19, "total_pages": 34, "image_filename": "19930086022_p19.jpg", "text": "NACA RM L9E24\n17\n\n<!-- Image (184, 110, 867, 882) -->\n\n(a) $C_{Z_a}$, $C_n$, and $C_l$ plotted against $\\alpha$.\nFigure 4.— Aileron characteristics of plain wing.", "timestamp": "2026-07-22T04:19:08.529531+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 24, "total_pages": 32, "image_filename": "19930085982_p24.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:19:12.843727+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 53, "total_pages": 104, "image_filename": "19930085842_p53.jpg", "text": "NACA RM L9C29\n49\n\nRight aileron hinge-moment coeff.\nper degree aileron deflection,\n$\\left(\\frac{\\partial C_{ha}}{\\partial \\delta_a}\\right)_{C_{ha}=0}$\n\nPitching-moment coefficient\nper degree elevator deflection,\n$\\left(\\frac{\\partial C_m}{\\partial \\delta_e}\\right)_{C_m=0}$\n\nLift coefficient, $C_L$\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 20.- Variation of $\\left(\\frac{\\partial C_m}{\\partial \\delta_e}\\right)_{C_m=0}$ and $\\left(\\frac{\\partial C_{ha}}{\\partial \\delta_a}\\right)_{C_{ha}=0}$ with lift\ncoefficient. Model in complete configuration; propellers\nremoved.", "timestamp": "2026-07-22T04:19:14.266700+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 55, "total_pages": 92, "image_filename": "19930085930_p55.jpg", "text": "54\nNACA RM L9G07\n\n[Figure: A graph plotting Average Mach number ratio against Area ratio. The y-axis ranges from 0 to 1.2. The x-axis ranges from 0.9 to 1.3. The curve shows an increasing trend that peaks around an area ratio of 1.2 and then slightly decreases. There are red \"UNCLASSIFIED\" stamps and black \"CONFIDENTIAL\" stamps over the graph area. A NACA logo is present in the bottom right corner of the graph.]\n\nAverage Mach number ratio, $\\frac{M_{2,av}}{M_1}$\n\nArea ratio, $\\frac{A_2}{A_1}$\n\nFigure 21.- The variation of the average Mach number ratio with area ratio for model 1.", "timestamp": "2026-07-22T04:19:16.409708+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 31, "total_pages": 43, "image_filename": "19930085958_p31.jpg", "text": "```markdown\n30\nNACA RM No. L9B11\n\n<!-- Image (109, 113, 859, 786) -->\n\nFigure 11.- Effects of wing-fuselage position on the characteristics of a $42^\\circ$ sweptback wing with $0.60\\frac{b}{2}$ drooped-nose flaps, split flaps, and upper-surfaces fences. $\\delta_n = 30^\\circ$.\n```", "timestamp": "2026-07-22T04:19:23.461136+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 7, "total_pages": 44, "image_filename": "19930086081_p7.jpg", "text": "NACA RM L9H05 CONFIDENTIAL 5\n\nair stream) of $6.6^\\circ$ and $15.4^\\circ$, respectively. The leading-edge wedge was modified by a small nose radius, and the sharp breaks in contour were modified by a slight fairing.\n\nThe control surface (outer one-third of the exposed span) was separated from the inner panel of the wing by a streamwise parting line and rotated about an axis perpendicular to the root chord. The axis was located at 63 percent of the control-surface root chord. The basic control was comprised of 3-percent-thick double-wedge airfoil sections measured parallel to the air stream modified by a 0.9-percent-chord leading-edge radius. A discontinuity in airfoil thickness existed at the parting line between the control surface and main panel. Errors in fabrication introduced a slight camber in the 3-percent-thick control with a maximum displacement of the mean line near the point of maximum thickness amounting to about 0.4 percent chord. An alternate control surface was tested, identical in plan form to the basic surface but having an airfoil section 7 percent thick with the maximum thickness far forward.\n\nFences of two different sizes were tested on the main wing panel at the parting line between the main wing panel and the control surface (fig. 2(b)).\n\nA few tests were made with a wing having no control surface but having 9-percent-thick tip sections on the outer one-third of the exposed wing span.\n\nAll tests of the wing and the control surface were made in the presence of a half-fuselage. Fuselages of two different sizes were used, both of which had the same nose shape. The nose section merged into a constant-diameter section at the station where the wing leading edge intersected the fuselage.\n\nTUNNEL AND TEST TECHNIQUE\n\nThe Langley 9- by 12-inch supersonic blowdown tunnel, in which the present tests were made, is a nonreturn tunnel utilizing the exhaust air from the Langley 19-foot pressure tunnel. The inlet air enters at an absolute pressure of about $2\\frac{1}{3}$ atmospheres and contains about 0.3 percent of water by weight.\n\nSemispan models are cantilevered from a 5-component strain-gage balance mounted flush with the tunnel wall. The balance rotates with the model as the angle of attack is changed and the forces and moments\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:19:25.283708+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 39, "total_pages": 50, "image_filename": "19930085952_p39.jpg", "text": "38\n\n2.0\n1.6\n$C_L$ 1.2\n.8\n.4\n0\n8 12 16 20 24 28 32 36\n$\\alpha$, deg\n\nO Articulated propeller\n□ Rigid propellers\n◇ Propellers removed\n\n.1 0 -.1\n$C_m$\n\nNACA\n\nFigure 18.- Effect of full-power operation on the variation of $\\alpha$ and $C_m$ with $C_L$ for the model with the all-movable horizontal tail removed; $\\delta_F = 0^\\circ$.\n\nNACA RM L5024", "timestamp": "2026-07-22T04:19:29.523126+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 27, "total_pages": 67, "image_filename": "19930085965_p27.jpg", "text": "26\nNACA RM E9E06\n\nper hour through air having a liquid-water content of 1 gram per cubic meter. Heat requirements were then calculated for an axial-flow compressor of the 4000-pound-thrust class for the following conditions:\n\n| Altitude (ft) | Engine speed (rpm) | Air flow (lb/sec) |\n| :--- | :--- | :--- |\n| 0 | 7600 | 75 |\n| | 7000 | 65 |\n| | 6000 | 50 |\n| 5,000 | 7600 | 65 |\n| | 7000 | 50 |\n| | 6000 | 45 |\n| 15,000 | 7600 | 50 |\n| | 7000 | 45 |\n| | 6000 | 35 |\n\nThe rate of heat dissipation was the greatest for the highest speed condition at sea level (fig. 18(a)). The average heat dissipation for this condition was approximately 5400 Btu per hour per square foot, which is equivalent to a power of approximately 11 watts per square inch of vane surface.\n\n### Shaft Power Required\n\nThe power required for protecting the inlet guide vanes under the design conditions is approximately 11.6 horsepower or about 8.7 kilowatts. At sea level and at rated engine speed of 7600 rpm, the compressor absorbs about 5680 horsepower. Thus the power required for protection against ice accretion would be only about 0.20 percent of the turbine load. In addition to the power furnished by the turbine, about 1 percent of the 11.6 horsepower required for protection will necessarily be supplied for excitation of the coils. The electric power for excitation can be supplied by batteries and generator.\n\n### Apparent Advantages and Disadvantages\n\nThe apparent advantages of an ice-protection system utilizing eddy-current heating lie in the fact that heat is produced where", "timestamp": "2026-07-22T04:19:33.209977+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 25, "total_pages": 54, "image_filename": "19930086015_p25.jpg", "text": "24\nCONFIDENTIAL\nNACA RM A9E24\n\n<!-- Image (102, 132, 864, 883) -->\n\nFigure 6.-Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:19:37.506615+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 1, "total_pages": 34, "image_filename": "19930086151_p1.jpg", "text": "NACA RM L9J28\nCONFIDENTIAL\nN62 60247\nCopy\nRM L9J28\n\nNACA\nRESEARCH MEMORANDUM\n\nLOW-SPEED INVESTIGATION OF DEFLECTABLE WING-TIP AILERONS\nON AN UNTAPERED 45° SWEPTBACK SEMISPAN WING\nWITH AND WITHOUT AN END PLATE\nBy Jack Fischel and James M. Watson\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nDATE 8-18-54\nAUTHORITY J.W.CHOWLEY\nCHANGE# 2469\nF.E.T.\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\nDecember 14, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:19:37.690474+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 70, "total_pages": 118, "image_filename": "19930085838_p70.jpg", "text": "68\nNACA RM No. L9B23\n\n<!-- Image (102, 109, 836, 927) -->\n\n(e) $\\delta_f = 25^\\circ$.\nFigure 9.- Continued.", "timestamp": "2026-07-22T04:19:39.934377+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 11, "total_pages": 20, "image_filename": "19930086060_p11.jpg", "text": "NACA RM L9F02\n9\n\nCONFIDENTIAL\n\n[Figure: Three rocket models displayed side-by-side on stands. A ruler marked \"INCHES\" is visible at the base of the rightmost model.]\n\nNACA\nL-56173\n\nFigure 2.- Test configurations having maximum diameters located at 20,\n40, and 60 percent of body length from nose.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:19:45.243052+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 43, "total_pages": 59, "image_filename": "19930085936_p43.jpg", "text": "42\nNACA RM No. E9B03\n\n<!-- Image (155, 110, 889, 838) -->\n\n(a) $\\theta = 0^\\circ$ longitudinal plane.\nFigure 8. - Pressure distributions along longitudinal planes at $10^\\circ$ angle of attack for range of yaw angles.", "timestamp": "2026-07-22T04:19:45.243353+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 9, "total_pages": 42, "image_filename": "19930086078_p9.jpg", "text": "NACA RM L9H04 CONFIDENTIAL 7\n\nvalue of 0.059 computed by the theoretical method of reference 6 (using a value of 0.1075 for section lift-curve slope (reference 7)).\n\nThe lift curves for M = 0.19 and M = 0.27 indicate an increase in $C_{L_{max}}$ with an increase in Mach number and Reynolds number; however, an adverse compressibility effect at high lift coefficients, accompanied by a decrease in $C_{L_{max}}$, can be noted for M = 0.37. These effects correlate well with data from references 7 and 8 pertaining to Reynolds number and Mach number effects.\n\n45° sweptback wing.— For the 45° sweptback wing configuration, $C_{L_\\alpha}$ increased very slightly and the drag and pitching-moment characteristics changed negligibly as Mach number and Reynolds number were increased (fig. 5). The pitching-moment data for the sweptback wing indicate that the aerodynamic center was about 5 percent mean aerodynamic chord ahead of the $\\bar{c}/4$ at low lift coefficients; however, at higher lift coefficients and through the stall region the wing was stable.\n\nThe experimental value of $C_{L_\\alpha}$ obtained on the sweptback wing at M = 0.19 is 0.036. This value compares very well with similar values of $C_{L_\\alpha}$ of 0.037 computed by the theoretical method of reference 6 (which accounts for sweep angle) and by the empirical method recommended in reference 5 (which considers aspect ratio as the only variable). This agreement between the estimated and measured values of $C_{L_\\alpha}$ tends to substantiate the point made in reference 6 that sweep angle has little effect on $C_{L_\\alpha}$ for wings of low aspect ratio.\n\nThe lift data for M = 0.37 were not obtained at lift coefficients high enough to observe any compressibility effect similar to that noted on the unswept wing configuration. However, the lift curves for M = 0.19 and M = 0.27 show a negligible change in $C_{L_{max}}$ with increase in Mach number and Reynolds number, possibly indicating the onset of adverse compressibility effects, or little effect of Reynolds numbers of $4.5 \\times 10^6$ and $6.3 \\times 10^6$.\n\nComparison of the unswept and sweptback plain-wing configurations.— Comparison of the plain-wing aerodynamic data for the unswept and swept-back configurations shows that the results vary with aspect ratio and sweep angle as would be predicted by theory. The value of $C_{L_\\alpha}$ was higher for the unswept configuration, primarily because of the higher aspect ratio of the unswept wing (reference 6). A larger value of $C_{L_{max}}$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:19:45.520065+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 3, "total_pages": 22, "image_filename": "19930086105_p3.jpg", "text": "NACA RM E9H12 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION OF PRESSURE FLUCTUATIONS\n\nIN 3.6-INCH RAM JET AT MACH NUMBER 1.92\n\nBy James F. Connors and Albert H. Schroeder\n\nSUMMARY\n\nAn experimental study of the pressure fluctuations at the combustion-chamber inlet of a 3.6-inch-diameter ram jet operating at a Mach number of 1.92 was conducted at the NACA Lewis 20-inch supersonic tunnel. Results based on single-point instantaneous pressure measurements in an unsteady-flow system are presented for two burner configurations of different operating characteristics in order to indicate trends in pressure magnitude, frequency, and wave form. The amplitude of the pressure fluctuations and the mean static pressure at the combustion-chamber inlet are shown as a function of fuel-air ratio for several values of outlet-inlet area ratio.\n\nFor each outlet-inlet area ratio investigated, the amplitude of the pressure fluctuations built up gradually as the fuel-air ratio was increased to the value corresponding to optimum diffuser mean static pressure (as indicated on manometers by static orifices). With further increase in fuel-air ratio, the pressure pulsations became more intense and the shock oscillated in and out of the diffuser inlet. The fundamental frequency increased from 6 to 35 cycles per second as the fuel-air ratio was increased beyond the value at optimum mean static pressure. A typical wave form is presented for this type of pressure fluctuation. With both burner configurations at conditions of optimum mean combustion-chamber static pressure, the maximum instantaneous pressure coincided (within experimental error) with the static pressure at optimum cold recovery. With the conical flame-holder configuration, the minimum instantaneous pressure at optimum mean pressure recovery approximated the cold static pressure corresponding to the particular outlet area.\n\nA marked improvement in diffuser performance was realized with a regenerative-type burner as compared with the perforated conical flame-holder configuration.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:19:47.792642+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 104, "total_pages": 122, "image_filename": "19930082090_p104.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:19:48.014664+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 47, "total_pages": 47, "image_filename": "19930085918_p47.jpg", "text": "```markdown\n46\nNACA RM A9D29\n\n<!-- Image (181, 98, 752, 414) -->\n\n(a) Experimental pressure distributions over section at 28.1% semispan.\n\n<!-- Image (181, 509, 752, 716) -->\n\n(b) Theoretical pressure distributions. [NACA logo]\n\nFigure 12. - Comparisons between upper-surface pressure distributions of basic wing, wing with full-span leading-edge flap deflected 30° down, and wing with full-span cambered nose.\n```", "timestamp": "2026-07-22T04:19:52.461980+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 20, "total_pages": 34, "image_filename": "19930086022_p20.jpg", "text": "18\nNACA RM L9E24\n\n<!-- Image (169, 103, 759, 885) -->\n\n(b) $P_R$ and $C_{h_a}$ plotted against $\\alpha$.\nFigure 4.- Continued.", "timestamp": "2026-07-22T04:19:53.415659+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 71, "total_pages": 82, "image_filename": "19930085551_p71.jpg", "text": "70\n\nNACA RM No. L8E30\n\n(b) Corrective control applied.\nFigure 17.-- Concluded.\n\n| Indicated airspeed, mph | Control angle, deg | Wheel force, lb | Angular velocity, rad/sec | Angle of bank, deg | Acceleration, g |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| | Down Up | Push Pull | Down Up | Sideslip angle, deg | Down Up |\n| | Left Right | Left Right | Left Right | Left Right | Left Right |\n| 100 | 10 | 40 | .2 | 10 | 1 |\n| 110 | 0 | 0 | 0 | 0 | 0 |\n| 0 | 10 | 40 | .2 | 5 | |\n| | | | | 0 | |\n| | | | | 5 | |\n\n| Time, sec | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| 0 | | | | | |\n| 4 | | | | | |\n| 8 | | | | | |\n| 12 | | | | | |\n| 16 | | | | | |\n| 20 | | | | | |\n| 24 | | | | | |\n| 28 | | | | | |\n\n[Graph showing Indicated airspeed, mph vs Time, sec. A line labeled NACA is plotted.]\n\n[Graph showing Control angle, deg vs Time, sec. Lines labeled Elevator, Rudder, and Total aileron are plotted.]\n\n[Graph showing Wheel force, lb vs Time, sec. Lines labeled Elevator, Rudder, and Aileron are plotted.]\n\n[Graph showing Angular velocity, rad/sec vs Time, sec. Lines labeled Pitch, Yaw, and Roll are plotted.]\n\n[Graph showing Angle of bank, deg and Sideslip angle, deg vs Time, sec. Lines labeled Bank and Sideslip are plotted.]\n\n[Graph showing Acceleration, g vs Time, sec. Lines labeled Normal and Transverse are plotted.]", "timestamp": "2026-07-22T04:19:58.261827+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 56, "total_pages": 92, "image_filename": "19930085930_p56.jpg", "text": "NACA RM L9G07\n55\n\n$$\\frac{P_{2av}}{P_o}$$\nAverage stagnation-pressure recovery,\n\n[Figure: A graph plotting Average stagnation-pressure recovery against Area ratio. The y-axis ranges from 0 to 1.0. The x-axis ranges from .9 to 1.3. Data points are plotted as circles, with a curve fitted through them. The graph contains two red \"CONFIDENTIAL\" stamps and a NACA logo.]\n\nArea ratio, $$\\frac{A_2}{A_1}$$\n\nFigure 22.- The variation of the average stagnation-pressure recovery with area ratio for model 1.", "timestamp": "2026-07-22T04:20:05.026096+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 40, "total_pages": 50, "image_filename": "19930085952_p40.jpg", "text": "NACA RM L9C24\n39\n\n<!-- Image (120, 135, 935, 806) -->\n\n(a) Articulated propellers.\nFigure 19.- Variation of $C_m$ with $C_L$ for several elevator deflections.\n$C_{DR} \\approx 0; \\delta_F = 0^\\circ$.", "timestamp": "2026-07-22T04:20:10.380133+00:00"}
{"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 78, "total_pages": 85, "image_filename": "19930085529_p78.jpg", "text": "```markdown\nNACA RM No. L8A30a\n77\n\nTABLE 71\n$$\n\\left[ \\Lambda = -30^\\circ, \\delta_{te} = 10.0^\\circ, \\alpha = 4^\\circ \\right]\n$$\n\nCONFIDENTIAL\n\n| UPPER SURFACE | | | | | | | | LOWER SURFACE | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Tube** | **Per-cent chord** | **Mach Number** | | | | | | **Tube** | **Per-cent chord** | **Mach Number** | | | | | |\n| | | **0.60** | **0.80** | **0.85** | **0.89** | **0.925** | **0.96** | | | **0.60** | **0.80** | **0.85** | **0.89** | **0.925** | **0.96** |\n| 1 | 2.0 | - | - | - | - | - | - | 86 | 3.0 | - | - | - | - | - | - |\n| 2 | 6.0 | - | - | - | - | - | - | 87 | 10.0 | - | - | - | - | - | - |\n| 3 | 15.0 | - | - | - | - | - | - | 88 | 25.0 | - | - | - | - | - | - |\n| 4 | 27.5 | - | - | - | - | - | - | 89 | 41.0 | - | - | - | - | - | - |\n| 5 | 40.0 | - | - | - | - | - | - | 90 | 52.5 | -0.003 | -0.067 | -0.109 | -0.140 | -0.176 | -0.199 |\n| 6 | 50.0 | -0.003 | -0.181 | -0.328 | -0.506 | -0.796 | -1.720 | 91 | 62.5 | -.025 | -.054 | -.088 | -.111 | -.168 | -.316 |\n| 7 | 59.0 | -.181 | -.462 | -.891 | -1.246 | -1.846 | -3.546 | 92 | 72.5 | -.011 | -.052 | -.098 | -.121 | -.181 | -.407 |\n| 8 | 67.5 | -.132 | -.387 | -.508 | -.511 | -.340 | -.532 | 93 | 84.0 | - | - | - | - | - | - |\n| 9 | 77.5 | - | - | - | - | - | - | 94 | 84.0 | - | - | - | - | - | - |\n| 10 | 89.5 | - | - | - | - | - | - | | | | | | | | |\n| 11 | 96.0 | - | - | - | - | - | - | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| 12 | 2.0 | -.166 | -.629 | -.560 | -.782 | -.787 | -.640 | 95 | 3.0 | .606 | .568 | .547 | .507 | .475 | .461 |\n| 13 | 6.0 | -.171 | -.638 | -.568 | -.738 | -.770 | -.756 | 96 | 10.0 | .278 | .298 | .242 | .213 | .192 | .196 |\n| 14 | 15.0 | -.031 | -.601 | -.518 | -.648 | -.874 | -.806 | 97 | 25.0 | .070 | .077 | .044 | .015 | -.008 | -.021 |\n| 15 | 27.5 | -.648 | -.633 | -.533 | -.575 | -.884 | -.841 | 98 | 41.0 | -.061 | -.103 | -.138 | -.189 | -.260 | -.272 |\n| 16 | 40.0 | -.474 | -.609 | -.551 | -.570 | -.792 | -.883 | 99 | 52.5 | -.068 | -.111 | -.144 | -.189 | -.270 | -.346 |\n| 17 | 50.0 | -.381 | -.571 | -.551 | -.515 | -.684 | -.838 | 100 | 62.5 | -.036 | -.071 | -.106 | -.131 | -.186 | -.315 |\n| 18 | 59.0 | -.299 | -.551 | -.551 | -.511 | -.486 | -.825 | 101 | 72.5 | -.017 | -.054 | -.091 | -.101 | -.165 | -.189 |\n| 19 | 67.5 | -.198 | -.440 | -.499 | -.435 | -.277 | -.755 | 102 | 84.0 | -.078 | -.094 | -.018 | .000 | .003 | -.117 |\n| 20 | 77.5 | -.110 | -.318 | -.399 | -.380 | -.461 | -.730 | 103 | 84.0 | .123 | .083 | .030 | .006 | .016 | -.039 |\n| 21 | 89.0 | -.002 | -.165 | -.236 | -.358 | -.368 | -.594 | | | | | | | | |\n| 22 | 95.5 | - | - | - | - | - | - | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| 23 | 2.0 | -.205 | -.087 | -.892 | -.653 | -.534 | -.387 | 104 | 3.0 | .514 | .548 | .532 | .504 | .484 | .481 |\n| 24 | 6.0 | -.097 | -.130 | -.968 | -.792 | -.668 | -.545 | 105 | 10.0 | .278 | .275 | .265 | .246 | .232 | .238 |\n| 25 | 15.0 | -.755 | -.060 | -.958 | -.798 | -.668 | -.562 | 106 | 25.0 | .070 | .077 | .044 | .015 | -.008 | -.028 |\n| 26 | 27.5 | -.618 | -.925 | -.913 | -.842 | -.711 | -.671 | 107 | 41.0 | -.031 | -.097 | -.077 | -.116 | -.133 | -.194 |\n| 27 | 40.0 | -.541 | -.921 | -.849 | -.838 | -.655 | -.671 | 108 | 52.5 | -.056 | -.111 | -.144 | -.168 | -.225 | -.253 |\n| 28 | 50.0 | -.442 | -.473 | -.492 | -.632 | -.775 | -.781 | 109 | 62.5 | -.028 | -.096 | -.076 | -.119 | -.185 | -.233 |\n| 29 | 59.0 | -.359 | -.394 | -.408 | -.628 | -.867 | -.861 | 110 | 72.5 | -.015 | -.054 | -.091 | -.101 | -.127 | -.168 |\n| 30 | 67.5 | -.240 | -.281 | -.263 | -.515 | -.749 | -.792 | 111 | 85.1 | .067 | .073 | .063 | .098 | .018 | -.048 |\n| 31 | 77.5 | -.134 | -.170 | -.160 | -.336 | -.644 | -.557 | 112 | 84.0 | .123 | .123 | .104 | .105 | .061 | -.022 |\n| 32 | 88.0 | -.019 | -.044 | -.072 | -.117 | -.244 | -.274 | | | | | | | | |\n| 33 | 95.5 | - | - | - | - | - | - | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| 34 | 2.0 | -.284 | -.030 | -.811 | -.627 | -.501 | -.381 | 113 | 3.0 | .508 | .516 | .505 | .476 | .459 | .454 |\n| 35 | 15.0 | -.687 | -.094 | -.913 | -.770 | -.668 | -.596 | 114 | 10.0 | .260 | .280 | .273 | .254 | .244 | .251 |\n| 36 | 27.5 | -.606 | -.900 | -.903 | -.830 | -.711 | -.630 | 115 | 25.0 | .070 | .080 | .059 | .032 | .019 | .030 |\n| 37 | 40.0 | -.511 | -.848 | -.854 | -.911 | -.632 | -.737 | 116 | 41.0 | -.069 | -.071 | .092 | -.090 | -.119 | -.115 |\n| 38 | 50.0 | -.468 | -.798 | -.814 | -.814 | -.614 | -.737 | 117 | 52.5 | -.071 | -.103 | -.071 | -.120 | -.157 | -.157 |\n| 39 | 59.0 | -.361 | -.409 | -.581 | -.809 | -.832 | -.798 | 118 | 62.5 | -.019 | -.036 | -.046 | -.097 | -.141 | -.169 |\n| 40 | 67.5 | - | - | - | - | - | - | 119 | 72.5 | -.015 | -.054 | -.091 | -.101 | -.127 | -.168 |\n| 41 | 77.5 | -.148 | -.137 | -.122 | -.227 | -.456 | -.779 | 120 | 87.4 | .078 | .106 | .123 | .068 | .029 | .011 |\n| 42 | 87.5 | -.039 | -.041 | -.061 | -.004 | -.452 | -.452 | 121 | 84.0 | .097 | .117 | .109 | .061 | .000 | .011 |\n| 43 | 94.2 | .014 | .041 | .061 | .004 | -.161 | -.269 | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| 44 | 2.0 | -.797 | -.182 | -.917 | -.724 | -.546 | - | 122 | 3.0 | .590 | .589 | .570 | .546 | .518 | - |\n| 45 | 6.0 | -.813 | -.125 | -.933 | -.730 | -.585 | - | 123 | 10.0 | .317 | .337 | .327 | .310 | .283 | - |\n| 46 | 15.0 | -.719 | -.153 | -.975 | -.838 | -.703 | - | 124 | 25.0 | .111 | .121 | .111 | .092 | .072 | - |\n| 47 | 27.5 | -.631 | -.900 | -.953 | -.861 | -.746 | - | 125 | 41.0 | .020 | .029 | .020 | .007 | -.012 | - |\n| 48 | 40.0 | -.593 | -.095 | -.962 | -.811 | -.759 | - | 126 | 52.5 | .014 | .019 | .011 | -.016 | -.041 | - |\n| 49 | 50.0 | -.477 | -.925 | -.942 | -.825 | -.746 | - | 127 | 62.5 | .014 | .019 | .011 | -.016 | -.041 | - |\n| 50 | 59.0 | -.388 | -.395 | -.905 | -.703 | -.692 | - | 128 | 72.5 | .040 | .044 | .035 | -.011 | -.066 | - |\n| 51 | 67.5 | -.275 | -.281 | -.263 | -.515 | -.749 | - | 129 | 78.0 | .077 | .083 | .077 | .000 | -.063 | - |\n| 52 | 77.5 | -.111 | -.135 | -.121 | -.493 | -.715 | - | 130 | 85.3 | .081 | .106 | .104 | .043 | -.023 | - |\n| 53 | 88.5 | -.019 | -.041 | -.061 | -.3", "timestamp": "2026-07-22T04:20:11.143166+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 54, "total_pages": 104, "image_filename": "19930085842_p54.jpg", "text": "50\nNACA RM L9C29\n\n[Figure: Graph plotting Pitching-moment coefficient, $C_m$ (y-axis) against Lift coefficient, $C_L$ (x-axis). The y-axis ranges from -4 to 2. The x-axis ranges from -2 to 8. The graph contains multiple curves representing different values of $\\delta_a, deg$ (15, 10, 5, 0, -5). A white rectangular box obscures part of the grid near the bottom center. The logo \"NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\" is present in the bottom right corner of the graph area.]\n\nFigure 21.- Variation of pitching-moment coefficient with lift coefficient. Model in complete configuration; propellers removed.", "timestamp": "2026-07-22T04:20:13.766967+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 2, "total_pages": 34, "image_filename": "19930086151_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:20:13.963996+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 8, "total_pages": 44, "image_filename": "19930086081_p8.jpg", "text": "6 CONFIDENTIAL NACA RM L9H05\n\nare measured with respect to the balance axes. In measuring the forces and moments acting on the control surface, the surface was connected with the balance independent of the wing panel by means of a mounting staff which extended spanwise through an internal slot in the main wing. The half-span wing models are tested in the presence of, but not attached to, a half-fuselage shimmed out 0.25 inch from the tunnel wall. The finite gap existing between the wing and fuselage is believed to have no influence on the flap loading. (See reference 2.)\n\nThe dynamic pressure and test Reynolds number decreased about 5 percent during the course of each run because of the decreased pressure of the inlet air. The average dynamic pressure was 11.8 pounds per square inch, and the average Reynolds number, based on the mean aerodynamic chord of the exposed wing, was $4.0 \\times 10^6$.\n\nPRECISION OF DATA\n\nFree-stream Mach number has been calibrated at $1.90 \\pm 0.02$. This Mach number was used in determining the dynamic pressure. Calibration tests which were made with the model removed indicated that the static pressure varied about $\\pm 1.5$ percent from a mean value for the region normally occupied by the wing. A discussion is given in reference 2 of the various factors which might influence the test results, such as humidity effects and method of mounting.\n\nAn estimate has been made of the probable errors to be found in the measured test points, when fluctuations in the readings of the measuring equipment, calibration errors, and shift of instrument no-load readings experienced during the course of each test are considered. The following table lists the errors that might be expected to exist between the test points for each particular figure.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:20:14.996565+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 26, "total_pages": 54, "image_filename": "19930086015_p26.jpg", "text": "NACA RM A59E24\nCONFIDENTIAL\n\nFlagged symbols\ndenote survey\noff center line\n\nVertical distance from tunnel center line, z, in.\n24\n16\n8\n0\n-8\n-16\n-24\n\nx=-24\nx=0\nx=21\n\n-.04 0 .04\n-.08 -.04 0 .04 .08\n-.08 -.04 0 .04\n\nStream pressure coefficient, $\\Delta p/q$\n\n(a) D=165.12; M=1.23.\n\nFigure 7.— The variation of static pressure vertically in the Ames 6- by 6-foot supersonic\nwind tunnel. y=0; stagnation pressure = 9 lb/sq in. abs.\n\nCONFIDENTIAL\n25", "timestamp": "2026-07-22T04:20:16.853321+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 71, "total_pages": 118, "image_filename": "19930085838_p71.jpg", "text": "```markdown\nNACA RM No. L9B23\n\nAileron section hinge-moment coefficient, $c_{h_a}$\n\nSection angle of attack, $\\alpha_o$, deg\n\n$\\delta_f = -20^\\circ$\n\n$\\delta_a$ (deg)\n-20\n-17\n\n0\n\nNACA\n\n(r) $\\delta_f = 25^\\circ$.\n\nFigure 9.- Continued.\n\n69\n```", "timestamp": "2026-07-22T04:20:17.044767+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 12, "total_pages": 20, "image_filename": "19930086060_p12.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:20:17.329345+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 28, "total_pages": 67, "image_filename": "19930085965_p28.jpg", "text": "NACA RM E9E06\n27\n\nneeded with little time lag. Most of the heat is used to warm the surface subject to icing and not the induction air as in the exhaust bleedback systems. The loss in net thrust would therefore be small because the increase in the temperature of the incoming mass of induction air due to surface heating would be slight.\n\nIf ice accumulations occur before heat is applied, de-icing can be accomplished because eddy-current heating will melt the under surface of the ice coating and permit washing-back of the accumulated ice by the induction air. This washing-back may prove to be a disadvantage, however, if the pieces broken off are allowed to become large enough to damage the compressor blading. The application of some heat before the ice is allowed to accumulate may therefore be necessary.\n\nBecause no special ducting is required for eddy-current heating, the design may be kept within the limits of the maximum diameter of the compressor. Eddy-current heating may have to be supplemented with some other system of heating, however, to protect the lips and the surface of the inlet cowling. Corrosive exhaust gases need not be fed back into the compressor, however, because they can be contained within the walls of the inlet.\n\nThe modifications with eddy-current heating appear feasible. The system of protection by means of eddy-current heating can be applied back through the compressor for as many stages as is deemed necessary. The heat requirement for successive stages would diminish with stage number. Application of this principle, however, becomes more difficult with successive stages because of limited clearances and complexity of electric circuits.\n\nWith eddy-current heating, the use of intensity-controlled heating would be possible because the controls would be simple.\n\nThe disadvantages of the application of eddy-current heating are the necessity for including magnetic material in the flux circuit and the corresponding weight increase of these materials and of the coil and the chopper.\n\nSUMMARY OF RESULTS\n\nFrom an investigation of eddy-current heating for icing protection of axial-flow-compressor blades, the following results were obtained:", "timestamp": "2026-07-22T04:20:21.089324+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 25, "total_pages": 32, "image_filename": "19930085982_p25.jpg", "text": "NACA RM E9E13\n\n- Orifice tank\n- Inlet throttle\n- Thin-plate orifice\n\n- High-speed motor\n- Stator\n- Guide vanes\n- Rotor\n- Collector\n- Outlet throttle\n- Screens\n- Depression tank\n\n[Figure: Schematic diagram of compressor installation.]\n\nFigure 3. — Schematic diagram of compressor installation.\n\nNACA\n\n23", "timestamp": "2026-07-22T04:20:24.317776+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 44, "total_pages": 59, "image_filename": "19930085936_p44.jpg", "text": "NACA RM No. E9B03\n43\n\nPressure coefficient, Cp\nAngle of yaw (deg)\nO -12\n□ -6\n◇ 0\n△ 6\n▽ 12\n\n.20\n.10\n0\n-.10\n-.20\n-.30\n\n0 .2 .4 .6 .8 1.0\nDistance from tip, x/L\n\n(b) $\\theta = 45^\\circ$ longitudinal plane.\n\nFigure 8. - Continued. Pressure distributions along longitudinal planes at $10^\\circ$ angle of attack for range of yaw angles.", "timestamp": "2026-07-22T04:20:24.933668+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 105, "total_pages": 122, "image_filename": "19930082090_p105.jpg", "text": "NACA TN No. 1455\n103\n\n[Figure: A black and white photograph of a cylindrical, finned object, likely a heat exchanger, with a circular flange at the top. A dark rectangular bar is visible above the object.]\n\nFigure 53.- Aluminum-alloy fin heat exchanger N.\n[annotation: NACA]", "timestamp": "2026-07-22T04:20:25.105952+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 21, "total_pages": 34, "image_filename": "19930086022_p21.jpg", "text": "NACA RM L9E24\n19\n\n<!-- Image (221, 119, 821, 879) -->\n\n(c) $C_m$ and $C_L$ plotted against $\\alpha$.\nFigure 4.- Concluded.", "timestamp": "2026-07-22T04:20:26.497373+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 72, "total_pages": 82, "image_filename": "19930085551_p72.jpg", "text": "NACA RM No. L8K30\n71\n\n<!-- Image (323, 136, 646, 801) -->\n\n(a) No corrective control applied. No. 1 engine fails at 0.7 second.\nFigure 18.- Time histories of airplane motions during a simulated take-off in which No. 1 engine fails. C-54D airplane; flaps 20°; gear down; power 43 in. Hg; 2550 rpm; No. 1 engine cut to idling.", "timestamp": "2026-07-22T04:20:29.881165+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 4, "total_pages": 22, "image_filename": "19930086105_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM E5H12\n\nINTRODUCTION\n\nA prime requisite for efficient ram-jet operation at supersonic velocities is the attainment of high total-pressure recoveries in the diffusion process. Several schemes (references 1 to 4) have been successful in attaining efficient diffusion in cold tests using steady-flow resistances (outlet-area restrictions). As shown in references 5 to 7, however, fluctuating pressures from the combustion process caused a deterioration of the optimum total-pressure recovery obtainable with three types of supersonic diffuser. These pressure fluctuations resulted in a decrease in optimum total-pressure recovery with either increasing outlet area or increasing rates of heat release. Thus the attainment of high diffuser total-pressure recoveries appears to be limited by the degree of roughness (pressure fluctuation) in the combustion process.\n\nIn order to study further the nature of these pressure pulsations and their detrimental effect on diffuser total-pressure recovery, a preliminary experimental investigation was conducted at the NACA Lewis laboratory on a 3.6-inch-diameter ram jet employing the same shock diffuser and burner configuration that had previously exhibited rough combustion characteristics (reference 5). In addition, a regenerative-type burner similar to that developed by the Research and Development Service Suboffice at Fort Bliss, Texas, was also investigated. High-speed schlieren photographs were taken of the unsteady-shock patterns at the inlet and instantaneous measurements were made of the amplitudes and frequencies of the pressure pulsations at the combustion-chamber inlet. Inasmuch as the unsteady flow produced by the burner probably set up standing waves along the duct, the single-point instantaneous measurements at the combustion-chamber inlet do not necessarily present a complete quantitative picture of the pressure fluctuations. Such data, however, should be of qualitative interest in establishing trends in pressure magnitude, frequency, and wave form.\n\nMaximum and minimum pressures at the combustion-chamber inlet and improved diffuser performance obtained with the regenerative-type burner are presented for the operating range of fuel-air ratios and outlet areas.\n\nAPPARATUS AND PROCEDURE\n\nExperiments were conducted in the NACA Lewis 20-inch supersonic tunnel, which was operated at a Mach number of $1.92 \\pm 0.04$. In the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:20:30.062946+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 10, "total_pages": 42, "image_filename": "19930086078_p10.jpg", "text": "8 CONFIDENTIAL NACA RM L9H04\n\nwas obtained on the sweptback configuration than on the unswept configuration, an effect which has been found previously in other investigations (for example, reference 9).\n\nDrag coefficients of the sweptback wing were generally larger than those of the unswept wing, especially at large lift coefficients. Calculations indicate that this is accounted for mostly by the lower aspect ratio and the consequent larger values of induced drag of the sweptback wing. The maximum lift-drag ratio (which occurred at $C_L \\approx 0.2$) was about 16 and 12 for the unswept and sweptback wing configurations, respectively.\n\nThe aerodynamic center was ahead of the $\\bar{c}/4$ approximately the same amount in percent mean aerodynamic chord at zero lift coefficient for both wing configurations. Both wing configurations exhibited stable stall characteristics; however, the unswept wing had a more nearly linear variation of $C_m$ with $\\alpha$.\n\n### Lateral Control Characteristics\n\nThe rolling-moment and yawing-moment coefficients obtained for several extensions and deflections of the various plan forms of extensible wing-tip ailerons on the unswept wing configuration are shown in figures 6 to 11, and similar data obtained on the sweptback wing configuration are shown in figures 12 to 17. As previously discussed under the section entitled \"Corrections,\" the rolling-moment and yawing-moment data presented in these figures are uncorrected for reflection-plane effects.\n\n**Unswept wing configuration.**— The rolling-moment coefficients generally increased with increase in $\\alpha$ for all aileron deflections and extensions on the unswept wing configuration.\n\nThe rolling-moment data indicate a reversal of direction of roll at some negative angles of attack, a highly undesirable condition for inverted flight and some maneuvers. Utilizing greater aileron deflections than those used in this investigation would probably relieve the undesirable condition somewhat since reversal of rolling moment occurs at larger negative angles of attack with increased aileron deflections.\n\nDeflecting the large-chord and triangular ailerons caused fairly linear increases in rolling moment for the deflection range tested (figs. 6 and 8). For the short-chord aileron, no appreciable increase in rolling moment was gained by deflecting the aileron beyond $4^\\circ$ at positive angles of attack, probably because of early separation over the aileron which was accentuated with increased deflection. However, the rolling moment produced by the short-chord aileron continued to\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:20:30.919755+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 32, "total_pages": 43, "image_filename": "19930085958_p32.jpg", "text": "NACA RM No. L9B11\n31\n\n<!-- Image (124, 80, 874, 828) -->\n\nFigure 12.- Effects of wing-fuselage position on the characteristics of a 42° sweptback wing with 0.60b/2 drooped-nose flaps and upper-surface fences. $\\delta_n = 30^\\circ$.", "timestamp": "2026-07-22T04:20:35.212276+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 57, "total_pages": 92, "image_filename": "19930085930_p57.jpg", "text": "```markdown\n56\n\nCONFIDENTIAL\nUNCLASSIFIED\n\n$M_{c2av}$\nAverage exit Mach number,\n\n| | |\n| :--- | :--- |\n| O | $\\frac{p_1}{p_A} = 1.28$ |\n| $\\square$ | $\\frac{p_1}{p_A} = 1.00$ |\n| $\\diamond$ | $\\frac{p_1}{p_A} = .85$ |\n\nConcave surface\n\nNACA\n\nDistance from convex surface\n\n(a) 50-percent-span station.\n\nFigure 23.- The variation of the average exit Mach number with distance from convex surface for three static-pressure ratios at an area ratio of 1.077 for model 2.\n\nNACA RM E9G07\n```", "timestamp": "2026-07-22T04:20:35.721159+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 41, "total_pages": 50, "image_filename": "19930085952_p41.jpg", "text": "40\nNACA RM L9C24\n\n<!-- Image (196, 97, 676, 820) -->\n\n(a) $\\alpha \\approx 41^\\circ$.\nFigure 20.- Effect of propeller articulation on the variation of $C_m$, $C_L$, and $C_{h_a}$ with $\\delta_a$ for conditions of $C_{D_R} = 0$. $\\beta = 11.5^\\circ$; $\\delta_F = 0^\\circ$.", "timestamp": "2026-07-22T04:20:37.434543+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 55, "total_pages": 104, "image_filename": "19930085842_p55.jpg", "text": "NACA RM L9C29\n51\n\nNeutral-point location, percent of MAC\nLift coefficient, $C_L$\nNormal c.g. location\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\nFigure 22.- Variation of stick-fixed neutral points with lift coefficient.\nModel in complete configuration; propellers removed.", "timestamp": "2026-07-22T04:20:45.036495+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 1, "total_pages": 37, "image_filename": "19930090382_p1.jpg", "text": "FILE COPY\nNO 2\nCONFIDENTIAL\nCopy\nRM L9I07\n230\n\nNACA\nRESEARCH MEMORANDUM\n\nINVESTIGATION OF THE NACA 4-(4)(06)-04 TWO-BLADE\nPROPELLER AT FORWARD MACH NUMBERS TO 0.925\n\nBy James B. Delano and Daniel E. Harrison\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nTHIS DOCUMENT ON LOAN FROM Langley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: J.W. CROWLEY\nDATE: 8-13-54\nCHANGE #2474\nWHL\n\nRETURN TO THE ABOVE ADDRESS.\n\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\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 may be disclosed\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nOctober 28, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:20:46.772211+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 3, "total_pages": 34, "image_filename": "19930086151_p3.jpg", "text": "NACA RM L9J28\nCONFIDENTIAL\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nRESEARCH MEMORANDUM\nLOW-SPEED INVESTIGATION OF DEFLECTABLE WING-TIP AILERONS\nON AN UNTAPERED 45° SWEPTBACK SEMISPAN WING\nWITH AND WITHOUT AN END PLATE\nBy Jack Fischel and James M. Watson\n\nSUMMARY\n\nA low-speed wind-tunnel investigation to determine the characteristics of deflectable wing-tip ailerons on an untapered 45° sweptback semispan wing was made in the Langley 300 MPH 7- by 10-foot tunnel. The ailerons investigated had triangular and parallelogram plan forms with a maximum chord of 0.625 wing chord and a flat-plate profile. These ailerons were tested on the plain wing and on the wing with a rectangular end plate (to simulate a vertical fin) mounted inboard of the ailerons.\n\nThe results of the investigation indicated that the plan form of the aileron had little effect on the lift, drag, and pitching-moment characteristics of the wing. The addition of the end plate, however, increased the wing lift-curve slope and the drag, but decreased the maximum lift and the lift-drag ratio of the wing.\n\nAileron plan form generally had little effect on the values of rolling-moment coefficient produced by aileron deflection; however, the ailerons were more effective on the plain wing than on the wing with end plate. The ailerons should provide adequate lateral control over the entire angle-of-attack range investigated. The yawing moments resulting from aileron deflection were generally adverse - particularly at large angles of attack and aileron deflections.\n\nINTRODUCTION\n\nThe NACA is currently investigating various devices for use in providing adequate lateral control on transonic and supersonic wing configurations. The deflectable wing-tip aileron is one of the control devices being investigated. This aileron consists of the entire tip of\nCONFIDENTIAL", "timestamp": "2026-07-22T04:20:52.783720+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 72, "total_pages": 118, "image_filename": "19930085838_p72.jpg", "text": "70\nNACA RM No. L9B23\n\n<!-- Image (109, 134, 835, 857) -->\n\n(g) $\\delta_f = 40^\\circ$.\nFigure 9.- Continued.", "timestamp": "2026-07-22T04:20:53.271076+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 27, "total_pages": 54, "image_filename": "19930086015_p27.jpg", "text": "26\nCONFIDENTIAL\nNACA RM A50E24\n\nVertical distance from tunnel center line, z, in.\n24\n16\n8\n0\n-8\n-16\n-24\n-.04 0 .04\nx=-24\n\nFlagged symbols denote survey off center line\n-.08 -.04 0 .04 .08\nx=0\n\n-.08 -.04 0 .04\nx=21\n\nStream pressure coefficient, $\\Delta p/q$\n\n(b) D=147.20; M=1.32. 9/22, 8-22\n\nFigure 7.- Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T04:20:53.486956+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 106, "total_pages": 122, "image_filename": "19930082090_p106.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:21:00.344003+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 73, "total_pages": 82, "image_filename": "19930085551_p73.jpg", "text": "```markdown\n72\nNACA RM No. L8K30\n\n<!-- Image (162, 116, 752, 881) -->\n\n(b) Corrective control applied. No. 1 engine fails at 0.8 second.\n\nFigure 18.— Concluded.\n```", "timestamp": "2026-07-22T04:21:04.988511+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 22, "total_pages": 34, "image_filename": "19930086022_p22.jpg", "text": "20\nNACA RM L9E24\n\n$$C_{Z_a}$$\n$$C_n$$\n$$C_l$$\n\n$$\\delta_a$$\n(deg)\n25\n20\n15\n10\n7\n6\n5\n4\n3\n2\n1\n0\n-1\n-2\n-3\n-4\n-5\n-10\n-15\n-20\n-25\n\nNACA\n\n$$\\alpha, \\text{deg}$$\n-4 0 4 8 12 16 20\n\n(a) $$C_{Z_a}$$, $$C_n$$, and $$C_l$$ plotted against $$\\alpha$$.\n\nFigure 5.- Aileron characteristics of wing with leading- and trailing-edge flaps deflected.", "timestamp": "2026-07-22T04:21:06.188363+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 58, "total_pages": 92, "image_filename": "19930085930_p58.jpg", "text": "```markdown\nNACA RM L9907\n\nCONFIDENTIAL\n\nAverage exit Mach number, $M_{2,av}$\n\n$$\n\\begin{array}{c|c}\n\\circ & \\frac{p_1}{p_A} = 1.28 \\\\\n\\hline\n\\square & \\frac{p_1}{p_A} = 1.00 \\\\\n\\hline\n\\diamond & \\frac{p_1}{p_A} = .85\n\\end{array}\n$$\n\nConcave surface\n\nDistance from convex surface\n\n(b) 10.15-percent-span station.\n\nFigure 23.- Concluded.\n\nCONFIDENTIAL\n\nNACA\n\n57\n```", "timestamp": "2026-07-22T04:21:15.056869+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 33, "total_pages": 43, "image_filename": "19930085958_p33.jpg", "text": "```markdown\n32\nNACA RM No. L9B11\n\n<!-- Image (63, 103, 884, 805) -->\n\nFigure 13.- Effects of wing-fuselage position on the characteristics of a\n42° sweptback wing with 0.55b/2 extensible leading-edge flaps, split flaps,\nand upper-surfaces fences.\n```", "timestamp": "2026-07-22T04:21:15.251661+00:00"}
{"citation_id": "19930086105", "source_url": "https://ntrs.nasa.gov/api/citations/19930086105/downloads/19930086105.pdf", "page_number": 5, "total_pages": 22, "image_filename": "19930086105_p5.jpg", "text": "NACA RM E9H12 CONFIDENTIAL 3\n\ntunnel test section, the ambient pressure was approximately 4.0 inches of mercury absolute (pressure altitude of 47,000 ft). The total temperature of the air stream was maintained at $220^\\circ \\pm 5^\\circ$ F and the dew point at $-15^\\circ \\pm 10^\\circ$ F. Schlieren studies were made at a Mach number of 1.91 in the 18- by 18-inch tunnel.\n\nThe 3.6-inch-diameter ram jet (fig. 1) is the same as that of reference 5, except for the outlet-area restriction. Instead of the variable outlet-area configuration, a series of fixed symmetrical nozzles was used to give the desired area ratios. The diffuser was the shock type with a triple-shock projecting cone and a curved inlet.\n\nTwo types of burner configuration were employed in this investigation. The first burner consisted of a perforated conical flame holder (references 5 to 7) used in conjunction with a liquid 62-octane-gasoline injection system and an acetylene-gas pilot. The second burner (fig. 1(b)) was the regenerative type in which the main fuel (propylene oxide) entered a 0.19-inch-diameter tube and discharged into a helical passage, where the addition of heat caused the fuel to flash into a vapor as it expanded through a 3/4-inch-diameter peripheral orifice with a gap setting of 0.005 inch. Combustion was initiated by acetylene pilot flames, which were issuing around the disk at four stations $90^\\circ$ apart. A photograph of the burner in combination with the acetylene pilot is shown in figure 1(c).\n\nInstantaneous pressure data for the experiments with liquid-gasoline injection were indicated by a variable-inductance pressure pickup. This gage, which was flush-mounted to the wall at the combustion-chamber inlet, consisted basically of a twisted Bourdon tube and a variable-inductance air-gap element. The data were recorded on a three-element galvanometer, which also furnished an atmospheric reference line and timing trace.\n\nBecause of a breakdown of the variable-inductance pressure pickup, maximum and minimum pressure data were obtained for the regenerative-type burner by a balanced diaphragm gage. No frequency data were obtained with this gage.\n\nA survey rake of 40 pitot-static-pressure tubes (fig. 1(d)) was located at the diffuser outlet. The pressures indicated on a multitube tetrabromoethane manometer board were photographically recorded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:21:16.633574+00:00"}