PDF/ocr_dataset_1078.jsonl

{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 29, "total_pages": 52, "image_filename": "19930085911_p29.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:50:47.473153+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 19, "total_pages": 34, "image_filename": "19930085913_p19.jpg", "text": "```markdown\n18\n\nTABLE I.- EXPERIMENTAL DATA - Concluded\n\n\\begin{tabular}{|l|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|c|}\n\\hline\n\\multirow{4}{*}{Model} & \\multirow{4}{*}{Run} & \\multirow{4}{*}{\\shortstack{q$_r$ \\\\ (lb/sq ft)}} & \\multirow{4}{*}{\\shortstack{v$_r$ \\\\ (fps)}} & \\multirow{4}{*}{\\shortstack{Mach \\\\ number}} & \\multirow{4}{*}{\\shortstack{Distance \\\\ of weight \\\\ from root \\\\ (percent l)}} & \\multicolumn{4}{c|}{\\shortstack{Frequencies \\\\ (cps)}} & \\multicolumn{12}{c|}{\\shortstack{Phase-angle relationship of bending and torsional stresses. \\\\ (Ref indicates reference strain-gage trace)}} \\\\\n\\cline{7-22}\n& & & & & & \\multicolumn{3}{c|}{Natural} & \\multirow{3}{*}{Flutter} & \\multicolumn{4}{c|}{2nd natural mode} & \\multicolumn{4}{c|}{3rd natural mode} & \\multicolumn{4}{c|}{Flutter mode} \\\\\n\\cline{7-9}\\cline{11-22}\n& & & & & & 1st & 2nd & 3rd & & 1 & 2 & 3 & 4 & 1 & 2 & 3 & 4 & 1 & 2 & 3 & 4 \\\\\n\\cline{11-22}\n& & & & & & & & & & (deg) & (deg) & (deg) & (deg) & (deg) & (deg) & (deg) & (deg) & (deg) & (deg) & (deg) & (deg) \\\\\n\\hline\n\\multirow{14}{*}{\\shortstack[l]{Model C \\\\ Swept untapered wing; \\\\ $\\Lambda = 60^\\circ$ \\\\ Weight moved along \\\\ midchord line; $\\alpha_w = 0$ \\\\ Reynolds number $\\cong 7514.6r_r$}} & 82 & 40.92 & 192.7 & 0.1670 & 0 & 2.94 & 17.52 & 22.35 & 13.33 & Ref & 180 & - & 0 & Ref & 0 & 0 & 180 & Ref & 166 & 0 & 9 \\\\\n\\cline{2-22}\n& 83 & 39.57 & 190.0 & .1642 & 13.88 & 2.96 & 17.07 & 22.25 & 13.64 & Ref & 180 & - & 0 & Ref & 0 & 0 & 180 & Ref & 170 & 0 & 28 \\\\\n\\cline{2-22}\n& 84 & 36.73 & 182.7 & .1581 & 25.00 & 2.93 & 14.45 & 22.10 & 12.00 & Ref & 180 & - & 0 & Ref & 0 & 0 & 180 & Ref & 151 & 21 & 8 \\\\\n\\cline{2-22}\n& 85 & 41.23 & 194.3 & .1678 & 30.55 & 2.90 & 12.41 & 21.80 & 9.82 & Ref & 180 & - & 0 & Ref & 0 & 0 & 180 & Ref & 90 & 14 & 34 \\\\\n\\cline{2-22}\n& 86 & 41.24 & 193.9 & .1678 & 36.11 & 2.85 & 11.07 & 21.20 & 10.00 & Ref & 180 & - & 0 & Ref & 0 & 0 & 180 & Ref & 90 & 13 & 27 \\\\\n\\cline{2-22}\n& 87 & 57.89 & 230.3 & .1995 & 41.67 & 2.73 & 10.34 & 20.35 & 9.52 & Ref & 180 & - & 0 & Ref & 0 & 0 & 180 & Ref & 13 & 0 & 340 \\\\\n\\cline{2-22}\n& 88 & 74.18 & 260.4 & .2268 & 47.22 & 2.60 & 10.21 & 19.77 & 15.63 & Ref & 180 & - & 0 & Ref & 0 & 0 & 180 & Ref & 0 & 0 & --- \\\\\n\\cline{2-22}\n& 89 & 73.12 & 258.3 & .2250 & 52.77 & 2.42 & 10.59 & 19.06 & 15.58 & Ref & 180 & - & 0 & Ref & 0 & 0 & --- & Ref & 0 & 0 & 0 \\\\\n\\cline{2-22}\n& 90 & 77.20 & 265.1 & .2310 & 58.33 & 2.24 & 11.42 & 18.25 & 15.28 & Ref & 180 & - & 0 & Ref & 0 & 0 & 180 & Ref & 0 & 0 & --- \\\\\n\\cline{2-22}\n& 91 & 95.92 & 296.1 & .2588 & 63.89 & 2.08 & 12.63 & 17.89 & 10.00 & Ref & 180 & - & 0 & Ref & 0 & 0 & 180 & Ref & 16 & 0 & --- \\\\\n\\cline{2-22}\n& 92 & 86.99 & 281.4 & .2460 & 69.44 & 1.90 & 14.21 & 17.80 & 10.00 & Ref & 180 & 0 & 0 & Ref & 0 & 0 & 180 & Ref & 20 & 0 & --- \\\\\n\\cline{2-22}\n& 93 & 75.97 & 262.8 & .2295 & 75.00 & 1.73 & 15.31 & 18.28 & 10.91 & Ref & 180 & 0 & 0 & Ref & 0 & 0 & 180 & Ref & 180 & 0 & --- \\\\\n\\cline{2-22}\n& 94 & 71.34 & 254.2 & .2220 & 80.56 & 1.60 & 15.20 & 18.90 & $^a$7.89 & Ref & 180 & 0 & 0 & Ref & 0 & 0 & 180 & Ref & 188 & 0 & --- \\\\\n\\cline{2-22}\n& -- & \\multicolumn{3}{c|}{No flutter test made} & 91.67 & 1.36 & 14.15 & 17.35 & No record & Ref & 180 & 0 & 0 & Ref & 0 & 0 & 180 & \\multicolumn{4}{c|}{No record} \\\\\n\\cline{2-22}\n& 95 & 53.16 & 218.8 & .1908 & 97.22 & 1.28 & 12.94 & 16.72 & 5.51 & Ref & 180 & 0 & 0 & Ref & 0 & 0 & 180 & Ref & 180 & 0 & 19 \\\\\n\\hline\n\\end{tabular}\n\n[Figure: Diagram of wing model with dimensions l, t, and 2b]\n\n$^a$Note oscillograph record, figure 1.\n\nNACA\nNACA RM L9F24\n```", "timestamp": "2026-07-22T06:50:48.331682+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 22, "total_pages": 22, "image_filename": "19930085928_p22.jpg", "text": "```markdown\nNACA RM No. A9A31\n\nCONFIDENTIAL\n\n<!-- Image (49, 109, 934, 832) -->\n\nFigure 7. -Variation of total-pressure ratio with mass-flow ratio for models with 0.085-inch by 0.300-inch slots.\n\n21\n```", "timestamp": "2026-07-22T06:50:52.156338+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 21, "total_pages": 47, "image_filename": "19930085919_p21.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:50:56.144147+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 2, "total_pages": 24, "image_filename": "19930085991_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:50:56.346292+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 42, "total_pages": 92, "image_filename": "19930085870_p42.jpg", "text": "NACA RM No. L9D07\n43\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O $C_L$\n                {□ $C_m$\n.16\nWedge L.E. {△ $C_L$\n           {◇ $C_m$\n\n.08\n$C_L$\n0\n-.08\n-.16\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\n0\n\n6\n$L/D$\n4\n2\n0\n\n-8 -6 -4 -2 0 2 4 6 8\n$\\alpha$, deg\n\n(b) Wing 2. $w=0.530$; $R=1,250,000$.\nFigure 6. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:56.346889+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 10, "total_pages": 55, "image_filename": "19930085966_p10.jpg", "text": "NACA RM L9B17 CONFIDENTIAL 9\n\nTotal temperature at station 5\n\n$$\nT_{t5} = \\frac{2\\gamma}{gR(\\gamma - 1)} \\frac{p_5^2}{\\left( \\frac{\\dot{W}_5}{A_5} \\right)^2} \\left[ \\left( \\frac{p_{t5}}{p_5} \\right)^{\\frac{\\gamma-1}{\\gamma}} - 1 \\right] \\left( \\frac{p_{t5}}{p_5} \\right)^{\\frac{\\gamma-1}{\\gamma}}\n$$\n\nMach number at station 0\n\n$$\nM_0 = \\sqrt{ \\frac{2}{\\gamma - 1} \\left[ \\left( \\frac{p_{t0}}{p_0} \\right)^{\\frac{\\gamma-1}{\\gamma}} - 1 \\right] }\n$$\n\nRESULTS AND DISCUSSION\n\nRam-jet performance can be expressed in many different terms according to what purpose is being accomplished. The thrust force is a significant quantity that can be directly compared with the drag force of the body to determine the resulting equilibrium level-flight speed or the possible acceleration at a given flight speed and rate of climb. Thrust is often more usefully thought of in terms of a dimensionless coefficient $C_F$, which is comparable to the drag coefficient of a body. Neither of these quantities reflects fuel economy or efficiency of energy conversion. Over-all efficiency is the product of burner, thermodynamic-cycle, and propulsive efficiencies and expresses the percentage of energy in the fuel converted to thrust energy. The reciprocal of the over-all efficiency is proportional to the specific fuel consumption and indicates the fuel rate required per unit thrust horsepower.\n\nAn analysis of the relation of ram-jet performance parameters to flight Mach number is given in reference 1. Performance curves similar to those of reference 1 have been prepared from the data taken on the thin-plate-burner ram-jet configuration. A plot of reduced thrust (thrust/$\\delta_0$) against simulated flight Mach number is presented in figure 11. Thrust was calculated using the following equation:\n\n$$\nF = \\frac{\\dot{W}_3}{g} (V_7 - V_0)\n$$\n\nIt is estimated that the thrust values shown in figure 11 are accurate to ±10 percent. An analysis of figure 11 shows that for a given temperature ratio the reduced thrust is proportional to the simulated flight Mach number to a power which varies between 1.8 and 2.1. The highest thrust attained in the tests was 410 pounds at a Mach number of 0.54 and a temperature ratio $T_t$ of 3.0.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:56.804805+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 6, "total_pages": 25, "image_filename": "19930085979_p6.jpg", "text": "```markdown\nNACA RM E9E12\n\n$P_f$ total pressure at front rakes, pounds per square foot absolute\n\n$q_0$ free-stream dynamic pressure, pounds per square foot\n\nA decrease in ram-pressure recovery of approximately 1 percent was observed for an increase in angle of attack from $0^\\circ$ to $8^\\circ$.\n\nLip-pressure distribution. - The effect of angle of attack on inlet-lip pressure distribution is shown in figure 3. The pressure distribution is presented in terms of a pressure coefficient\n\n$$S = 1 - \\frac{p-p_0}{q_0}$$\n\nwhere $p$ is the local surface static pressure in pounds per square foot absolute and $p_0$ is the free-stream static pressure in pounds per square foot absolute.\n\nAerodynamic Investigation with Cold-Gas Bleedback\n\nMass-flow characteristics. - No measurable change in mass flow through the model was observed with increasing cold-gas bleedback.\n\nRam-pressure recovery. - In order to determine the effect of the jets alone, cold gas was bled into the inlet-air stream. The effect of bleedback on ram-pressure recovery is shown in figure 4 for tunnel-air velocities of 200, 280, 355, and 435 feet per second. The loss in ram-pressure recovery is linearly related to bleedback and no effect of velocity on ram-pressure recovery is evident.\n\nLip-pressure distribution. - A slight movement of the stagnation point to a position further inside the lip was observed with cold-gas bleedback (fig. 5). This movement increased with increasing bleedback and was caused by the decrease in the inlet-velocity ratio with increasing bleedback, as evidenced by the reduced pressure coefficients in the inlet. The slight change in lip-pressure distribution with bleedback should have a small effect on the external nacelle drag.\n\nNo measurable change in mass flow through the model occurred as a result of cold-gas bleedback. A decreasing inlet-velocity ratio must therefore occur with increasing bleedback because an increasing part of the total flow through the model is represented by the bleedback gas and as a consequence the air flow entering the inlet is reduced.\n\n```", "timestamp": "2026-07-22T06:50:59.440258+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 19, "total_pages": 92, "image_filename": "19930085951_p19.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n\nNACA RM L59D29\n\n[Figure: A large propeller dynamometer in a test section with tunnel open. A person is visible in the background observing the setup.]\n\nNACA\nL-45763\n\nFigure 1.- Propeller dynamometer in test section with tunnel open.\n\nCONFIDENTIAL\n\n17", "timestamp": "2026-07-22T06:51:03.830539+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 10, "total_pages": 30, "image_filename": "19930085970_p10.jpg", "text": "8 CONFIDENTIAL NACA RM A9E09\n\nfrom the discrepancy between the respective drag-rise factors noted previously.\n\nThe principal conclusion to be drawn from these results is that, for the type of plan form investigated, maximum lift-drag ratios measured at low Reynolds numbers cannot be assumed to be reliable for prediction of performance at the full-scale Reynolds numbers. It may also be concluded for the present case that the failure to realize theoretical maximum lift-drag ratios is due, in a large part, to the corresponding failure to realize theoretical drag-rise factors.\n\nPitching Moment\n\nCurves of pitching-moment coefficient as a function of lift coefficient for each test Mach number are shown in figure 11. The results of reference 3 for the higher Reynolds number and the results of reference 2 are also plotted in the figure. The large differences between the present results and those of reference 3 that are apparent in figure 11 are believed to be due to the dissimilarity of the boundary-layer conditions existing at the widely different test Reynolds numbers. The correspondence of the present result and that of reference 2, at a Mach number of about 1.5 and at similar Reynolds numbers, is good.\n\nThe locations of the aerodynamic center of the model at the various Mach numbers have been determined from the slopes of the pitching-moment curves between 0 and 0.2 lift coefficient and are shown in figure 12.\n\nCorresponding values from reference 3 are also shown as are values calculated by the methods of references 6 and 7 wherever applicable.² At subsonic Mach numbers, the present variation of location of the aerodynamic center with Mach number is similar to that of reference 3, although the rate of rearward movement at the higher Mach numbers exhibited by the present results is probably too large. Although the present results are not believed to offer a quantitative representation of aerodynamic-center locations at high Reynolds numbers, they roughly verify the magnitude of the total theoretical rearward shift of that parameter over the range of the test Mach numbers.\n\n---\n\n²The method of reference 7 cannot be used at Mach numbers below that at which the Mach cones emanating from the trailing-edge apex crosses the leading edges (1.43 Mach number).\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:05.266518+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 88, "total_pages": 114, "image_filename": "19930086061_p88.jpg", "text": "84\nNACA RM L9J07\n\n$$c_l$$\n16\n12\n8\n4\n0\n$$\\alpha = 4.1^\\circ$$\n$$C_L = 0.13$$\n\n$$\\alpha = 24.1^\\circ$$\n$$C_L = 0.83$$\n\n$$c_l$$\n16\n12\n8\n4\n0\n$$\\alpha = 8.1^\\circ$$\n$$C_L = 0.28$$\n\n$$\\alpha = 39.1^\\circ$$\n$$C_L = 1.17$$\n\n$$c_l$$\n8\n4\n0\n100 80 60 40 20 0\n$$\\frac{y}{b/2}$$ , percent\n$$\\alpha = 14.1^\\circ$$\n$$C_L = 0.50$$\n\n100 80 60 40 20 0\n$$\\alpha = 50.1^\\circ$$\n$$C_L = 0.63$$\n[Figure: NACA logo]\n\nFigure 33.- Effect of $$\\alpha$$ on the spanwise $$c_l$$ variation of wing 3;\n$$\\psi = 0^\\circ$$.", "timestamp": "2026-07-22T06:51:06.217846+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 57, "total_pages": 72, "image_filename": "19930085491_p57.jpg", "text": "```markdown\n56\n\nCONFIDENTIAL\n\n(a)\nWF-57\nR=0.62 million\n\n(b)\nWF-60\nR=0.62 million\n\n(c)\nWF-63\nR=0.62 million\n\n(d)\nWF-63\nR=0.84 million\n\n(e)\nWF-67\nR=0.95 million\n\n(f)\nWF-70\nR=0.62 million\n\n$\\circ \\nabla \\diamond C_m$ (Experiment)\n— — $C_m$ (Approximate linear theory)\n- - - - Center-of-pressure location (Experiment)\n\nPitching-moment coefficient, $C_m, \\frac{1}{4}$\nDistance from moment axis to center of pressure, percent m.a.c.\nLift coefficient, $C_L$\n\n[Figure: Six graphs showing pitching-moment coefficient and center-of-pressure location variations with lift coefficient for different wing-fuselage configurations (WF-57, WF-60, WF-63, WF-67, WF-70) at various Reynolds numbers (R).]\n\nFigure 8—Pitching-moment coefficient and center-of-pressure location variations with lift coefficient of swept-back wing and fuselage configurations. Moment reference axis at 50 percent of mean aerodynamic chord.\n\nNACA\n\nNACA RM No. A53J04\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:51:13.075050+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 5, "total_pages": 132, "image_filename": "19930085990_p5.jpg", "text": "NACA RM A9I01\nCONFIDENTIAL\n3\n\nThe effective downwash angle at the tail, the Mach number at the tail,\nand the tail efficiency factor are presented herein.\n\nCOEFFICIENTS AND SYMBOLS\n\nThe following coefficients are used in this report:\n\n$C_L$ lift coefficient $\\left(\\frac{\\text{lift}}{qS}\\right)$\n\n$C_D$ drag coefficient $\\left(\\frac{\\text{drag}}{qS}\\right)$\n\n$C_m$ pitching-moment coefficient about an axis normal to the plane of\nsymmetry passing through the quarter point of the wing mean\naerodynamic chord $\\left(\\frac{\\text{pitching moment}}{qSc'}\\right)$\n\n$\\frac{\\Delta H}{q}$ total-pressure-loss coefficient $\\left(\\frac{H_0-H}{q}\\right)$\n\nThe following symbols are used in this report:\n\na speed of sound, feet per second\n\nb twice the span of the semispan wing, feet\n\nc local wing chord, feet\n\nc' wing mean aerodynamic chord, chord through centroid of the\nwing semispan plan form $\\left(\\frac{\\int_0^{b/2} c^2 dy}{\\int_0^{b/2} c dy}\\right)$\n\nH local stagnation pressure in the region of the horizontal tail,\npounds per square foot.\n\n$H_0$ free-stream stagnation pressure, pounds per square foot\n\n$i_t$ angle of the stabilizer setting with respect to the wing-chord\nplane, degrees\n\n$l_t$ tail length, distance from quarter point of the wing mean\naerodynamic chord to the quarter point of the horizontal-\ntail mean aerodynamic chord, feet\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:16.027935+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 52, "total_pages": 149, "image_filename": "19930083192_p52.jpg", "text": "```markdown\n48\nNACA TN 1976\n\nthe complete airplane was then divided so that the upper wing of the equivalent biplane would have the same static deflection as the original wing tip.\n\nIn addition to and as a check on the method of calculation, tests were made in the Langley gust tunnel with a model having two wings of different frequencies. The wing deflection and fuselage acceleration were the primary quantities measured. The model used is shown diagrammatically in figure 42 and its characteristics together with the test conditions are given in table XVI.\n\nFor each of the two wing frequencies tests were made at one forward speed and three gust sizes. Time histories of pitch, acceleration increment of the fuselage, and wing deflection were obtained. Some results of the tests are shown in figure 43, in which the maximum wing-tip deflection per unit acceleration increment is plotted for one wing as a function of gradient distance.\n\nThe analytical procedure described in reference 38 was utilized to compute the response of the system when damping of the motion is included and when aerodynamic damping is neglected. The results have been included in figure 43 for comparison with experiment.\n\n### Analytical Study\n\nIn order to determine the effect of the various parameters on dynamic response, a series of calculations was made at three gust-gradient distances for four airplanes. The characteristics of the airplanes, which were designated as A, B, C, and D, are given in table XVII. The parameters studied were the wing stiffness, airplane weight, and forward speed, with all damping neglected. The gust velocity was not varied since ratios of dynamic to static conditions were utilized and it was not a factor for linear equations. The conditions used in the calculations are given in table XVII.\n\nThe calculated results for each airplane of the maximum deflection ratio, wing-tip acceleration ratio, and fuselage acceleration ratio are shown in figure 44.\n\n### Discussion\n\nInspection of the results of figure 43 indicates, in general, that dynamic response is overestimated when damping is neglected; whereas the inclusion of damping leads to fair agreement with experimental results. The results obtained were not sufficiently accurate to provide an absolute\n```", "timestamp": "2026-07-22T06:51:19.012609+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 30, "total_pages": 52, "image_filename": "19930085911_p30.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 29\n\n1152\n\n[Figure: Three-quarter rear view of a flame holder assembly with multiple cylindrical tubes and star-shaped vanes arranged in a circular pattern. A scale bar labeled \"INCHES\" is visible at the bottom right. NACA logo, C-21694, 6-9-48.]\n\n(a) Three-quarter rear view.\n\n[Figure: Rear view of the same flame holder assembly, showing the circular arrangement of seven star-shaped vanes around a central vane. A scale bar labeled \"INCHES\" is visible at the bottom right. NACA logo, C-21695, 6-9-48.]\n\n(b) Rear view.\n\nFigure 6. - Flame holder for supersonic ram-jet unit 16-A-5.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:20.335978+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 8, "total_pages": 46, "image_filename": "19930085983_p8.jpg", "text": "6\nCONFIDENTIAL\nNACA RM A9I27\n\nConstriction\n\nThe constriction effects of the tunnel walls have been evaluated by\nthe method of reference 9. No modification of this method has been made\nto account for the effects of sweepback. The magnitude of the correc-\ntions applied to the Mach number and to the dynamic pressure is illus-\ntrated by the following table:\n\n| Corrected Mach number | Uncorrected Mach number | q, corrected / q, uncorrected |\n| :--- | :--- | :--- |\n| 0.930 | 0.919 | 1.012 |\n| .890 | .884 | 1.007 |\n| .800 | .798 | 1.003 |\n| .600 | .599 | 1.002 |\n| .200 | .200 | 1.001 |\n\nBase Pressure\n\nThe pressure on the base of the model fuselage was measured and,\nin an effort to correct for support interference, the drag data were\ncorrected to correspond to a base pressure equal to the static pressure\nof the free stream. The base-pressure correction to the drag was less\nthan 5 percent for Mach numbers up to 0.75, and increased to approxi-\nmately 20 percent at a Mach number of 0.93. The base-pressure correc-\ntion reduced the drag.\n\nTares\n\nThere were no tares due to direct air forces on the model-support\nequipment, since the balance was within the model. Corrections were\nmade for the change in static tares due to angle of attack.\n\nRESULTS AND DISCUSSION\n\nLongitudinal Characteristics\n\nElevon effectiveness and hinge moments.- Angle of attack, drag\ncoefficient, and pitching-moment coefficient as functions of lift coef-\nficient, and hinge-moment coefficient as a function of angle of attack\nare presented in figures 4 to 8, inclusive, for various elevon deflec-\ntions for Mach numbers ranging from 0.20 to 0.93. The angle of attack\nfor zero lift became more positive as the elevon was deflected upward\nand the minimum drag coefficient was increased considerably by negative\nelevon deflections greater than -5°.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:22.926490+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 22, "total_pages": 47, "image_filename": "19930085919_p22.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 21\n\n[Figure: (a) Long fuselage. NACA A-11788]\n\n[Figure: (b) Short fuselage and split flap. NACA A-11892]\n\nFigure 3.- The wing-fuselage combinations mounted in one of the Ames 7- by 10-foot wind tunnels.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:25.914029+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 20, "total_pages": 92, "image_filename": "19930085951_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:51:26.943191+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 1, "total_pages": 40, "image_filename": "19930085997_p1.jpg", "text": "NACA RM A9I29\n\nCONFIDENTIAL\n\nCopy 310\nRM A9I29\n\nNACA\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION AT SUPERSONIC SPEEDS OF\nSIDE SCOOPS EMPLOYING BOUNDARY-LAYER SUCTION\n\nBy Sherman S. Edwards\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nDecember 12, 1949\n\nCONFIDENTIAL\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: NACA RESEARCH ABSTRACT NO. 120\nEFFECTIVE DATE: SEPTEMBER 13, 1957\nWHL", "timestamp": "2026-07-22T06:51:28.544302+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 43, "total_pages": 92, "image_filename": "19930085870_p43.jpg", "text": "44\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. $\\circledcirc$ $C_L$ $\\square$ $C_m$\nWedge L.E. $\\triangle$ $C_L$ $\\diamond$ $C_m$\n.16\n.08\n$C_L$\n0\n-.08\n-.16\n-.24\n.01\n$C_m$\n0\n-.01\n\n.06\nElliptical L.E. $\\circledcirc$ $C_D$ $\\square$ $L/D$\nWedge L.E. $\\triangle$ $C_D$ $\\diamond$ $L/D$\n.04\n$C_D$\n.02\n0\n-8\n-6\n-4\n-2\n0\n2\n4\n6\n8\n$\\alpha$, deg\n6\n4\n$L/D$\n2\n0\nNACA\n\n(c) Wing 3. $w = 0.661$; $R = 1,230,000$.\nFigure 6. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:28.868255+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 89, "total_pages": 114, "image_filename": "19930086061_p89.jpg", "text": "```markdown\nNACA RM L59J07\n\n| Station | $\\frac{y}{b/2}$ |\n| :--- | :--- |\n| 1 | 0 |\n| 2 | 0.167 |\n| 3 | 0.333 |\n| 4 | 0.500 |\n| 5 | 0.667 |\n| 6 | 0.833 |\n| 7 | 0.916 |\n\n[Graph plotting $c_l$ vs $\\alpha$]\n\n$c_l$\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\n$\\alpha$, deg\n0\n4\n8\n12\n16\n20\n24\n28\n32\n36\n40\n44\n48\n52\n\n[NACA logo]\n\nFigure 34.- Variation of $c_l$ with $\\alpha$ at seven stations along the semispan of wing 1; $\\psi = 0^\\circ$.\n\n85\n```", "timestamp": "2026-07-22T06:51:32.456059+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 58, "total_pages": 72, "image_filename": "19930085491_p58.jpg", "text": "--- - Approximate linear theory\n\nCONFIDENTIAL\n\n(a)\nWF-57\nR=0.62 million\n\n(b)\nWF-60\nR=0.62 million\n\n(c)\nWF-63\nR=0.62 million\n\nPitching-moment coefficient, $C_{m_{\\frac{1}{4}}}$\n\n(d)\nWF-63\nR=0.84 million\n\n(e)\nWF-67\nR=0.95 million\n\n(f)\nWF-70\nR=0.62 million\n\nLift coefficient, $C_L$\n\nNACA\n\nLift coefficient, $C_L$\n\nFigure 9.- Pitching-moment coefficient variations with lift coefficient of swept-back wing and fuselage configurations. Moment reference axis at 25 percent of mean aerodynamic chord.\n\nNACA RM No. A8J04\n\nCONFIDENTIAL\n\n57", "timestamp": "2026-07-22T06:51:33.932829+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 11, "total_pages": 55, "image_filename": "19930085966_p11.jpg", "text": "10 CONFIDENTIAL NACA RM L9B17\n\nThe relation of thrust coefficient $C_F$ to simulated flight Mach number is presented in figure 12. Thrust coefficient was defined as follows:\n\n$$\nC_F = \\frac{F}{q_0 A_3}\n\\tag{1}\n$$\n\nThe area $A_3$ was used as the reference area because the cooling shroud would not necessarily be used in a flight model. Thrust coefficients from 0.370 to 0.397 were attained at a temperature ratio of 5.0. Figure 12 indicates that within the accuracy of the data the ram-jet unit produced no variation of $C_F$ in the Mach number range covered by the tests for constant temperature ratio $T_t$.\n\nThe range of simulated flight Mach number obtained in the test was limited and the relation of the low Mach number data to possible high Mach number performance was not obvious; therefore, an estimate based on the subsonic test-stand data was made of the thrust-coefficient variation with flight Mach number. The combustion-chamber performance in terms of Mach numbers and pressure and temperature ratios was held to those values obtained in the tests regardless of the flight Mach number. It was believed reasonable to restrict the combustion-chamber inlet velocity by limiting the inlet Mach number to test values since most ram-jet burners depreciate in performance if the air velocity is increased beyond certain values. This limitation would be imposed physically by regulation of the nozzle exit area. The limitation of the combustion-chamber temperature ratios to test values is considered conservative since the higher levels of pressures and temperatures associated with higher flight Mach numbers are favorable to combustion. Further discussion of the assumptions and methods used in the calculations is given in the appendix.\n\nTwo cases were calculated, one for a temperature-rise ratio of 3.85 and a combustion-chamber inlet Mach number of 0.065 (fig. 13) and one for a temperature-rise ratio of 2.89 and an inlet Mach number of 0.085 (fig. 14). For inlet-total-pressure-recovery ratios of 80 and 90 percent both figures 13 and 14 show a continuously rising thrust coefficient with flight Mach number; however, for 100-percent inlet-total-pressure-recovery ratio both cases show a peak thrust coefficient in the region of a flight Mach number of 0.25 to 0.40 and a minimum thrust coefficient in the region of a flight Mach number of 1.0 to 1.4.\n\nIn discussing the calculated curves comparisons will be drawn with the test-data curves of figure 12, which differ significantly from the calculated curves in that the test-data curves are for a variable instead of a constant combustion-chamber inlet Mach number and a constant instead\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:34.381280+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 7, "total_pages": 25, "image_filename": "19930085979_p7.jpg", "text": "6\nNACA RM E52L2\n\nAerodynamic Investigation with Hot-Gas Bleedback\n\nOptimum orifice configuration. - The results presented herein are for the orifice configuration that gave the most uniform temperature distribution at the calculated value of bleedback (4.9 percent) and plenum-chamber-gas temperature (1000° F) necessary for adequate ice prevention corresponding to an icing condition of 1.4 grams per cubic meter at a tunnel total temperature of 0° F. Preliminary computations utilizing the equations and the method of reference 2 showed the impracticality of providing ice protection for the accessory housing. Because of the large inlet diameter and the short mixing distance (11 in.) from the orifices to the accessory-housing tip, extremely large-diameter orifices would be required to obtain adequate jet penetration. Inasmuch as these large-diameter orifices would necessarily be few in number in order to pass only the required amount of bleedback gas, the temperature distribution inside the model would be very poor.\n\nThe investigation was therefore undertaken with the purpose of providing ice protection only for the inlet screen under the most severe conditions used in the icing investigation. By use of the method of reference 2 and assuming that the depth of penetration of the jets required at the screen was one-half of the duct diameter at the accessory-housing tip, an orifice diameter of 0.52 inch was obtained at the design conditions for the model. The model was designed to handle 31 pounds of air per second at a tunnel-air velocity of 425 feet per second and a free-stream density of 0.0024 slug per cubic foot. The calculated total jet area required to pass the required amount of bleedback gas (1.38 lb/sec at maximum air flow through the model) was 2.65 square inches. The results of reference 2 indicate that for a circular inlet, a minimum of 12 orifice holes are required to obtain a good temperature distribution inside the model. In order to obtain the required total jet area with the minimum number of orifices, a configuration consisting of 12 orifices of 17/32-inch diameter equally spaced around the inlet was used in the investigation. This configuration gave the most uniform temperature distribution inside the model for the calculated values of bleedback and plenum-chamber-gas temperatures required for ice prevention.\n\nIn order to be certain that the orifice configuration found by the method of reference 2 was the best, several other configurations were experimentally investigated. A configuration consisting of nine equally spaced 5/8-inch-diameter orifices was investigated. The results showed that, although greater penetrations were obtained than with the 12-orifice configuration, the local-air-temperature", "timestamp": "2026-07-22T06:51:34.677464+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 11, "total_pages": 30, "image_filename": "19930085970_p11.jpg", "text": "NACA RM A9E09 CONFIDENTIAL 9\n\nCONCLUSIONS\n\nFrom the results of tests performed on a wing with the leading edge swept back 63° and of symmetrical section in combination with a body at Mach numbers from 0.5 to 0.95 and from 1.09 to 1.51, it is concluded that for highly swept thin wings of moderate aspect ratio:\n\n1. The variations of lift and drag coefficients with Mach number are continuous and small.\n\n2. The total rearward shift of the location of the aerodynamic center occurring between Mach numbers of 0.5 and 1.5 corresponds approximately to that predicted by the use of theoretical methods.\n\n3. The lift characteristics can be estimated with reasonable accuracy by analytical methods for Mach numbers as high as 1.5.\n\n4. The trend with Mach number of the values of minimum drag coefficient is similar to that indicated by theoretical methods.\n\n5. Measurements of drag due to lift and pitching moment at subsonic Mach numbers and low Reynolds numbers cannot be considered quantitatively representative of the corresponding characteristics at much higher Reynolds numbers.\n\nAmes Aeronautical Laboratory,  \nNational Advisory Committee for Aeronautics,  \nMoffett Field, Calif.\n\nREFERENCES\n\n1. Jones, R. T.: Estimated Lift-Drag Ratios at Supersonic Speeds. NACA TN 1350, 1947.\n\n2. Madden, Robert T.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back 63°.— Characteristics at a Mach Number of 1.53 Including Effect of Small Variations of Sweep. NACA RM A8J04, 1949.\n\n3. Reynolds, Robert M., and Smith, Donald W.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back 63°.— Subsonic Mach and Reynolds Number Effects on the Characteristics of the Wing and on the Effectiveness of an Elevon. NACA RM A8D20, 1948.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:35.383183+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 31, "total_pages": 52, "image_filename": "19930085911_p31.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:51:36.636553+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 6, "total_pages": 132, "image_filename": "19930085990_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM A9I01\n\nM Mach number (V/a)\n\n$M_t$ Mach number at the position corresponding to the centroid of the semitail area\n\nn normal-acceleration factor of the airplane\n\nq free-stream dynamic pressure ($\\frac{1}{2}\\rho V^2$), pounds per square foot\n\n$q_t$ dynamic pressure at the position corresponding to the centroid of the semitail area, pounds per square foot\n\nR Reynolds number $\\left(\\frac{\\rho V c'}{\\mu}\\right)$\n\nS area of the semispan wing, square feet\n\n$S_t$ area of the horizontal semitail, square feet\n\nu local airspeed in the tunnel-floor boundary layer, feet per second\n\nV airspeed, feet per second\n\ny distance from the plane of symmetry, feet\n\n$\\alpha_t$ effective angle of attack of the horizontal tail, degrees\n\n$\\alpha$ angle of attack of the wing-chord plane, degrees\n\n$\\delta$ tunnel-wall boundary-layer thickness, inches\n\n$\\delta^*$ displacement thickness of the boundary layer $\\left[\\int_1^\\delta (1-u/V)dy\\right]$, inches\n\n$\\delta_e$ elevator deflection, measured in a plane perpendicular to the elevator hinge axis, positive downward, degrees\n\n$\\delta_f$ trailing-edge flap deflection, measured in a plane perpendicular to the flap hinge axis, positive downward, degrees\n\n$\\delta_n$ leading-edge flap deflection, measured in a plane perpendicular to the flap hinge axis, positive downward, degrees\n\n$\\epsilon$ effective average angle of downwash, positive when the air is deflected downward, degrees\n\n$\\eta$ efficiency of the horizontal tail\n\n$\\mu$ viscosity of air, slugs per foot-second\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:36.848403+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 23, "total_pages": 47, "image_filename": "19930085919_p23.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:51:38.893935+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 21, "total_pages": 92, "image_filename": "19930085951_p21.jpg", "text": "NACA RM L9D29\n\n[Figure: Propeller dynamometer in test section with tunnel closed. NACA L-45764]\n\nFigure 2.— Propeller dynamometer in test section with tunnel closed.\n\nCONFIDENTIAL\nUNCLASSIFIED\n\n19", "timestamp": "2026-07-22T06:51:42.781032+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 2, "total_pages": 40, "image_filename": "19930085997_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:51:44.400224+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 53, "total_pages": 149, "image_filename": "19930083192_p53.jpg", "text": "NACA TN 1976\n\ncheck of the calculation. The results do indicate, however, that the calculations give a fair prediction of the ratio of maximum wing deflection to maximum fuselage acceleration.\n\nTime histories given in reference 38 but not shown herein indicate that the method of calculation did not predict correctly the amplitude of the wing oscillations after passing the peak load. The result is due in part to the fact that the damping coefficient was adjusted to give a good estimate for the first wing motion and would be overestimated for the subsequent vibratory motion. Putnam (reference 39) attempted to resolve this difficulty when he extended the method of reference 38. Another possible factor is that the simplifications used are still too drastic. Other results (reference 38) indicate that the use of a constant damping factor is reasonable if its magnitude is adjusted for the average effect of unsteady lift.\n\nInspection of figure 44 indicates that the dynamic-stress ratios $\\delta_{d_{\\max}} / \\delta_{st}$ and wing-tip acceleration ratios $D^2 \\delta_{w_{\\max}} / \\Delta n_{r_{\\max}}$ increase as the gradient distance decreases. Recent calculations for more modern aircraft bear out this result and indicate that elastic response is becoming increasingly important. The fuselage acceleration ratio $D^2 \\delta_{r_{\\max}} / \\Delta n_{r_{\\max}}$ did not appear to differ greatly from 1.0 and does not appear to be seriously affected by gust size.\n\nThe calculations shown in figure 45 for the effect of wing stiffness on the dynamic-stress ratios indicate that at the \"critical\" gradient distance of 10 chords, the high-frequency wing shows a dynamic-stress ratio of 14 percent below the low-frequency wing. Further analysis is needed before a definite conclusion can be reached that a reduction of wing frequency by changing wing stiffness tends to increase the dynamic-stress ratio at any gradient distance.\n\nThe results given in figure 46 illustrate the variations of the three ratios due to increasing the forward velocity of model C from 200 to 400 miles per hour. The increase in velocity together with the corresponding increase in the rate of application of the gust load would appear to result in an increase in the dynamic-stress ratio. Figure 46, however, shows that the dynamic-stress ratio does not vary much as the speed increases and this lack of variation is thought to be caused partly by the increase in aerodynamic damping with speed. Although the results for the fuselage acceleration ratio show little effect of speed, the wing-tip acceleration ratio increases from about 1.8 to 2.5 as the speed is doubled.\n\nThe calculations for model D were made to show the effect on the dynamic-stress ratio of a change in flight condition from normal gross", "timestamp": "2026-07-22T06:51:44.607617+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 20, "total_pages": 34, "image_filename": "19930085913_p20.jpg", "text": "NACA RM L9F24\n19\n\n| | 1 | 2 | 3 | 4 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 20 | Gage 20 | 30 | |\n| | | Failure | | |\n| | 1 | | | |\n\n| | 1 | 2 | 3 | 4 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 30 | 30 | 30 | 30 |\n| | 2 | | | |\n\n| | 1 | 2 | 3 | 4 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 30 | 30 | 30 | 30 |\n| | 3 | | | |\n\n| | 30 | 30 | 20 | 20 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 4 | | | |\n\n| | 30 | 30 | 20 | 20 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 5 | | | |\n\n| | 30 | 30 | 20 | 20 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 6 | | | |\n\n| | 30 | 30 | 20 | 20 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 7 | | | |\n\n| | 50 | 50 | 30 | 30 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 10 | | | |\n\n| | 50 | 50 | 30 | 30 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 11 | | | |\n\nNo records for runs 8 and 9 -- Divergence\n\n| | 30 | 30 | 20 | 20 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 12 | | | |\n\n| | 30 | 30 | 20 | 20 |\n| :--- | :--- | :--- | :--- | :--- |\n| | 13 | | | |\n\nNACA\n1- Root torsion\n2- Root bending\n3- Tip torsion\n4 Tip bending\n\nNumbers on gage traces are attenuations\n\n(a) Model A; $\\Lambda = 0^\\circ$; $e_w = -1$.\n\nFigure 1.- Oscillograph records taken at flutter.", "timestamp": "2026-07-22T06:51:45.878394+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 90, "total_pages": 114, "image_filename": "19930086061_p90.jpg", "text": "```markdown\n86\n\n| Station | y/b/2 |\n| :--- | :--- |\n| 1 | 0 |\n| 2 | 0.167 |\n| 3 | 0.333 |\n| 4 | 0.500 |\n| 5 | 0.667 |\n| 6 | 0.833 |\n| 7 | 0.916 |\n\n[Figure: A graph plotting $c_l$ versus $\\alpha, \\text{deg}$ with multiple curves corresponding to the stations listed in the legend.]\n\nFigure 35.- Variation of $c_l$ with $\\alpha$ at seven stations along the semispan of wing 2; $\\psi = 0^\\circ$.\n\nNACA RM L9J07\n```", "timestamp": "2026-07-22T06:51:46.380237+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 59, "total_pages": 72, "image_filename": "19930085491_p59.jpg", "text": "```markdown\n58\n\nAll experimental data are for a\nReynolds number of 0.62 million\n\nCONFIDENTIAL\n\n<!-- Image (68, 215, 928, 718) -->\n\nFigure 10.- Variation of parameters affected by sweep angle.\n\nCONFIDENTIAL\n\nNACA TM NO. A8J04\n```", "timestamp": "2026-07-22T06:51:49.642686+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 44, "total_pages": 92, "image_filename": "19930085870_p44.jpg", "text": "```markdown\nNACA RM No. L9D07\n45\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O CL, {□ Cm\nWedge L.E. {△ CL, {◇ Cm\n.16\n.08\nCL\n0\n-.08\n-.16\n-.24\n\n.01\nCm\n0\n-.01\n\n.06\nElliptical L.E. {O CD, {□ L/D\nWedge L.E. {△ CD, {◇ L/D\n.04\nCD\n.02\n0.8\n-6\n-4\n-2\n0\n2\n4\n6\n8\n0\nα, deg\n\n6\n4\nL/D\n2\n0\n\n[NACA logo]\n\n(d) Wing 4. w=0.765; R=1,080,000.\nFigure 6. - Continued.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:51:51.211409+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 9, "total_pages": 46, "image_filename": "19930085983_p9.jpg", "text": "NACA RM A9I27 CONFIDENTIAL 7\n\nThe elevon had sufficient pitching-moment effectiveness to provide longitudinal balance at all test Mach numbers for all positive lift coefficients at which the model had static longitudinal stability. The positive lift coefficient at which the loss of static longitudinal stability occurred (about 0.5) was reduced with increasing negative elevon deflection at a Mach number of 0.20, and generally increased with negative elevon deflection greater than -5° at higher Mach numbers. A slight forward movement of the aerodynamic center at zero lift was noted as the elevon was deflected negatively, and the movement became larger at the higher Mach numbers.\n\nThe change of elevon hinge moment with angle of attack was nearly uniform between angles of attack of -1° and +8° at a Mach number of 0.20 and between -1° and +6° for all other test Mach numbers. The variation of hinge-moment coefficient with angle of attack became considerably larger at angles of attack beyond these ranges. The sharply defined change of slope of the hinge-moment curves occurred coincidentally with the rearward movement of the aerodynamic center noted in the pitching-moment data.\n\nThe variations of lift coefficient, pitching-moment coefficient, and hinge-moment coefficient with elevon deflection are presented in figure 9 for constant angles of attack at several Mach numbers. The pitching-moment effectiveness of the elevons was generally maintained throughout the entire range of elevon deflection.\n\nThe effect of Mach number on the pitching-moment effectiveness of the elevons and on the lift coefficient for longitudinal balance is shown in figure 10. The pitching-moment effectiveness was nearly independent of Mach number at lift coefficients below 0.20 over the test range of Mach numbers. The effectiveness $-C_{m\\delta}^*$ increased with increasing Mach number at lift coefficients greater than 0.20. The lift coefficient for longitudinal balance was essentially unaffected by compressibility up to a Mach number of 0.80 for negative elevon deflection of 10° or less, and it is indicated that for negative deflections of 5° or less the lift coefficient for longitudinal balance was little affected by compressibility throughout the entire test range of Mach numbers.\n\nLift-drag ratio.— Figure 11 presents the variation of lift-drag ratio with lift coefficient for various elevon deflections at several Mach numbers. The highest maximum lift-drag ratio occurred with an elevon deflection of -5°, which suggests that increasing the wing twist would result in a higher maximum lift-drag ratio for the wing with the elevons undeflected.\n\nLateral Control\n\nElevon effectiveness and hinge moments.— Rolling-moment coefficients due to elevon deflection are presented in figure 12 as a function\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:51.298664+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 32, "total_pages": 52, "image_filename": "19930085911_p32.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 31\n\n[Figure: Graph showing two plots over time. Top plot: Net acceleration, $a_n$, g's vs. Time after release, $\\tau$, sec. Bottom plot: Free-stream Mach number, $M_0$ vs. Time after release, $\\tau$, sec. An arrow labeled \"Impact\" points to the end of the curves at approximately 50 seconds. The NACA logo is in the bottom right corner of the graph area.]\n\n(a) Resultant flight conditions.\n\nFigure 7. - Time history of flight data and performance of ram-jet unit 16-A-2.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:51:52.392025+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 22, "total_pages": 92, "image_filename": "19930085951_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:51:57.516749+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 88, "total_pages": 98, "image_filename": "19930086073_p88.jpg", "text": "86\nNACA RM A9H04\n\nLift coefficient, $C_L$\nPitching-moment coefficient, $C_m$\n\n| | | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | 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| | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | | | | | | | | | | | | |\n| | |", "timestamp": "2026-07-22T06:52:00.123487+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 8, "total_pages": 25, "image_filename": "19930085979_p8.jpg", "text": "NACA RM E9E12\n\ndeviations inside the model were greater. Another configuration consisting of 16 orifices of 15/32-inch diameter was also investigated. The best temperature distribution obtained with this configuration was slightly inferior to the optimum temperature distribution obtained with the 12-orifice configuration. The best temperature distribution with the 16-orifice configuration was obtained at higher than the calculated value of bleedback, which resulted when high plenum-chamber-gas pressures were employed in order to obtain adequate penetration.\n\nModel-air-temperature distribution. — Once the optimum temperature distribution inside the model was obtained at the calculated value of bleedback, it could be maintained for a range of tunnel-air velocities provided that the bleedback was held constant. The plenum-chamber-gas pressure corresponding to the optimum temperature distribution varied linearly with tunnel-air velocity, as in reference 2. This variation of plenum-chamber-gas pressure with tunnel-air velocity for the optimum temperature distribution is shown in figure 6 for a plenum-chamber-gas temperature of $1000^\\circ$ F. For the curve shown, the bleedback in each case was approximately 4.9 percent, which yielded average model-air-temperature rises of $50^\\circ$ F.\n\nAlthough the radial temperature distribution observed on each thermocouple rake was very good and the maximum deviation from the average temperature for the rake was only $2^\\circ$ to $4^\\circ$ F; when the temperatures of all thermocouple rakes were considered, the maximum temperature deviation was $9^\\circ$ to $10^\\circ$ F. This rather large temperature variation resulted from incomplete jet mixing at the simulated inlet screen, as evidenced by the fact that the average air temperature measured on a thermocouple rake located directly downstream of an orifice was higher than the average model-air temperature; whereas the temperature measured on a thermocouple rake located between orifices was consistently lower than the average model-air temperature. Mixing of the hot gases with the inlet air was apparently suppressed by the converging annulus formed by the accessory housing and the outer duct wall.\n\nThe use of plenum-chamber-gas pressures corresponding to the optimum temperature distribution inside the model gave insufficient penetration to provide adequate ice protection for the accessory-housing tip. Temperature distributions on the accessory housing for several values of hot-gas bleedback at a plenum-chamber-gas temperature of $1000^\\circ$ F and a tunnel-air velocity of 200 feet per second are shown in figure 7. A bleedback of approximately 6.0 percent would be required for complete icing protection of the accessory housing for this tunnel-air velocity (fig. 7). The use", "timestamp": "2026-07-22T06:52:04.172753+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 7, "total_pages": 132, "image_filename": "19930085990_p7.jpg", "text": "NACA RM A9I01 CONFIDENTIAL 5\n\np mass density of air, slugs per cubic foot\n\nMODEL AND APPARATUS\n\nThe tests were conducted in the Ames 12-foot pressure wind tunnel, which is a closed-throat variable-density wind tunnel with a low-turbulence level closely approximating that of free air.\n\nThe steel semispan model wing used for this investigation was the one used in the tests reported in reference 1 and represented a wing of aspect ratio 4 and taper ratio 0.50. The midchord line of the wing was perpendicular to the plane of symmetry. The wing profile was a faired double wedge having a thickness-chord ratio of 0.042. The horizontal tail was identical in plan form and profile to the wing and had an area equal to one quarter of the wing area. Dimensions of the semifuselage and its location with respect to the wing are given in figure 1. The semifuselage was fitted tightly to the wing and tail without fillets at the intersections. For a portion of the tests, the rear part of the fuselage was modified as shown in figures 1(b) and 2(c) to study the effects of such a modification on the pitching-moment characteristics of the model.\n\nThe wing was equipped with a full-span, constant-chord, leading-edge plain flap and a 60.9-percent-span, constant-chord, trailing-edge plain flap. The area of the leading-edge flap was 15 percent of the total area of the semispan wing and that of the trailing-edge flap was 12 percent of the total area of the semispan wing. The unsealed gaps between the flaps and the wing were 0.015 inch with the flaps undeflected.\n\nThe horizontal tail was mounted in the extended wing-chord plane (figs. 1(a) and 2(a)) and alternately 13 inches (0.696c') above the extended wing-chord plane (figs. 1(b) and 2(b)). To mount the tail above the fuselage, a bracket with a fairing body to enclose the fittings at the point of attachment of the tail surface was added to the fuselage. With the tail mounted in either position, provision was made to vary the angle of the stabilizer by pivoting it about its 50-percent-chord line.\n\nAs shown in figure 2, the semispan model was mounted with the wing perpendicular to the floor which served as a reflection plane. The gap between the model and the tunnel floor was maintained between 0.010 inch and 0.150 inch. No attempt was made to remove the tunnel-floor boundary layer which, at the location of the model, had a displacement thickness $\\delta^*$ of 0.5 inch. The velocity characteristics of the wing-fuselage wake at the longitudinal location of the horizontal tail were measured with a rake consisting of 61 total-pressure tubes and 3 static-pressure tubes. The rake was mounted from the tunnel floor with the total-pressure tubes at a position corresponding to the centroid of the semitail area.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:05.838770+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 24, "total_pages": 47, "image_filename": "19930085919_p24.jpg", "text": "```markdown\nNACA RM No. A9C21\n\nCONFIDENTIAL\n\nhinge line of split flap located on these\nconstant-percent-chord lines\n\nForward position of split\nflap normal to the air-\nstream\n\nRearward position of\nsplit flap normal to\nthe airstream\n\n<!-- Image (108, 237, 869, 752) -->\n\nDimensions in inches\n\nFlap of triangular plan form\nhas same area as the 0.25c\nsplit flap (1.561 sq ft)\n\nSection A-A\n\nFigure 4.- The geometry and positions of the split flaps on the wing.\n\nCONFIDENTIAL\n\n23\n```", "timestamp": "2026-07-22T06:52:07.146985+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 12, "total_pages": 55, "image_filename": "19930085966_p12.jpg", "text": "NACA RM L9B17 CONFIDENTIAL 11\n\nof a variable nozzle exit area. Only one point from either figure 13 or 14 corresponds in every respect with a group of conditions on the test-data plot. Such a point common to figures 12 and 13 occurs at a thrust coefficient of 0.367 and a flight Mach number of 0.375. This point has a total-pressure-recovery ratio of 99.5 percent, a total-temperature ratio of 4.75, a combustion-chamber inlet Mach number of 0.065, and a nozzle exit area equal to that of the test setup.\n\nMoving, in figure 13, from this common point to lower flight Mach numbers along a 99.5-total-pressure-recovery-ratio curve results in higher thrust coefficients, whereas in figure 12 on the 4.75-temperature-ratio curve the same procedure results in almost the same thrust coefficient. Higher thrust coefficients are obtained along the 99.5-total-pressure-recovery-ratio curve because the inlet Mach number and therefore the air flow are held constant by opening up the nozzle exit, but in the test-data curves the rate of thrust-coefficient increase is less because the nozzle exit is held constant and the inlet Mach number of air mass flow is allowed to decrease. However, continued movement to lower flight Mach numbers along the 99.5-pressure-recovery curve leads to a condition where relatively high internal losses due to maintenance of approximately a constant air flow, a large nozzle exit area, and low ram pressures combine to reduce the exit velocity to the same order of magnitude as the flight velocity, and the thrust coefficient approaches zero rapidly. This situation never occurs in the constant-nozzle-exit-area case (fig. 12) because the inlet Mach number is allowed to decrease and the internal losses stay more in proportion to the decreasing ram pressures. Movement from the common point to higher flight Mach numbers invokes arguments converse to those for movement in the opposite direction. The higher rate of decrease in thrust coefficient of the 99.5-pressure-recovery curve with respect to the 4.75-temperature-ratio curve of figure 12 is principally due to the limitation of the air flow in figure 13 caused by the decreasing nozzle exit area as compared to the constant nozzle exit area and increasing air flow of figure 12.\n\nThe final rise in thrust coefficient of the curves of figures 13 and 14 is due to the thrust coefficient being referenced to flight dynamic pressure instead of flight stagnation-pressure rise, as is evident from the dashed curve of figure 13.\n\nThe thrust coefficients shown in figures 13 and 14 for the supersonic flight range are regarded as too low to be of practical value. The possibility of increasing the thrust coefficient by increasing the inlet Mach number or air flow and/or the temperature rise will be shown in a later discussion to be remote. Therefore, it appears that the thin-plate burner does not have direct application to supersonic aircraft.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:08.625023+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 3, "total_pages": 40, "image_filename": "19930085997_p3.jpg", "text": "NACA RM A9I29 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION AT SUPERSONIC SPEEDS OF  \nSIDE SCOOPS EMPLOYING BOUNDARY-LAYER SUCTION  \n\nBy Sherman S. Edwards\n\nSUMMARY\n\nThe pressure-recovery characteristics of a model having two scoops situated on the aft portion of a long forebody and connected through diffusers to a common settling chamber were determined at Mach numbers between 1.36 and 2.01 and Reynolds numbers (based upon the length of the model ahead of the inlets) between 2.6 and 3.4 million. The boundary layer present on the forebody of the model ahead of the main scoops was removed by means of boundary-layer suction scoops. Total pressure and mass flow in the main and boundary-layer ducts were measured in tests in which the approach to the inlets and the model angle of attack were varied. The effects of interaction between the flow in the two air-induction systems and of varying the mass flow through the boundary-layer scoops were studied.\n\nAt Mach numbers less than 1.8, it was found that, by properly designing the approach to the inlets and neglecting the energy expended in boundary-layer removal, total-pressure recovery within 0.05 of that of nose inlets could be maintained over a large range of mass-flow ratios. By full-scale extrapolation of the data, it was estimated that the energy required for removal of the boundary layer was equivalent to a loss in total pressure of approximately 0.04 of the measured recovery in the main scoops. The total-pressure ratio was found to decrease with increasing positive angles of attack. An improvement in the pressure recovery occurred at angle of attack when the forebody was drooped with respect to the duct inlets.\n\nINTRODUCTION\n\nIn reference 1 it was found that the total-pressure recovery obtained with scoop inlets compared favorably with that obtained with nose inlets over a relatively large range of mass-flow ratio and up to free-stream Mach numbers of about 1.70. The results indicated that improved pressure recovery depended primarily upon boundary-layer-control measures designed both to remove low-energy air from the ducts and to prevent premature separation of the boundary layer ahead of the scoops. In the models tested, the boundary layer was diverted through slots in\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:09.221357+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 54, "total_pages": 149, "image_filename": "19930083192_p54.jpg", "text": "50\nNACA TN 1976\n\nweight to overloaded gross weight. Table XVII shows that the forward velocity is different in the two cases, but consideration of the foregoing discussion may justify the assumption that speed has a negligible effect on the findings. The results are given in figure 47 with dynamic-stress ratio plotted as a function of the gradient distance of the gust. A reduction in wing frequency brought about by the addition of mass is shown to result in an increase in the dynamic-stress ratio.\n\nAvailable information on repeated gusts such as given in table IV has been used (reference 38) to determine the dynamic effects on wing deflection in repeated gusts. The following table (from reference 38) indicates the dynamic-deflection ratios for two gusts of given intensity spaced 25 chords apart. Also given in the table are the deflection ratios for the single gusts of greater intensity which have the same probability of occurrence as the sequence of two gusts used.\n\n| Model | Condition | \"Design\" repeated gusts ($\\delta_{d_{max}}/\\delta_{st}$) | \"Design\" single gusts ($\\delta_{d_{max}}/\\delta_{st}$) |\n| :--- | :--- | :--- | :--- |\n| C | 1 | 0.96 | 1.07 |\n| D | 1 | 1.08 | .92 |\n| D | 2 | 1.10 | 1.09 |\n\nThe variation in the results indicates that the dynamic-stress ratios for a repeated gust should be investigated. The results also indicate that these dynamic-stress ratios are probably not much greater than those determined for a single gust likely to be encountered in flight.\n\nConcluding Remarks Concerning Elastic-Airplane Reactions\n\nStudy of available results and unpublished data indicates that dynamic response is becoming of greater importance with modern advances in airplane design. As mass tends to be distributed along the wing, more exact solutions are needed and, until the various factors such as unsteady-lift functions, inertia terms, elastic constants, and the span-wise gust distribution are known more exactly, such calculations should be utilized on a relative basis. Although new and more elaborate methods of calculation are available, serious problems still remain as to the basic aerodynamic and elastic coefficients.", "timestamp": "2026-07-22T06:52:09.903343+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 91, "total_pages": 114, "image_filename": "19930086061_p91.jpg", "text": "NACA RM L9J07\n\n1.4\nStation y/b/2\n1 0.167\n2 0.333\n3 0.500\n4 0.667\n5 0.833\n6 0.916\n7\n\nc_l\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\n0 4 8 12 16 20 24 28 32 36 40 44 48 52\nα, deg\n\nNACA\n\nFigure 36.- Variation of c_l with α at seven stations along the semispan of wing 3; ψ = 0°.\n\n87", "timestamp": "2026-07-22T06:52:10.561810+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 60, "total_pages": 72, "image_filename": "19930085491_p60.jpg", "text": "```markdown\nNACA RM No. A8J04\n\nAll experimental data are for a Reynolds number\nof 0.62 million except where noted.\n\nCONFIDENTIAL\n\n<!-- Image (109, 196, 499, 764) -->\n\n(c) $\\Delta C_D/(\\Delta C_L)^2$ and $k_a$\n\n<!-- Image (533, 196, 881, 764) -->\n\n(d) $(L/D)_{max}$ and $C_{L_{opt}}$\n\nFigure 10.- Concluded.\n\nCONFIDENTIAL\n\n59\n```", "timestamp": "2026-07-22T06:52:11.274882+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 9, "total_pages": 30, "image_filename": "19930085975_p9.jpg", "text": "NACA RM L9E10 CONFIDENTIAL 7\n\n$$\nC_{l\\delta} = KC_{l\\delta_{\\text{test}}}\n$$\n\nand\n\n$$\nC_{l_p} = \\frac{\\partial C_l}{\\partial \\left( \\frac{pb}{2V} \\right)} = - \\frac{C_{l\\delta}}{\\left( \\frac{pb}{2V} \\right)_\\delta}\n$$\n\nwhere the expressions $C_{l\\delta_{\\text{test}}}$ and $\\left( pb/2V \\right)_\\delta$ were evaluated graphically as the slopes of the static rolling-moment coefficient $C_l$ plotted against aileron deflection $\\delta$ and the nondimensional steady rate of rolling $pb/2V$ plotted against aileron deflection $\\delta$, respectively. This method of determining $C_{l_p}$ assumes that the effects of rolling on $C_{l\\delta}$ are negligible (except for distortion corrections previously discussed) and that $C_{l_p}$ is independent of aileron deflection.\n\nRESULTS AND DISCUSSION\n\nThe results of the investigation are presented in the following figures:\n\nFigures\n\nRolling-moment data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T06:52:13.939705+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 45, "total_pages": 92, "image_filename": "19930085870_p45.jpg", "text": "46\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\nElliptical L.E. {O CL, □ Cm}\nWedge L.E. {△ CL, ◇ Cm}\n\nCL\n.24\n.16\n.08\n0\n-.08\n-.16\n-.24\n\nCm\n.01\n0\n-.01\n\nElliptical L.E. {O CD, □ L/D}\nWedge L.E. {△ CD, ◇ L/D}\n\nCD\n.06\n.04\n.02\n\nL/D\n6\n4\n2\n0\n\nα, deg\n0.8\n-6\n-4\n-2\n0\n2\n4\n6\n8\n\n[NACA logo]\n\n(e) Wing 5. w=0.869; R = 960,000.\nFigure 6.-Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:14.127186+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 33, "total_pages": 52, "image_filename": "19930085911_p33.jpg", "text": "32\nCONFIDENTIAL\nNACA RM E9F22\n\nDiffuser Mach number, station 2, $M_2$\nDiffuser total-pressure recovery, $P_4/P_0$\nExternal-drag coefficient, $C_D$\nNet-thrust coefficient, $C_F$\n\nTime after release, $\\tau$, sec\n\n(b) Diffuser conditions and external-drag and net-thrust coefficients.\nFigure 7. - Concluded. Time history of flight data and performance of ram-jet unit 16-A-2.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:52:15.303262+00:00"}