PDF/ocr_dataset_1082.jsonl

{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 61, "total_pages": 149, "image_filename": "19930083192_p61.jpg", "text": "NACA TN 1976\n57\n\nAlthough no data have been analyzed with regard to the imposed\ngust loads for the postwar period, on the basis of the current design\nrequirements and if the level of roughness is assumed to remain the\nsame as for the prewar period, the imposed gust acceleration increment\nfor a given number of flight hours will be about 12 percent higher on\nthe average for the postwar period than for the prewar period. Alter-\nnatively, the increased probable speed for the postwar period would\nresult in an appreciable decrease in the flight miles to exceed a given\nacceleration increment.\n\n**Maximum speeds.** - Inspection of figure 49 indicates that the\nprobability of exceeding a given value of airspeed is a function of the\nflight condition. Figure 49 indicates that the descent is most import-\nant, and for one postwar airplane the placard speed might be exceeded\non the average about once in 200 hours in descent. A simple sorting of\nthe data with regard to the smooth air and rough air showed no apparent\neffect of rough air on the speed tendencies.\n\nFor complete flight operations, the data given in table XXI show\nthat the flight miles to exceed the placard speed decreased from many\nmillions of miles during the prewar period to about a quarter of million\nof miles during the postwar period. The reduction in the flight miles\nto exceed the placard speed for the postwar period is of considerable\nsignificance.\n\n**Speed-time distributions.** - Figure 48 shows that a relatively higher\npercentage of time is spent above high speed in level flight than would\nbe expected from estimates made by airline personnel, which estimates\nhave indicated a value of less than 1 percent. The high percentage may\nbe due in part to performance abilities of the airplanes in excess of\nthat indicated by the regulated speeds or due to high-speed descents at\nthe end of flights. Until further information is obtained, the results\nin figure 48 might be assumed to be representative of postwar operations.\n\n**Disturbed motions.** - Telefson (reference 44) showed that the maximum\nangular displacements of an airplane do not correlate with the maximum\nintensity of gusts encountered during a given run. The correlation\ncoefficient between the angular motions and the maximum gust intensity\nwas about 0.3. The disturbed motions may depend more on the sequence of\nthe gusts encountered than on the intensity of any individual gust\nencountered. The data cannot be extended to other flight conditions\nexcept by assuming dynamically similar conditions.\n\nInspection of figure 51(a) indicates that the amplitude of the\nrolling motion is much greater than either pitch or yaw. Similar results\nare indicated in figure 51(b) for the angular velocities. The data of\nfigure 51 can be utilized, if dynamically similar airplanes and a", "timestamp": "2026-07-22T06:56:35.891525+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 31, "total_pages": 92, "image_filename": "19930085951_p31.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n\nNACA RM L9D29\n\nThrust coefficient, $C_T$\n\nAdvance ratio, $J$\n\n$\\beta_{0.75R}$ 20° 25° 30° 35° 40° 45° 50°\n\nNACA\n\n(a) Thrust coefficient.\n\nFigure 9.- Characteristics of NACA 10-(3)(062)-045A propeller. Rotational speed, 1350 rpm.\n\nCONFIDENTIAL\n\n29", "timestamp": "2026-07-22T06:56:39.105908+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 71, "total_pages": 72, "image_filename": "19930085491_p71.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:56:41.388158+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 52, "total_pages": 92, "image_filename": "19930085870_p52.jpg", "text": "```markdown\nNACA RM No. L9D07\n53\n\nCONFIDENTIAL\n\n<!-- Image (65, 110, 922, 904) -->\n\n(a) Wing 1. w = 0.382; R = 1,000,000.\nFigure 7.- Aerodynamic characteristics of 8-percent-thick triangular wings at M = 2.40.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:56:44.180080+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 13, "total_pages": 17, "image_filename": "19930085988_p13.jpg", "text": "```markdown\n12\n\nCONFIDENTIAL\n\n28 x10⁶\n24\n20\n16\n12\n8\n4\n0\nReynolds number\n\nBody alone\nWing-body\n\n0.8 .9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0\nMach number, M\n\nCONFIDENTIAL\nNACA\n\nFigure 4.- Variation of Reynolds number with Mach number for test models. Reynolds numbers based on\nwing mean aerodynamic chord of 1.99 feet.\n\nNACA RM L9H30\n```", "timestamp": "2026-07-22T06:56:45.512526+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 18, "total_pages": 25, "image_filename": "19930085979_p18.jpg", "text": "```markdown\n1130\n\nNACA RM E9E12\n\nRam-pressure recovery, $\\eta$\n\n| Tunnel-air velocity (ft/sec) | |\n|---|---|\n| $\\circ$ | 200 |\n| $\\square$ | 280 |\n| $\\diamond$ | 355 |\n| $\\triangle$ | 435 |\n\n[Figure: A graph plotting Ram-pressure recovery against Cold-gas bleedback. The y-axis ranges from .88 to .98. The x-axis ranges from 1 to 10. Data points for four different tunnel-air velocities are plotted, showing a decreasing trend. A solid line represents the general trend of the data.]\n\nCold-gas bleedback, percent\n\nFigure 4. - Variation of ram-pressure recovery with cold-gas bleedback. Angle of attack, $0^\\circ$.\n\n17\n```", "timestamp": "2026-07-22T06:56:45.801785+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 41, "total_pages": 52, "image_filename": "19930085911_p41.jpg", "text": "40\nCONFIDENTIAL\nNACA RM E9F22\n\n[Figure: NACA logo]\n\nDiffuser Mach number, station 2, $M_2$\nDiffuser total-pressure recovery, $P_4/P_0$\nTime after release, $\\tau$, sec\n\nNo data obtained from 35 1/2 to 49 seconds\n\nImpact\n\n(c) Diffuser variables.\n\nFigure 9. - Continued. Time history of flight data and performance of ram-jet unit 16-A-4.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:56:49.435079+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 34, "total_pages": 47, "image_filename": "19930085919_p34.jpg", "text": "```markdown\nNACA RM No. A9C21\nCONFIDENTIAL\n\n1.6\n1.2\n.8\n.4\nLift coefficient, $C_L$\n0\n-.4\n\n$\\delta_{el}$ deg $\\delta_{el}$ deg\n$\\nabla$ 40 $\\nabla$ -5\n$\\circ$ 30 $\\circ$ -10\n$\\square$ 20 $\\square$ -20\n$\\diamond$ 10 $\\diamond$ -30\n$\\triangle$ 0 $\\triangle$ -40\n\n0.25c\n0.25c elevon\n0.25c\nShort fuselage\n\n$\\delta_{el}$ deg\n40\n30\n20\n10\n0\n-10\n-20\n-30\n-40\n\n$\\delta_{el}$ deg\n-40 -30 -20 -10 0 5 10 20 30 40 $\\delta_{el}$ deg\n\n1.2\n.8\n.4\nLift coefficient, $C_L$\n0\n-.4\n\n$\\delta_{el}$ deg $\\delta_{el}$ deg\n$\\nabla$ 40 $\\nabla$ -5\n$\\circ$ 30 $\\circ$ -10\n$\\square$ 20 $\\square$ -20\n$\\diamond$ 10 $\\diamond$ -30\n$\\triangle$ 0 $\\triangle$ -40\n\nGEN\nconstant-\nchord elevon\n0.25c\nShort fuselage\n\n$\\delta_{el}$ deg\n40\n30\n20\n10\n0\n-10\n-20\n-30\n-40\n\n$\\delta_{el}$ deg\n-40 -30 -20 -10 0 5 10 20 $\\delta_{el}$ deg\n\n-8 0 8 16 24 32 40\nAngle of attack, $\\alpha$, deg\n\n.24 .16 .08 0 -.08 -.16\nPitching-moment coefficient, $C_m$\n\nNACA\n\n(a) $C_L$ vs $\\alpha$ and $C_m$.\nFigure 11- Comparative effectiveness of the 25-percent chord elevon and the constant-\nchord elevon. R, 4.2 x $10^6$.\n\nCONFIDENTIAL\n33\n```", "timestamp": "2026-07-22T06:56:57.458811+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 5, "total_pages": 20, "image_filename": "19930085999_p5.jpg", "text": "NACA RM E9I07\n\nobtained during the investigation of loosely mounted turbine blades. The comparison was based on the effects produced by loose blade-base mounting on natural frequency, response to excitation forces, and damping of vibration in the turbine blades.\n\nIn the analysis of the results, attention was given to the magnitude of the excitation forces. An investigation has been made of the effects produced by varying the magnitude of an excitation force applied to a cantilever blade fastened in a rotor in such a manner as to have a considerable amount of mechanical damping in the base mounting (reference 5). The damping of this type of blade mounting was affected to a considerable extent by the magnitude of the exciting force. It is therefore probable that this factor is important in interpretation of the data obtained from the turbine blades investigated, in which the vibratory-stress level was generally low.\n\nEXPERIMENTAL EQUIPMENT AND PROCEDURE\n\nThe turbojet engine used in this investigation is described in reference 4. The engine is a straight-flow type with a centrifugal compressor, 14 combustion chambers, 48 nozzle vanes, and a single-stage turbine. The turbine blades are unshrouded and are attached in the rim of the turbine wheel with serrated dovetails.\n\nCertain modifications were made to the turbojet engine in order to provide a passage for the lead wires from the strain gages mounted on the turbine blades to the slip-ring unit at the forward end of the engine. Axial holes were bored through the turbine wheel and shaft, the compressor, and the several shafts connecting these components. A hollow auxiliary shaft provided a connection between the forward hub of the compressor and the slip-ring unit mounted on the accessory housing.\n\nModification of the bases of the turbine blades used in the experimental investigation was necessary in order to obtain the desired amount of looseness in fit between the blade bases and the serrations in the turbine-wheel rim. The turbine blades were 4 inches long, and the bases were modified to provide two degrees of looseness. One blade had an amplitude of tip movement in the plane of the wheel of 0.03 inch; a similar blade had an amplitude of movement at the tip of 0.06 inch. These amplitudes were selected as approximately representative of average and maximum looseness for turbine blades of the type used in the investigation.", "timestamp": "2026-07-22T06:56:59.165666+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 62, "total_pages": 149, "image_filename": "19930083192_p62.jpg", "text": "58\nNACA TN 1976\n\nconstant level of atmospheric turbulence are assumed, to estimate the\nfraction of the airplane life flown before displacements or angular\nvelocities of the airplane in excess of selected values are experienced.\n\nRelatively little data have been obtained on angular accelerations\nin rough air and the data for roll are discussed in the section about\ngust structure. Twenty-one values of angular acceleration in pitch were\navailable from flights of the XB-15 airplane and show that the linear\nacceleration increments at the tail were 0 to 40 percent greater than\nthe acceleration increment at the center of gravity. The average\nincrease was about 25 percent. If dynamically similar airplanes are\nassumed, the results can be expressed as an average value of the angular\nacceleration in pitch for any airplane as\n\n$$\n\\frac{dq}{dt} = \\frac{20 \\Delta n}{c^{3/2}}\n$$\n\nAt present exact estimations of the angular acceleration in pitch are\nnot possible.\n\nThe effect of disturbed motions on imposed loads is appreciable\nand may affect the frequency distribution by as much as 50 percent\n(fig. 50). A previously mentioned analysis of piloting effects that\ncould be construed as disturbed motion indicated a variation of some\n20 percent in the imposed loads. Although such variations cannot be\npredicted at this time, current methods of estimating loads consider\nthis factor in that the measured flight load experience includes such\neffects.\n\nR É S U M É\n\nA résumé of the more important findings under gust structure, air-\nplane reactions, and operating statistics is included because of the\nvariety of subjects covered.\n\nGust Structure\n\nThe gust-gradient distance and gust size are largely independent of\nweather, altitude, and airplane when expressed in terms of the mean\ngeometric chord. The gradient distance for a vertical gust has a\nprobable value of 10 to 14 chords for the higher gust intensities and is\nthe same for both spanwise and flight directions.", "timestamp": "2026-07-22T06:57:06.482460+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 53, "total_pages": 92, "image_filename": "19930085870_p53.jpg", "text": "54\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O CL\n                {□ Cm\n.16\nWedge L.E. {△ CL\n           {◇ Cm\n.08\nCL\n0\n-.08\n-.16\n-.24\n.01\nCm\n0\n-.01\n\n.06\nElliptical L.E. {O CD\n                {□ L/D\n.04\nCD\n.02\n0\n-8 -6 -4 -2 0 2 4 6 8\nα, deg\nWedge L.E. {△ CD\n           {◇ L/D\n6\n4\nL/D\n2\n0\nNACA\n\n(b) Wing 2. w=0.705; R=1,000,000.\nFigure 7.- Continued\nCONFIDENTIAL", "timestamp": "2026-07-22T06:57:09.111486+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 14, "total_pages": 17, "image_filename": "19930085988_p14.jpg", "text": "NACA RM L9H30\n\nCONFIDENTIAL\n\n$C_D$\n\n.03\n.02\n.01\n0\n-.01\n\n.8 .9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0\nM\n\nCONFIDENTIAL\n\nTotal $\\begin{cases} \\text{Wing-body, 2 fins} \\\\ \\text{Body, 4 fins} \\end{cases}$\n\nBase $\\begin{cases} \\text{Body, 4 fins} \\\\ \\text{Wing-body, 2 fins} \\end{cases}$\n\n$\\circ$ Doppler\n$\\square$ Telemeter\n\nNACA\n\nFigure 5.- Variation of total- and base-drag coefficients with Mach number for the test models.\n$C_D$ referred to wing-plan-form area.\n\n13", "timestamp": "2026-07-22T06:57:11.236405+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 28, "total_pages": 34, "image_filename": "19930085913_p28.jpg", "text": "NACA RM L9F24\n27\n\nNo weight, -19.44%\n0.0%\n5.55%\n11.11%\n16.66%\n22.20%\n27.77%\n33.33%\n38.90%\n44.40%\n\n50.00%\n55.50%\n61.11%\n66.66%\n72.20%\n77.80%\n83.30%\n88.90%\n94.40%\n100.00%\n\n2nd natural frequency nodal lines\n3rd natural frequency nodal lines\nNACA\n\n(c) Swept, untapered wing; $\\Lambda = 60^\\circ$, $e_w = -1$.\n\nFigure 2.- Concluded.", "timestamp": "2026-07-22T06:57:13.327623+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 19, "total_pages": 25, "image_filename": "19930085979_p19.jpg", "text": "18\nNACA RM E9E12\n\nCold-gas\nbleedback\n(percent)\nO 0\n□ 9.5\n\n[Figure: A graph showing pressure distribution on an inlet-lip. The vertical axis is labeled \"Pressure coefficient, s\" with values .0, .3, .6, .9. The graph contains two curves with data points marked by circles and squares.]\n\nNACA\n\nFigure 5. - Effect of cold-gas bleedback on inlet-lip pressure distribution.\n\n1130", "timestamp": "2026-07-22T06:57:14.922947+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 12, "total_pages": 40, "image_filename": "19930085997_p12.jpg", "text": "10 CONFIDENTIAL NACA RM A9I29\n\n3. The mass-flow ratio $m_4/m_0$, required at full scale to produce the same pressure recoveries as were recorded in the model tests, was equal to\n\n$$(m_4/m_0)_{\\text{full scale}} = \\frac{(m_4/m_0)_{\\text{model}}}{[(Rx)_{\\text{full scale}}/(Rx)_{\\text{model}}]^{1/5}}$$\n\nThe results of applying these assumptions regarding the influence of the model scale to the previous calculations are also shown in figure 11. In this case the energy required to remove the boundary layer was equivalent to a decrease in the total pressure ratio $H_2/H_0$ of approximately 0.04.\n\nAngle of Attack\n\nConfiguration D.- The variation of total-pressure recovery with mass flow is shown in figure 12 for configuration D at angle of attack. An examination of these results indicates that $(H_3/H_0)_{\\text{max}}$ decreased with increasing positive angles of attack, and that the range of mass-flow ratios in which large values of $H_3/H_0$ could be maintained was markedly reduced. The schlieren photographs of figure 13 indicate that the thickness of the boundary layer along the forebody ahead of the scoop increased as the angle of attack of the model was increased.¹ Thus the boundary-layer scoops were apparently inadequate in handling this thicker boundary layer. The greater boundary-layer thickness at angle of attack may have been caused by a secondary flow in the boundary layer due to the pressure difference between the upper and lower surfaces of the forebody. The low values of the total-pressure recovery measured at position 3 probably are greatly influenced by the presence of this boundary layer in the subsonic diffuser at angle of attack. In all tests of the model at angle of attack, the mass flow in the boundary-layer scoops was maintained at the choked condition.\n\nForebody incidence.- To improve the pressure recovery at angle of attack the forebody was drooped 2° and 6° with respect to the inlets. (See fig. 2.) Maximum values of $H_3/H_0$ as functions of $M_0$ are presented in figure 14 for configurations E and F at various angles of attack. The values of $(H_3/H_0)_{\\text{max}}$ measured in tests of configuration D are also shown for purposes of comparison. Improvements in $(H_3/H_0)_{\\text{max}}$ occurred in tests of E and F; however, the total-pressure recovery at 6° and 9° angle of attack still was low when compared to that at 0°. In figure 14, configuration F at 6° angle of attack did not give the same pressure recovery as configuration D at 0° angle of attack due to the fact that\n\n---\n\n¹In these photographs the model is both at an angle of attack and sideslip because it was necessary to rotate the model about its longitudinal axis in order to photograph the boundary layer on the forebody ahead of one scoop.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:57:15.227026+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 16, "total_pages": 46, "image_filename": "19930085983_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:57:17.647577+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 42, "total_pages": 52, "image_filename": "19930085911_p42.jpg", "text": "NACA RM E9F22\nCONFIDENTIAL\n41\n\n[Figure: A multi-panel graph showing time history of flight data. The x-axis is \"Time after release, $\\tau$, sec\" ranging from 0 to 50. The y-axes represent various combustion-chamber-inlet variables. The graph contains four panels stacked vertically. The top panel shows \"Combustion-chamber-inlet static temperature, $t_4$, °R\" from 400 to 700. The second panel shows \"Combustion-chamber-inlet static pressure, $P_4$, lb/sq ft\" from 0 to 3000. The third panel shows \"Combustion-chamber-inlet velocity, $V_4$, ft/sec\" from 0 to 300 and \"Combustion-chamber-inlet Mach number, $M_4$\" from 0 to .20. The bottom panel shows \"Fuel-air ratio, $W_f/W_a$\" from .06 to .14. There is a vertical dashed line at approximately 35 seconds with a note: \"No data obtained from 35 1/2 to 49 seconds\". An arrow labeled \"Impact\" points to the 50-second mark on the x-axis.]\n\n(d) Combustion-chamber-inlet variables.\nFigure 9. - Continued. Time history of flight data and performance of ram-jet unit 16-A-4.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:57:18.330910+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 35, "total_pages": 47, "image_filename": "19930085919_p35.jpg", "text": "34\n\nCONFIDENTIAL\n\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\nLift coefficient, $C_L$\n\n0 .08 .16 .24 .32 .40 .48 .56 .64 .72 .80 .88 .96 1.04\nDrag coefficient, $C_D$\n\n(b) $C_L$ vs $C_D$.\nFigure 11.- Continued.\n\n$\\delta_f$, deg\n$\\Delta$ 20\n$\\circ$ 0\n$\\triangle$ -20\n$\\triangledown$ -40\n\n0.25c\n-0.25c aileron\n-0.25c\nShort fuselage\n\n$\\delta_f$, deg\n$\\Delta$ 20\n$\\circ$ 0\n$\\triangle$ -20\n$\\triangledown$ -40\n\n0.62c\nSpoilers-\nshort aileron\n-0.25c\nShort fuselage\n\nNACA\n\nNACA RM No. A9C21\nCONFIDENTIAL", "timestamp": "2026-07-22T06:57:25.832679+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 100, "total_pages": 114, "image_filename": "19930086061_p100.jpg", "text": "96\nNACA RM L9J07\n\n$$ \\frac{c_l c}{C_L C_{av}} $$\nTheoretical load distribution\n$$ \\alpha, \\text{deg} \\quad C_L $$\n$$ \\square \\quad 4.1 \\quad 0.13 $$\n$$ \\circ \\quad 8.1 \\quad 0.29 $$\n$$ \\frac{y}{b/2}, \\text{ percent} $$\n(a) Angles of attack : 4.1°, 8.1°.\n\n$$ \\frac{c_l c}{C_L C_{av}} $$\n$$ \\alpha, \\text{deg} \\quad C_L $$\n$$ \\diamond \\quad 14.1 \\quad 0.50 $$\n$$ \\triangle \\quad 24.1 \\quad 0.83 $$\n$$ \\triangledown \\quad 34.1 \\quad 1.12 $$\n$$ \\frac{y}{b/2}, \\text{ percent} $$\n(b) Angles of attack : 14.1°, 24.1°, 34.1°.\n\n$$ \\frac{c_l c}{C_L C_{av}} $$\n$$ \\alpha, \\text{deg} \\quad C_L $$\n$$ \\triangledown \\quad 39.1 \\quad 1.17 $$\n$$ \\square \\quad 44.1 \\quad 0.94 $$\n$$ \\diamond \\quad 50.1 \\quad 0.62 $$\n$$ \\frac{y}{b/2}, \\text{ percent} $$\n(c) Angles of attack : 39.1°, 44.1°, 50.1°.\n\nFigure 45.- Span load distribution of wing 3 at various angles of attack;\n$$ \\psi = 0^\\circ $$. All data are taken over left semispan.", "timestamp": "2026-07-22T06:57:26.613032+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 15, "total_pages": 132, "image_filename": "19930085990_p15.jpg", "text": "NACA RM A9I01 CONFIDENTIAL 13\n\nof the fuselage is included in the value of the downwash computed from the data.\n\nEfficiency of the horizontal tail.— The tail efficiency factor $\\eta(q_t/q)$ computed from the force and moment data is presented in figures 35 and 36. The tail efficiency factor, defined as the ratio of the lift produced by the tail in the presence of the fuselage to the lift produced by the isolated tail operating at the same Mach number, was computed by means of the following expression:\n\n$$\n\\eta \\frac{q_t}{q} = \\left( \\frac{\\partial C_m}{\\partial i_t} \\right)_\\alpha \\frac{1}{(dC_L/d\\alpha)_t (S_t l_t / Sc)}\n$$\n\nwhere $(dC_L/d\\alpha)_t$ is the lift-curve slope of the isolated horizontal tail operating at the free-stream Mach number of the horizontal tail (figs. 5 and 32). No attempt was made to separate the effects of dynamic-pressure ratio at the tail from the tail efficiency due to the possible large variation of $q_t/q$ along the tail span. The tail efficiency factor is presented as a function of Mach number in figure 35. For either position of the tail with the flaps neutral, the tail efficiency factor was less than 80 percent and varied approximately 10 percent over the test range of Mach numbers and angles of attack.\n\nThe Effects of Compressibility\n\nThe effects of compressibility on the lift, drag, pitching moment, and downwash of the complete model are summarized in figures 37 through 46.\n\nLift and drag.— The variation with Mach number of the angle of attack for a constant lift coefficient was small (fig. 37), increasing Mach number usually being accompanied by a decrease in the angle of attack for a given lift coefficient.\n\nThe variation with Mach number of the drag coefficient for several constant lift coefficients is shown in figure 38. At a lift coefficient of zero, the drag coefficient of the model with the tail in the extended wing-chord plane started to increase at a Mach number of about 0.80. For the model with the high tail, the drag increase started at a Mach number of about 0.75. The Mach number for drag divergence, defined as the Mach number at which $(\\partial C_D/\\partial M)_{C_L=0} = 0.10$, was approximately 0.92 for the model with either tail position.\n\nStatic longitudinal stability and control.— The variation with Mach number of the pitching-moment coefficient for several constant lift coefficients is shown in figure 39. In general, the pitching-moment coefficient increased with increasing Mach number. The static longitudinal\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:57:29.316781+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 72, "total_pages": 72, "image_filename": "19930085491_p72.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 71\n\n[Figure: Schlieren photograph of a wing model with shock waves visible, labeled (i) WF-70; C_L, 0.]\n\n[Figure: Schlieren photograph of a wing model with shock waves visible, labeled (j) WF-70: C_L, 0.21.]\n\nFigure 16.— Concluded.\n\n[Figure: Schlieren photograph of a wing tip, labeled (a) R, 0.31 million.]\n\n[Figure: Schlieren photograph of a wing tip with NACA logo and number A-12339, labeled (b) R, 0.62 million.]\n\nFigure 17.— Schlieren photographs of WF-63 wing-tip flow patterns at zero lift.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:57:32.616476+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 8, "total_pages": 32, "image_filename": "19930085992_p8.jpg", "text": "6\nNACA RM L9E17\n\nWeight 3 (fig. 3(g)) and weight 4 (fig. 3(h)) were the streamlined and the nonstreamlined bodies, respectively, used to pursue further the investigation of effects of aerodynamic shape. The weights were made so that they could be varied over a wide range of spanwise positions with their inertial properties remaining constant. The center of gravity of each weight was located close to the elastic axis of the wing.\n\nThe concentrated weight properties are given in the following table, in which the negative values of $e_w$ indicate weight locations forward of the wing elastic axis:\n\n| Weight | $\\frac{W_w}{W}$ | $e_w$ | $\\frac{I_w}{I_{EA}}$ |\n| :--- | :--- | :--- | :--- |\n| 1a | 0.0744 | 0.045 | 0.380 |\n| 1b | .0754 | -.513 | .509 |\n| 2a | .0755 | .045 | .403 |\n| 2b | .0755 | -.505 | .507 |\n| 2c | .0752 | .045 | .219 |\n| 2d | .0749 | .045 | .070 |\n| 3 | .0485 | .023 | .310 |\n| 4 | .0485 | .023 | .308 |\n\nThe flutter tests described herein were conducted in the Langley 4.5-foot flutter research tunnel, the essential features of which are discussed in reference 1.\n\nStrain gages mounted on the wing near the root, as shown in figure 1, permitted vibration records to be made of the bending and torsional oscillations of the wing during flutter. The square indicates the location of the bending gages and the circles indicate the locations of the torsion gages. The strain-gage signals were recorded on a recording oscillograph.", "timestamp": "2026-07-22T06:57:34.848376+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 63, "total_pages": 149, "image_filename": "19930083192_p63.jpg", "text": "NACA TN 1976\n59\n\nThe gust shapes vary widely but triangular or sinusoidal profiles in the direction of flight seem good approximations.\n\nAn approximation to the spanwise gust profile for unsymmetrical gusts would be a uniform distribution on which is superimposed a linear variation of gust velocity with equal and opposite gust intensities at each wing tip.\n\nGust spacing for use in estimating the number of gusts in rough air as a function of gust intensity is about 11 chords for velocities above a threshold of 0.3 foot per second. The spacing is independent of other factors.\n\nLateral, longitudinal, and vertical gusts have essentially the same characteristics.\n\nThe frequency distribution of gust intensities in rough air is independent of altitude and airplane when the gust intensity is expressed in terms of indicated velocity.\n\nAirplane Reactions\n\nAerodynamics.- Although the two-dimensional unsteady-lift functions yield the most satisfactory estimates of the load increments due to gusts for conventional airplanes with fuselages, finite-aspect-ratio unsteady-lift functions appear pertinent for flying wings. For unconventional configurations such as swept wings, strip theory with two-dimensional unsteady-lift functions appears adequate.\n\nThe slope of the lift curve for gust-load calculation can be best estimated for engineering purposes by using an arbitrary section value of 6 per radian and simple aspect-ratio corrections. For swept wings, the best estimate of lift-curve slope is obtained by correcting the value for the equivalent straight wing by the cosine of the sweep angle.\n\nAvailable information on compressibility effects indicates that the unsteady-lift functions would be only slightly affected below the critical Mach number and would tend to disappear for supersonic speeds. At this time, the use of high-speed wind-tunnel data is suggested for estimating the slope of the lift curve.\n\nRigid-body reactions.- Analysis indicates that the mass parameter is a significant measure of airplane reactions and that the use of net wing area results in best agreement between calculation and experiment. Analyses neglecting unsteady-lift effects are not considered adequate.", "timestamp": "2026-07-22T06:57:37.594573+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 6, "total_pages": 20, "image_filename": "19930085999_p6.jpg", "text": "```markdown\n4\nNACA RM E9I07\n\nThe turbine blades were originally intended to be mounted\ntightly and then positioned positively with respect to axial posi-\ntion in the turbine wheel by peening the blade-base dovetail. The\nremoval of sufficient material from the blade-base serrations to\nprovide increased clearance in the attachment made it necessary to\nemploy a different method of retaining the blade in the wheel. The\nprincipal requirement of the fastening was positive positioning with\nminimum restriction to movement of the turbine blade in the plane of\nthe wheel and within the limits of the desired blade-tip amplitude.\nIt was also considered necessary that the method of attachment per-\nmit removal and replacement of experimental blades without a major\ndisassembly of the engine. The modifications made to the turbine-\nwheel rim and to the dovetail sections of the experimental turbine\nblades and the blade-base modification made to accommodate strain-\ngage lead wires are shown in figure 1.\n\nWhen the turbine blades were assembled in the wheel, a posi-\ntioning guide and a hardened-steel sphere were placed in that part\nof the wheel passage having the larger diameter; a modified blade\nwas then inserted in the wheel dovetail and the adjustment screw\ntightened until the sphere was seated in the hemispherical cavity\nmachined in the blade dovetail. The adjustment screw was then\nloosened until there was no appreciable restraint to blade movement\nin the plane of the wheel. When the adjustment was satisfactory,\nthe screw was locked in position.\n\nHigh-temperature resistance-wire strain gages were used to\nobtain data from the loosely mounted turbine blades during service\noperation of the engine. The construction and the mounting of the\nstrain gages were similar to those of the multiple-loop type,\ndescribed in references 4 and 6, with some improvements that have\nresulted from continued research on high-temperature strain gages.\nThe strain-sensitive wire was a platinum-iridium alloy; Sauereisen\nNo. 32 cement was used as the mounting material. A precoat of a\nhigh-temperature ceramic was fired on the surface of an experimental\nblade before the strain gage was mounted. This ceramic coating,\nL-6AC, is a development of the National Bureau of Standards. After\nthe strain gage had been baked on a turbine blade at low temperature,\nthe blade was placed in a high-temperature oven and heated suffi-\nciently to stabilize the strain-sensitive characteristics of the wire\nand to complete the bonding of the cement. The strain gages were\napplied near the blade bases along the trailing edges on the convex\nside.\n\nA photograph (fig. 2) of one of the instrumented turbine blades\nshows the precoat, the location of the high-temperature strain gage,\nand the ceramic-lined Inconel conduit enclosing the strain-gage lead\n\n1191\n```", "timestamp": "2026-07-22T06:57:39.573764+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 20, "total_pages": 30, "image_filename": "19930085975_p20.jpg", "text": "18\nNACA RM L9E10\n\nCONFIDENTIAL\n\nRolling-moment coefficient, $C_l$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:57:48.317073+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 15, "total_pages": 17, "image_filename": "19930085988_p15.jpg", "text": "14\n\nCONFIDENTIAL\n\n$C_{Pb}$\n\nWing-body (two fins)\n\nBody alone (four fins)\n\nCONFIDENTIAL\n\nNACA\n\nM\n\nFigure 6.— Variation of base-pressure coefficient with Mach number for the test models.\n\nNACA RM L9H30", "timestamp": "2026-07-22T06:57:53.389110+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 36, "total_pages": 47, "image_filename": "19930085919_p36.jpg", "text": "```markdown\nNACA RM No. A9C21\n\nCONFIDENTIAL\n\nRolling-moment coefficient, $C_l$\n\n$\\delta_e$, deg\n-40\n-30\n-20\n-10\n-5\n5\n10\n20\n30\n40\n\n25-percent-chord elevon\n\nConstant-chord elevon\n\nNACA\n\nAngle of attack, $\\alpha$, deg\n\n(c) $C_l$ vs $\\alpha$.\n\nFigure 11.- Concluded.\n\nCONFIDENTIAL\n\n35\n```", "timestamp": "2026-07-22T06:58:09.585547+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 64, "total_pages": 149, "image_filename": "19930083192_p64.jpg", "text": "60\nNACA TN 1976\n\nNo one significant parameter of airplane stability for flight in\ngusty air has been found. The center-of-gravity position is an important\nfactor for a given airplane and arrangement and tail volume have some-\nwhat less effect. Comparison of experiment and calculation shows that\nminor differences in the constants used in calculation can vary the agree-\nment from good to poor. Detailed calculations of airplane reactions for\nestimating loads are not warranted at this time because the accuracy\nrequired is beyond the scope of present information.\n\nAvailable data on horizontal and vertical tail loads indicate that\ndetailed calculations of tail load are not warranted. The lift on\neither tail surface due to a gust can be estimated apparently by\nutilizing a true gust velocity and neglecting the alleviating effects of\nunsteady lift and airplane motion but considering the downwash.\n\nElastic-airplane reactions.- Analytical and experimental studies\nindicate that extensive simplifications lead to unrealistic answers and\nmore exact solutions are required as the wing mass becomes a greater\nfraction of the total mass.\n\nThe effect of forward speed on elastic response was negligible,\nbut the wing-tip accelerations increased with speed.\n\nIt does not appear feasible to obtain generalized solutions of\ndynamic response.\n\nOperating Statistics\n\nThe average miles to exceed limit load factor varies widely accord-\ning to operator and period. A most important factor in setting the level\nof load appears to be the probable speed, that is, the speed at which\nmaximum acceleration is most likely to be experienced, since the gust\nexperience on all routes was about the same. The probable-speed ratio\n(ratio of probable speed to the high speed in level flight) is increasing\nas new airplanes are introduced with a consequent reduction in the\naverage number of flight miles required to exceed limit acceleration\nincrement.\n\nThe flight miles to exceed the placard speed has decreased for the\nnewer types of airplanes.\n\nThe total number of gusts encountered by an airplane during its\nlife is a function of the amount of rough air flown. The amount of\nrough air is on the average about 10 percent of the total mileage. The\nnumber of gusts encountered in rough air depends on airplane size and\nan average figure would be $500/\\sqrt{6}$ gusts per mile greater than 0.3 foot\nper second.", "timestamp": "2026-07-22T06:58:15.258855+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 21, "total_pages": 30, "image_filename": "19930085975_p21.jpg", "text": "NACA RM L9E10\n19\n\nCONFIDENTIAL\n\n<!-- Image (280, 111, 747, 299) -->\n\n<!-- Image (280, 354, 747, 548) -->\n\n<!-- Image (280, 597, 747, 884) -->\n\nFigure 8.- The variation with Mach number of the rolling-moment characteristics of a wing for various aileron deflections at several angles of attack. $\\Lambda_{c/4} = 46.7^\\circ$.", "timestamp": "2026-07-22T06:58:15.966062+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 101, "total_pages": 114, "image_filename": "19930086061_p101.jpg", "text": "NACA RM L9J07\n97\n\n$$ \\frac{c_l c}{C_{L_{C_{av}}} 8} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 4.1 | 0.13 |\n| 8.1 | 0.26 |\n\n(a) Angles of attack : 4.1°, 8.1°.\n\n$$ \\frac{c_l c}{C_{L_{C_{av}}} 8} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 14.1 | 0.44 |\n| 24.1 | 0.78 |\n| 34.1 | 1.05 |\n\n(b) Angles of attack : 14.1°, 24.1°, 34.1°.\n\n$$ \\frac{c_l c}{C_{L_{C_{av}}}} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 39.1 | 1.10 |\n| 44.1 | 1.00 |\n| 50.1 | 0.56 |\n\n(c) Angles of attack : 39.1°, 44.1°, 50°\n\nFigure 46.- Span load distribution of wing 3 at various angles of attack; $\\psi = 10^\\circ$. Flagged symbols represent data taken with left semispan at $\\psi = -10^\\circ$.", "timestamp": "2026-07-22T06:58:16.468686+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 16, "total_pages": 17, "image_filename": "19930085988_p16.jpg", "text": "NACA RM L9H30\n\nCONFIDENTIAL\n\nReference 3\nPresent test\n\n$C_D$\n.02\n.01\n0\n.8 .9 1.0 1.1 1.2 1.3 1.4 1.5 1.6\nM\n\nCONFIDENTIAL\nNACA\n\nFigure 7.- Variation of wing-plus-interference drag coefficient with Mach number. $C_D$ based on wing-plan-form area.\n\n15", "timestamp": "2026-07-22T06:58:16.837901+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 32, "total_pages": 92, "image_filename": "19930085951_p32.jpg", "text": "```markdown\n30\n\nCONFIDENTIAL\nUNCLASSIFIED\n\nPower coefficient, $C_P$\n.48\n.44\n.40\n.36\n.32\n.28\n.24\n.20\n.16\n.12\n.08\n.04\n0\n\n$\\beta_{0.75R}$ 20° 25° 30° 35° 40° 45° 50°\n\nAdvance ratio, J\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0 3.2 3.4 3.6 3.8\n\nNACA\n\n(b) Power coefficient.\nFigure 9.— Continued. Rotational speed, 1350 rpm.\nCONFIDENTIAL\nUNCLASSIFIED\n\nNACA RM L9D29\n```", "timestamp": "2026-07-22T06:58:18.481216+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 29, "total_pages": 34, "image_filename": "19930085913_p29.jpg", "text": "28\nNACA RM L9F24\n\n<!-- Image (117, 110, 813, 435) -->\n\n(a) Model A; $\\Lambda = 0^\\circ$; $e_w = -1$.\n\n<!-- Image (117, 493, 813, 820) -->\n\n(b) Model A; $\\Lambda = 0^\\circ$; $e_w = 0$.\n\nFigure 3.— Variation of first three natural frequencies and flutter frequency with weight position for the various models tested.", "timestamp": "2026-07-22T06:58:23.010031+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 16, "total_pages": 132, "image_filename": "19930085990_p16.jpg", "text": "14 CONFIDENTIAL NACA RM A9I01\n\ninstability at Mach numbers above about 0.85 of the model with the tail in the extended wing-chord plane, as mentioned previously, is evident from the data of figure 39(a).\n\nThe variation with Mach number of the effective angle of downwash at several constant values of the lift coefficient is shown in figure 40, and the variation of $\\partial \\epsilon / \\partial \\alpha$ with Mach number is shown in figure 41. For either location of the horizontal tail, $\\partial \\epsilon / \\partial \\alpha$ increased with increasing Mach number but the value of $\\partial \\epsilon / \\partial \\alpha$ and the rate of increase with Mach number was much larger for the model with the tail in the extended wing-chord plane. The static longitudinal instability at high subsonic Mach numbers with the tail in the extended wing-chord plane was principally a result of this large value of $\\partial \\epsilon / \\partial \\alpha$.\n\nThe variation with Mach number of the lift coefficient for balance about the quarter point of the wing mean aerodynamic chord is presented in figure 42 for various angles of stabilizer setting. The model with the tail in the extended wing-chord plane was neutrally stable at a Mach number of 0.86 and unstable at higher Mach numbers when the stabilizer setting was $0^\\circ$. With a stabilizer setting of $-1^\\circ$ or $-2^\\circ$ the model was longitudinally stable, but the lift coefficient for balance varied erratically with Mach number at Mach numbers above about 0.70.\n\nWith the tail mounted above the extended wing-chord plane, the model possessed static longitudinal stability at all stabilizer settings and all Mach numbers. For positive values of lift coefficient, the balanced lift coefficient for a given stabilizer angle increased as the Mach number was increased to about 0.90 and decreased with further increase in the Mach number.\n\nFor the model with either position of the horizontal tail, the all-movable stabilizer required only $4^\\circ$ to $6^\\circ$ of deflection to balance the model at the stall with the flaps up.\n\nThe experimental results of this investigation have been used to predict the static longitudinal-stability-and-control characteristics of a hypothetical airplane with a wing loading of 100 pounds per square foot in flight at an altitude of 10,000 feet. The airplane center of gravity has been assumed to be on an axis perpendicular to the plane of symmetry passing through the quarter point of the wing mean aerodynamic chord. The variation of airplane lift coefficient with Mach number for several values of normal-acceleration factor is presented in figure 43. The calculated effects of flight-path curvature on the flow at the tail were negligible for the assumed flight condition.\n\nThe variation with Mach number of the stabilizer angle required to balance the airplane is shown in figure 44 for several values of normal-acceleration factor. With the horizontal tail in the extended wing-chord plane, the airplane would be longitudinally unstable with a normal-\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:58:27.300493+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 20, "total_pages": 25, "image_filename": "19930085979_p20.jpg", "text": "NACA RM E9E12\n19\n\nPlenum-chamber-gas pressure, lb/sq ft absolute\nTunnel-air velocity, ft/sec\n\n[Figure: Graph showing relation between plenum-chamber-gas pressure and tunnel-air velocity with data points and a trend line. NACA logo in bottom right corner of graph.]\n\nFigure 6. - Relation between plenum-chamber-gas pressure and tunnel-air velocity corresponding to optimum temperature distribution in model. Bleedback, 4.9 percent; plenum-chamber-gas temperature, $1000^\\circ$ F; average model-air-temperature rise, $50^\\circ$ F.", "timestamp": "2026-07-22T06:58:27.452716+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 43, "total_pages": 52, "image_filename": "19930085911_p43.jpg", "text": "42\nCONFIDENTIAL\nNACA RM E9F22\n\nOutlet static pressure, $P_7$, lb/sq ft\n3000\n2000\n1000\n0\nNo data obtained from 33 1/2 to 49 seconds\n\nNet-thrust coefficient, $C_F$\n.40\n.20\n0\n-.20\n\nGas total-temperature ratio, $T_7/T_0$\n5.0\n3.0\n1.0\n\nCombustion efficiency, $\\eta_b$, percent\n40\n20\n0\nNACA\n\nTime after release, $\\tau$, sec\n10\n20\n30\n40\n50\nImpact\n\n(e) Performance variables.\n\nFigure 9. - Concluded. Time history of flight data and performance of ram-jet unit 16-A-4.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:58:28.218678+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 21, "total_pages": 30, "image_filename": "19930085970_p21.jpg", "text": "NACA RM A9E09 CONFIDENTIAL 19\n\n[Figure: (e) M = 1.24, plan view.]\n\n[Figure: (f) M = 1.24, side view.]\n\n[Figure: (g) M = 1.51, plan view.]\n\n[Figure: (h) M = 1.51, side view.]\n\nFigure 4.— Concluded.\n\nCONFIDENTIAL\n\nNACA\nA-13972", "timestamp": "2026-07-22T06:58:28.540318+00:00"}
{"citation_id": "19930086020", "source_url": "https://ntrs.nasa.gov/api/citations/19930086020/downloads/19930086020.pdf", "page_number": 1, "total_pages": 22, "image_filename": "19930086020_p1.jpg", "text": "NACA RM A9J06\n2353\n\nCONFIDENTIAL\nCopy 185\nRM A9J06\n\nNACA\nCASE FILE\nCOPY\n\nRESEARCH MEMORANDUM\n\nTHE EFFECTS OF SCALE AND TEST TECHNIQUE ON THE VALIDITY\nOF SMALL-SCALE MEASUREMENTS OF THE AERODYNAMIC\nCHARACTERISTICS OF A WING WITH THE LEADING\nEDGE SWEPT BACK $63^\\circ$\n\nBy L. Stewart Rolls\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\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 9, 1949\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: NACA RESEARCH ABSTRACT NO. 121\nEFFECTIVE DATE: OCTOBER 14, 1957\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:58:31.419429+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 9, "total_pages": 32, "image_filename": "19930085992_p9.jpg", "text": "NACA RM L9E17\n\nIn order to prevent destruction of the wing as a result of divergence, restraining wires were attached from the tunnel walls to the wing quarter chord near the tip. As can be seen in figure 1, these wires had sufficient slack in them to permit adequate amplitude in flutter but could still save the wing if divergence occurred.\n\nTEST PROCEDURE\n\nIn the flutter testing of the model, velocity of flow in the tunnel was increased slowly until the critical flutter speed was attained. At this point, the tunnel conditions were observed and, simultaneously, an oscillograph record of the vibrations of the model was taken. These data, from which the experimental flutter speed and flutter frequency were obtained, have been recorded in table I. For most runs, the natural frequencies were tabulated both before and after the actual run to determine whether or not the wing had been damaged by flutter. The remarks in table I regarding the flutter characteristics are based almost entirely on visual observations made at the time of the run; because of the sudden and violent occurrence of flutter, these remarks are inclined to be somewhat arbitrary. The structural damping coefficients recorded in table I have been determined from the rate of decay of oscillations on the vibration records of the natural frequencies.\n\nRESULTS AND DISCUSSION\n\nIn presenting the results of this investigation, three phases of the problem are considered: first, some effects of aerodynamic shape of concentrated weights; second, effects of variation in the spanwise position of light concentrated weights; third, effects of moment of inertia of light concentrated weights. The second and third phases are included as logical outgrowths of this program and may be regarded as incidental to the first phase, which is concerned with the primary objective of the paper. As in reference 1, the variations in flutter speed and flutter frequency, the two flutter parameters studied, have been compared to the corresponding parameters of the unweighted wing.\n\nEffects of Aerodynamic Shape of Concentrated Weights\n\nAttention may first be directed to the effect on flutter speed and flutter frequency of the aerodynamic shape of concentrated weights", "timestamp": "2026-07-22T06:58:32.096597+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 22, "total_pages": 55, "image_filename": "19930085966_p22.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9B17\n\n[Figure: A cutaway view of a combustion-chamber shell and burner assembly. The left half shows the internal structure with ribs, tubes, and wiring. The right half shows the exterior surface. A ruler is placed on top for scale. A NACA logo with the identifier L-31970 is visible in the bottom right corner of the image.]\n\nFigure 1.— Combustion-chamber shell and burner assembly.\n\nCONFIDENTIAL\n\n21", "timestamp": "2026-07-22T06:58:34.660038+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 17, "total_pages": 46, "image_filename": "19930085983_p17.jpg", "text": "Equation for fuselage ordinates:\n\n$$\n\\frac{r}{r_{max}} = \\left[ 1 - \\left( 1 - \\frac{2x}{l} \\right)^2 \\right]^{\\frac{3}{4}}\n$$\n\nNote: All dimensions given in feet unless otherwise specified.\n\nFineness ratio; $\\frac{l}{2r_{max}} = 12.5$\n\n[Figure: Diagram showing dimensions of wing and fuselage with various measurements and angles labeled]\n\nCONFIDENTIAL\n\n$l = 6.375$\n\n$5.036$\n\n$\\frac{l}{2} = 3.186$\n\n$1.768$\n\n$1.472$\n\n$1.714$\n\n$63^\\circ$\n\n$2r_{max} = .510$\n\n$\\bar{c} = 1.200$\n\n$.750$\n\n$.938$\n\n$.268$\n\n$.429$\n\n$3.750$\n\nCONFIDENTIAL\n\nNACA RM A9I27\n\nFigure 2.— Dimensions of wing and fuselage.\n\nNACA\n\n15", "timestamp": "2026-07-22T06:58:34.863978+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 7, "total_pages": 20, "image_filename": "19930085999_p7.jpg", "text": "NACA RM E9I07\n\nwires. The turbine blades were inserted and fastened in the wheel by the method previously described, and the lead-wire conduit was attached to the surface of the turbine wheel by small metal straps spot-welded to the wheel. The lead wires were then attached to a terminal plate at the center of the turbine wheel.\n\nLead wires were run from the terminal plate, through the engine, and attached to the slip-ring rotor on which were mounted the inactive arms of the Wheatstone bridges. Brushes bearing on the slip rings carried current to the strain-gage bridges and transmitted the bridge output signal to the stationary instrumentation in the control room.\n\nThe experimental procedure consisted in operating the engine in a pendulum-type, sea-level test stand over the entire range of turbine speeds at exhaust temperatures similar to those that occur during regular service operation. The speed ranged from idling (4000 rpm) to full turbine speed (11,500 rpm). As the speed was slowly increased, the strain-gage signals were under constant observation. At the appearance of signals indicative of vibration, the turbine speed was held constant and the signals were recorded. These oscillograph records were used in the computation of turbine speeds, vibration frequency, and vibratory stress. (Reference 4 describes details of the methods.)\n\nRESULTS AND DISCUSSION\n\nTwo oscillograph records representative of the data obtained during engine operation are given in figure 3. These records show strain-gage signals indicating turbine-blade vibration at nominal engine speeds of 10,000 and 11,500 rpm. In each of the records shown, the uppermost oscilloscope trace was produced by a 400-cycle-per-second signal from a vacuum-tube tuning fork. This trace served as a frequency standard for accurate determination of turbine speed and vibration frequency. The two oscilloscope traces below the frequency standard show the variation in the output of the two strain-gage bridges, one arm of each being a high-temperature strain gage. The remaining signal was produced by an impulse tachometer mounted on the turbojet engine. Two impulses occurred during each revolution of the turbine wheel.\n\nIn figure 3(a), the upper strain-gage trace shows the presence of vibration in a turbine blade; the other turbine blade was not vibrating and the signal shown in the lower strain-gage trace is therefore indicative only of the level of electrical interference caused by small variations in resistance between the slip rings and the brushes.", "timestamp": "2026-07-22T06:58:38.355404+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 33, "total_pages": 92, "image_filename": "19930085951_p33.jpg", "text": "```markdown\nNACA RM L9D29\n\nCONFIDENTIAL\n\nEfficiency, $\\eta$\n1.0\n.9\n.8\n.7\n.6\n.5\n.4\n.3\n.2\n.1\n0\n\nHelical-tip Mach number\nAir-stream Mach number\n\n$\\beta_{0.75R}$ 20° 25° 30° 35° 40° 45° 50°\n\nAdvance ratio, J\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0 3.2 3.4 3.6 3.8\n\nMach number, M\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\nNACA\n\n(c) Efficiency.\n\nFigure 9.— Concluded. Rotational speed, 1350 rpm.\nCONFIDENTIAL\n\n31\n```", "timestamp": "2026-07-22T06:58:43.053095+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 54, "total_pages": 92, "image_filename": "19930085870_p54.jpg", "text": "NACA RM No. L9D07\n55\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. { O $C_L$\n                 { □ $C_m$\nWedge L.E.     { △ $C_L$\n                 { ◇ $C_m$\n\n.16\n\n.08\n$C_L$\n0\n\n-.08\n\n-.16\n\n-.24\n\n.01\n$C_m$\n0\n-.01\n\n.06\nElliptical L.E. { O $C_D$\n                 { □ $L/D$\nWedge L.E.     { △ $C_D$\n                 { ◇ $L/D$\n\n.04\n$C_D$\n.02\n\n0\n-8   -6   -4   -2   0   2   4   6   8\n$\\alpha$, deg\n\n6\n4\n$L/D$\n2\n0\n\n[NACA logo]\n\n(c) Wing 3. w = 0.880; R = 990,000.\nFigure 7. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:58:43.754944+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 30, "total_pages": 34, "image_filename": "19930085913_p30.jpg", "text": "NACA RM L9F24\n29\n\n<!-- Image (168, 109, 835, 468) -->\n\n(c) Model B; $\\Lambda = 45^\\circ$; $\\sigma_W = -1$.\n\n<!-- Image (190, 535, 802, 846) -->\n\n(d) Model B; $\\Lambda = 45^\\circ$; $\\sigma_W = 0$.\n\nFigure 3.— Continued.", "timestamp": "2026-07-22T06:58:45.120225+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 21, "total_pages": 25, "image_filename": "19930085979_p21.jpg", "text": "20\n\nHot-gas Average model-air\nbleedback temperature rise\n(percent) ($^\\circ$F)\nO 2.90 28\n$\\square$ 4.59 45\n$\\diamondsuit$ 5.83 60\n$\\triangle$ 7.13 72\n\n<!-- Image (48, 100, 927, 802) -->\n\nFigure 7. - Temperature distributions measured on accessory housing for several values of hot-gas bleedback.\nPlenum-chamber-gas temperature, 1000$^\\circ$ F; tunnel-air velocity, 200 feet per second; tunnel total temperature,\n0$^\\circ$ F.\n\nNACA RM E9E12", "timestamp": "2026-07-22T06:58:47.431151+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 22, "total_pages": 30, "image_filename": "19930085970_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:58:49.301775+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 44, "total_pages": 52, "image_filename": "19930085911_p44.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 43\n\n1152\n\nNet acceleration, $a_n$, g's\n\nFree-stream Mach number, $M_0$\n\nTime after release, $\\tau$, sec\n\n(a) Resultant flight conditions.\n\nFigure 10. - Time history of flight data and performance of ram-jet unit 16-A-5.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:58:49.500098+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 13, "total_pages": 40, "image_filename": "19930085997_p13.jpg", "text": "NACA RM A9I29 CONFIDENTIAL 11\n\nthe models were slightly different. In configuration F, an expansion occurs at station 2.681 which probably changed the shock-wave pattern immediately ahead of the inlet.\n\nThe variation of $H_2/H_0$ with $m_1/m_0$ measured in tests of configurations E and F is shown in figures 15 and 16. Comparison of these results with similar curves of configuration D, shown in figures 7 and 12, indicates that, at an angle of attack of $0^\\circ$ and at a Mach number of 2.01, drooping the forebody of the model improved the range of mass-flow ratios over which $H_2/H_0$ was maintained at relatively high values. A similar improvement, but to a lesser degree, can be noted at angle of attack.\n\nCONCLUSIONS\n\nFrom tests at Mach numbers between 1.36 and 2.01 and Reynolds numbers between 2.6 and 3.4 million (based upon the length of the model ahead of the inlets) of several configurations of a duct-inlet model having side scoops and employing boundary-layer suction, the following conclusions are drawn:\n\n1. For the best configuration developed in the investigation, it was found that the arrangement advantages of side scoops can be utilized with total-pressure recovery within 0.05 of that of nose inlets neglecting the energy expended in boundary-layer removal at free-stream Mach numbers less than 1.8.\n\n2. The total-pressure distribution within the ducts was nonuniform, and the variation increased with free-stream Mach number.\n\n3. The energy expended in removing the boundary layer at full-scale flight conditions was estimated to be equivalent to a reduction of approximately 0.04 of the measured total-pressure recovery.\n\n4. The total-pressure recovery decreased with increasing positive angles of attack. Drooping the forebody of the model with respect to the inlets improved the pressure recovery at angle of attack, and the range of mass-flow ratios in which high pressure recovery could be maintained was increased.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:58:50.777136+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 17, "total_pages": 132, "image_filename": "19930085990_p17.jpg", "text": "NACA RM A9I01 CONFIDENTIAL 15\n\nacceleration factor of unity at Mach numbers above about 0.87. Below this Mach number, the variation of stabilizer angle with speed was stable and a total change of stabilizer angle of $1.7^\\circ$ would be necessary to balance the airplane in level flight between Mach numbers of 0.50 and 0.87.\n\nWith the tail mounted above the extended wing-chord plane, the airplane would possess static longitudinal stability at all Mach numbers but the variation of stabilizer angle with velocity would be unstable at Mach numbers above about 0.90. A change of $2.4^\\circ$ in the stabilizer angle would be required to balance the airplane in level flight between Mach numbers of 0.50 and 0.95.\n\nTo compare the longitudinal control afforded by the all-movable stabilizer with that which could be accomplished with a fixed stabilizer and an elevator, elevator-effectiveness data from reference 6 were applied to the hypothetical airplane. The tail model of reference 6 was equipped with a 20-percent area, constant-chord elevator and the plan form and profile were identical with those of the horizontal tail investigated herein. The elevator-effectiveness data of reference 6 are reproduced herein in figure 45 and in application of the data it was assumed that there was no effect of scale between Reynolds numbers of 2,000,000 and 1,000,000 and that the elevator efficiency factor was 100 percent.\n\nThe variation with Mach number of the elevator deflection required to balance the airplane at the previously assumed flight conditions is presented in figure 46.\n\nThe calculated static longitudinal stability and control of the airplane with a fixed stabilizer and an elevator are similar to those previously discussed for the airplane with the all-movable stabilizer. About 50-percent greater deflection would be required of the elevator to produce the same balance lift coefficient as the all-movable stabilizer.\n\nLongitudinal Characteristics with the Flaps Deflected\n\nThe variation with lift coefficient of the stabilizer angle required to balance the model with the flaps deflected is presented in figure 47 for the model with the horizontal tail in the extended wing-chord plane. The corresponding drag coefficient is shown in the same figure and the lift-drag ratio as a function of lift coefficient for balance is shown in figure 48.\n\nThese experimental results have been used to predict the power-off gliding speed and sinking speed at sea level of a hypothetical airplane with a wing loading of 100 pounds per square foot. The effects of the proximity of the ground and the increased drag due to landing gear have\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:58:52.946195+00:00"}