PDF/ocr_dataset_1043.jsonl

{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 53, "total_pages": 58, "image_filename": "19930082617_p53.jpg", "text": "52\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:38:40.833760+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 15, "total_pages": 114, "image_filename": "19930086061_p15.jpg", "text": "NACA RM L9J07\n11\n\nComparison of pressure distributions and flow characteristics of\nwings 1, 2, and 3 at zero yaw.- The zero-yaw pressure distributions of\nwings 1 and 3 were quite similar in nature to those of wing 2, although\nthree primary differences were evident. First, at comparable stations\nand angles of attack the widths of the negative-pressure peaks and dips,\nmeasured in percent of chord, increased with aspect ratio with the\ngreater successive difference being between wings 1 and 2. For example,\nat $\\alpha = 14.1^\\circ$ the peak-negative pressure in the region of the separation\nvortex was at about 12, 8, and 6 percent of the chord of station 2 for\nwings 1, 2, and 3, respectively. Thus at equal angles of attack the\nvortex at any specified station was generally about the same absolute\ndistance from the leading edge of each wing. Second, the maximum\nnegative pressure coefficient increased with decreased wing aspect ratio.\nAs shown in figures 6, 9, and 11, the highest measured pressure coeffi-\ncient at station 2 ($0.167 \\text{ b/2}$) was -2.25 at $\\alpha = 24.1^\\circ$ for wing 1,\n-3.13 at $\\alpha = 34.1^\\circ$ for wing 2, and -3.50 at $\\alpha = 34.1^\\circ$ for wing 3.\nThird, the extent of tip stall was progressively greater for the wings\nof higher aspect ratio although the boundary-layer-flow tuft diagrams\nof figures 27(a), 28(a), and 29(a) indicate that the flow direction was\nsimilar for each wing. The pressure distributions of figures 6 to 26\nshow the same trend.\n\nComparison of theoretical and experimental pressure distributions\nat zero yaw.- Theoretical two-dimensional pressure distributions\n(calculated at equal $c_l$ by use of reference 10) are compared with the\nmeasured distributions for each station at $\\alpha = 4.1^\\circ$ in figures 6, 8,\nand 10 for wings 1, 2, and 3, respectively. Station 1 for all wings had\na favorable pressure gradient extending well behind midchord, as is\npredicted by the theory of reference 11. With the exception of the vortex\nregion at the leading edge, station 2 exhibited the same tendency to a\nlesser degree. Because the leading edge was swept back, the measured\nstagnation pressures were much less than 1.0q. For an infinitely long\n$60^\\circ$-sweptback airfoil, the stagnation pressure corresponding to the\nvelocity normal to the leading edge should be $0.25q$, which may be\ncompared with the values of $0.15q$ to $0.42q$ measured for the present\nwings at $\\alpha = 4.1^\\circ$.\n\nPressure Distributions and Flow Characteristics in Yaw\n\nPresentation of data.- The effects of yaw angles of $10^\\circ$, $20^\\circ$, and $35^\\circ$\non the pressure distributions of the three related wings are presented\nin figures 12 to 26, and the effects on the boundary-layer flow as\nindicated by surface tufts are shown in figures 27 to 29. The pressure\ncoefficients for positive wing yaw shown over the isometric view of the\nright semispan in figures 12 to 26 were actually measured over the left\nsemispan with the wings at equal negative yaw angles.", "timestamp": "2026-07-22T05:38:42.964694+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 73, "total_pages": 99, "image_filename": "19930082511_p73.jpg", "text": "NACA TN No. 1826\n71\n\n(a) Straight exit.\n\n(b) Bell-mouth exit.\n\n(c) Enclosed space beneath the lower free surface (two-dimensional tunnel).\n\nFigure 3.- Spillage from the lower lip of the exit.", "timestamp": "2026-07-22T05:38:45.719564+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 12, "total_pages": 46, "image_filename": "19930085972_p12.jpg", "text": "```markdown\n10\n\n<!-- Image (51, 69, 876, 841) -->\n\nFigure 1.- Drawing of variable-sweep model showing different sweep configurations and horizontal-tail locations.\n\nNACA RM L59L18\n```", "timestamp": "2026-07-22T05:38:45.946284+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 22, "total_pages": 23, "image_filename": "19930085934_p22.jpg", "text": "NACA RM E9G12\n21\n\n1164\n\nIsentropic enthalpy change, $\\Delta h_{s,a}$, Btu/lb\n\nWater-air\nratio, w/a\n0\n.02\n.03\n.04\n.05\n.06\n\nPressure ratio, $P_2/P_1$\n(b) Pressure ratio, 5.0 to 8.5.\n\nFigure 2. - Concluded. Variation of enthalpy change with pressure ratio for isentropic\ncompression. Compressor-inlet conditions: pressure, 14 inches mercury absolute;\ntemperature, 77° F; specific humidity, 0.", "timestamp": "2026-07-22T05:38:47.179903+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 49, "total_pages": 53, "image_filename": "19930082542_p49.jpg", "text": "NACA TN No. 1867\n57\n\nProperties at room temperature\nBrinell hardness\nBrinell hardness\n400\n300\n200\n100\n\nTensile strength\nStress, psi\n160,000\n140,000\n120,000\n100,000\n80,000\n60,000\n\n0.2-percent-offset yield strength\nStress, psi\n140,000\n120,000\n100,000\n80,000\n60,000\n40,000\n20,000\n\n0.02-percent-offset yield strength\nStress, psi\n140,000\n120,000\n100,000\n80,000\n60,000\n40,000\n20,000\n\nElongation\nElongation, percent\n60\n40\n20\n\nRupture properties at 1300° F\nRupture strength\nStress, psi\n70,000\n60,000\n50,000\n40,000\n30,000\n100 hr\n1000 hr\n\n100-hr rupture elongation\nElongation, percent\n40\n20\n\nAs-rolled\nAs-rolled, aged\nSolution-treated\nSolution-treated, aged\nAs-rolled, hot-cold-worked\nSolution-treated, hot-cold-worked\n\nNACA\n\nFigure 14.- Range in properties of low-carbon N-155 alloy for various types of treatment. Ranges shown are those now available and may not cover all possible treatment. (Aging treatments: 1350° - 1750° F; solution treatments: 1800° - 2300° F; hot-cold-work: 5- to 25-percent reduction at 1000° - 1800° F.)", "timestamp": "2026-07-22T05:38:48.746537+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 62, "total_pages": 66, "image_filename": "19930082914_p62.jpg", "text": "NACA TN No. 1857\n\n61\n\n[Figure: Composite interferogram of jet, 0 to $8\\frac{1}{2}$ inches from nozzle.]", "timestamp": "2026-07-22T05:38:51.531544+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 15, "total_pages": 42, "image_filename": "19930085938_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:38:52.680459+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 22, "total_pages": 42, "image_filename": "19930085914_p22.jpg", "text": "```markdown\nNACA RM A57D55\n\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n\nLift coefficient, $C_L$\n\n.04 0 -.04 -.08 for M=.20\nPitching-moment coefficient, $C_m$\n\nM=.20 M=.40 M=.60 M=.70 M=.80 M=.85 M=.89 M=.92 M=.93\n\n(c) $C_L$ vs $C_m$.\n\nFigure 4.- Concluded.\n\n[Figure: NACA logo]\n\n21\n```", "timestamp": "2026-07-22T05:38:56.230457+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 6, "total_pages": 50, "image_filename": "19930086076_p6.jpg", "text": "```markdown\n4\nNACA RM E9F09\n\nFlame holders. - The various flame-holder designs investigated are discussed in the Results and Discussion.\n\nINSTRUMENTATION\n\nStatic pressures were measured by mercury manometers at stations 1 and 2 (fig. 1). Total temperatures were measured at station 1 by two iron-constantan thermocouples and at station 3 by nine chromel-alumel thermocouples located at the approximate centers of equal areas.\n\nWater-flow rates in the quenching spray and cooling jackets and fuel-flow rates were measured by rotameters. Air-flow rates were metered with a variable, calibrated, slit-type orifice in the 12-inch inlet-air line. The air was heated to the desired inlet-air temperatures by electric heaters.\n\nAn automatic pressure recorder was used to determine the regularity and the magnitude of pressure pulsations during combustion with a few of the more efficient combustor designs. High-speed motion pictures (2400 frames/sec) were taken of combustion with one flame holder to affirm the absence of cyclic combustion.\n\nPROCEDURE\n\nThe procedure followed in the development of the flame holders was largely experimental rather than theoretical. Often more than one change in the fuel type or state, fuel injector, flame holder, or other equipment were made simultaneously in an effort to arrive at a satisfactory type of combustor in as short a time as possible.\n\nIt was possible to obtain data at only one operating condition for some flame holders because of the limited life of Inconel, the best readily obtainable metal for high temperature use. In order to make comparisons of the efficiency of different flame holders, an attempt was made to investigate all flame holders at a standard set of inlet-air conditions. The conditions selected were as follows:\n\n| | |\n| :--- | :--- |\n| Inlet-air average velocity, ft/sec | 200 |\n| Inlet-air static pressure, in./Hg absolute | 60 |\n| Inlet-air temperature, $^\\circ$F | 200 |\n| Fuel-air ratio | 0.05 |\n```", "timestamp": "2026-07-22T05:38:56.745249+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 9, "total_pages": 98, "image_filename": "19930086073_p9.jpg", "text": "NACA RM A9H04\n\nwas pointed out in reference 1 that the variation of these coefficients with sideslip was not always linear for the triangular wing with the sharp leading edge, particularly at the higher lift coefficients. Sample curves of $C_{l}$, $C_{n}$, and $C_{y}$ versus sideslip angle for the models reported herein are presented in figure 24. While some nonlinearities exist, they are not as severe as for the wing with sharp leading edge.\n\nAll three model configurations had a positive dihedral effect as will be seen from figure 23. When the body was added, $C_{l\\beta}$ became more negative, particularly at the higher values of $C_{L}$. On the other hand, $C_{l\\beta}$ became more negative at the lower values of $C_{L}$ with the addition of the vertical tail.\n\nThe wing alone was directionally stable up to the stall and the addition of the vertical tail overcame, up to a $C_{L}$ of 0.7, the directional instability caused by the body. Between a $C_{L}$ of 0.7 and 1.1, however, the directional stability of the model with body and tail decreased to zero and, by the time wing stall was reached, was considerably negative. This loss in directional stability is traceable to the increments of yawing-moment coefficient contributed by the vertical tail. Above a $C_{L}$ of 0.7 these increments decreased to nearly zero at a rate which increased with angle of sideslip. (See fig. 25.) Such a loss in directional stability is apparently connected with the effect on the vertical tail of the separation-vortex type of flow which exists over this wing. (Consult reference 3 for a description of the separation vortices.) That it was not connected with a loss in dynamic pressure at the tail is indicated by the rudder-effectiveness data as will be discussed later.\n\nThe influence of the separation vortices on the angle of attack of the vertical tail may very well account for the loss in tail effectiveness. The pattern of the separation vortices over the wing in sideslip is shown in figure 26(a). The apparent point of origin of the separation vortices moves inboard with angle of attack. In side view, the vortices form an angle with respect to the chord plane of the wing. The magnitude of this angle is approximately one-third of the angle of attack of the wing. Thus, in the view looking upstream (fig. 26(b)) the vertical displacement of the vortices, back at the tail, increases with increase in wing angle of attack. The vortex on the right side moves closer to the plane of the vertical tail than does the vortex on the left, for the model is in positive sideslip and hence the vortices are under the influence of the free-stream air flow from the right. The effective angle of attack on the upstream panel of a triangular plan-form wing in sideslip is greater than that on the downstream panel. Consequently, the strength of the vortex on the right in figure 26(b) will be greater than that of the one on the left for a given wing angle of attack. Above the core of the vortices and in the plane of the vertical tail, then, there will be a velocity component to the left which is the resultant of the velocity vectors from the two vortex flows. Below the core of the vortices there will be a velocity component to the right.", "timestamp": "2026-07-22T05:39:01.531762+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 52, "total_pages": 78, "image_filename": "19930082618_p52.jpg", "text": "```markdown\n.028\n.024\n.020\n.016\n.012\n.008\n.004\nSection drag coefficient, $c_d$\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\nFlagged symbols denote\nstandard roughness\n-.8 -.4 0 .4 .8 1.2\nSection lift coefficient, $c_l$\n\n.028\n.024\n.020\n.016\n.012\n.008\n.004\nSection drag coefficient, $c_d$\nR\n$\\nabla$ 3.0 x $10^6$\n$\\triangleright$ 6.0\n$\\triangleleft$ 9.0\nFlagged symbols denote\nstandard roughness\n-1.2 -.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\n.008\n.004\n0\n-.004\n-.008\n-.012\n-.016\nMoment coefficient, $c_{mac}$\n-.8 -.4 0 .4 .8 1.2 1.6\nSection lift coefficient, $c_l$\n\nNACA\nR a.c. position\nx/c y/c\n$\\circ$ 0.7 x $10^6$ .257 -.022\n$\\square$ 1.0 .261 -.015\n$\\diamond$ 1.5 .263 -.003\n$\\triangle$ 2.0 .266 .015\n$\\nabla$ 3.0 .258 -.106\n$\\triangleright$ 6.0 .259 -.085\n$\\triangleleft$ 9.0 .260 -.073\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic\ncenter of the plain NACA 66$_2$-415 airfoil section.\n\nFigure 10.- Concluded.\n\n50\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:39:08.745572+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 45, "total_pages": 96, "image_filename": "19930085880_p45.jpg", "text": "NACA RM No. L9C03\n43\n\n[Figure: A graph plotting Resistance (lb) on the y-axis against Wetted area (sq ft) on the x-axis. The y-axis ranges from 0 to 8. The x-axis ranges from 0 to .35. There are five curves representing different speeds (10, 15, 20, 25, 30 fps). The curves for 10, 15, 20, and 25 fps are marked with circles, squares, diamonds, and triangles respectively. The curve for 30 fps is marked with triangles and is located in the lower left portion of the graph.]\n\nResistance, lb\nSpeed (fps)\n30\n25\n20\n15\n10\nWetted area, sq ft\n(d) $\\tau = 16^d$.\nFigure 15.- Continued.\nNACA", "timestamp": "2026-07-22T05:39:11.458787+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 2, "total_pages": 36, "image_filename": "19930086097_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:39:30.868802+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 43, "total_pages": 46, "image_filename": "19930085542_p43.jpg", "text": "6E\nNACA RM No. L8L29\n41\n\n$$ \\frac{\\partial C_{Yp}}{\\partial C_L} $$\n$$ \\frac{\\partial C_{np}}{\\partial C_L} $$\n$$ C_{lp} $$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:39:34.561514+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 74, "total_pages": 99, "image_filename": "19930082511_p74.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:39:34.790197+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 23, "total_pages": 42, "image_filename": "19930085914_p23.jpg", "text": "```markdown\n22\n\nLift coefficient, $C_L$\n\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n\n0 .04 .08 .12 .16 .20 .24 for M=.20\n\nDrag coefficient, $C_D$\n\nM=.20\nM=.40\nM=.60\nM=.70\nM=.80\nM=.85\nM=.89\nM=.92\nM=.93\n\nNACA\n\n(a) $C_L$ vs $C_D$.\n\nFigure 5.- The effect of Mach number on the aerodynamic characteristics of the wing-fuselage combination at a Reynolds number of 2,000,000.\n\nNACA RM A9D25\n```", "timestamp": "2026-07-22T05:39:36.600333+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 23, "total_pages": 23, "image_filename": "19930085934_p23.jpg", "text": "22\nNACA RM E9G12\n\n<!-- Image (145, 202, 850, 736) -->\n\nFigure 3. - Variation of adiabatic efficiency with pressure ratio for impeller tip speed of 1560 feet per second and slip factor of 0.94. Compressor-inlet conditions: pressure, 14 inches mercury absolute; temperature, 77° F; specific humidity, 0.\n\nNACA-Langley - 9-7-49 - 300", "timestamp": "2026-07-22T05:39:40.193931+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 7, "total_pages": 50, "image_filename": "19930086076_p7.jpg", "text": "NACA RM E5F09\n\nUnless otherwise stated, combustion efficiency was determined at these conditions.\n\nBecause no simple method has been found for directly measuring exhaust-gas temperatures in excess of $3000^\\circ$ F, combustion efficiency was determined by means of the heat-balance method. In this method, sufficient water is introduced through the water spray to reduce the exhaust-gas temperatures to a value that can be directly measured with thermocouples. The temperature was maintained between $500^\\circ$ and $700^\\circ$ F. In this temperature range, it was assumed that all the water had been vaporized and the combustion process arrested. Combustion efficiency is the ratio of actual enthalpy rise of the exhaust products, cooling water, and steam to the heating value of the fuel.\n\nCombustion limits were determined in the following manner: Inlet-air pressure, temperature, and velocity were set at the desired values; fuel-air ratio was set at a value where combustion was stable and then was either increased or decreased until combustion ceased.\n\nRESULTS AND DISCUSSION\n\nPreliminary experiments. - Preliminary investigation of several different flame holders, both the single-stage gutter type and the multiple-stage type with gutters immersed in the combustion zone, gave similar results. Combustion was unstable at inlet-air velocities higher than 60 feet per second; and at fuel-air ratios richer than 0.025 combustion could not be stabilized at velocities greater than 40 feet per second. It was concluded that the source of combustion instability was not in the combustor but rather in the auxiliary ducting. The inlet-air duct comprised 500 feet of 12-inch pipe between the piston compressor supply and the test installation and this pipe might give rise to low-frequency pulsations.\n\nTo improve the stability, a variable-area diffuser was installed upstream of the combustor so that sonic flow could be maintained at all times in the diffuser. The effect was pronounced. A flame holder that could not be operated above the limits mentioned in the foregoing paragraph before installation of this diffuser could thereafter be operated up to inlet-air velocities of 140 feet per second at fuel-air ratios from 0.02 to 0.07. This flame holder consisted of two rows of four gutters each, as suggested by reference 1. All succeeding runs were conducted using a variable-area diffuser upstream of the combustor.", "timestamp": "2026-07-22T05:39:44.108135+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 13, "total_pages": 46, "image_filename": "19930085972_p13.jpg", "text": "NACA RM L9B18\n11\n\n[Figure: System of axes and control-surface hinge moments and deflections. Positive values of forces, moments, and angles are indicated by arrows. Positive values of tab hinge moments and deflections are in the same directions as the positive values for the control surfaces to which the tabs are attached.]\n\nFigure 2.- System of axes and control-surface hinge moments and deflections. Positive values of forces, moments, and angles are indicated by arrows. Positive values of tab hinge moments and deflections are in the same directions as the positive values for the control surfaces to which the tabs are attached.", "timestamp": "2026-07-22T05:39:50.942450+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 50, "total_pages": 53, "image_filename": "19930082542_p50.jpg", "text": "```markdown\n58\nNACA TN No. 1867\n\n<!-- Image (65, 109, 888, 997) -->\n\nFigure 15.- Typical curves of stress against rupture time for low-carbon N-155 alloy.\n```", "timestamp": "2026-07-22T05:39:51.164715+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 26, "total_pages": 36, "image_filename": "19930085912_p26.jpg", "text": "24\nNACA RM No. E9C16\n\n1105\n\nRam-pressure recovery, $\\eta$\n\nTunnel\nvelocity\n(ft/sec)\n$\\diamond$ 220\n$\\square$ 300\n$\\circ$ 380\n$\\triangle$ 450\n\n[Figure: A graph plotting Ram-pressure recovery against Cold-gas bleedback. The y-axis ranges from .74 to .98. The x-axis ranges from 0 to 12. A solid line trends downwards from approximately (0, .95) to (10.5, .75). Various data points ($\\diamond$, $\\square$, $\\circ$, $\\triangle$) are scattered around the line. A NACA logo is present in the bottom left corner of the plot area.]\n\nCold-gas bleedback, percent\n\nFigure 5. - Variation of ram-pressure recovery with cold-gas bleedback.\nFree-stream total temperature, $0^\\circ$ F; angle of attack, $0^\\circ$.", "timestamp": "2026-07-22T05:39:54.196649+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 30, "total_pages": 31, "image_filename": "19930085881_p30.jpg", "text": "28\nNACA RM L9D12\n\nCONFIDENTIAL\n\n$$\n\\frac{\\Delta p}{\\Delta p_0} / \\delta_a\n$$\n\n| Model number | $\\delta_a$ |\n| :--- | :--- |\n| 53c | $3.5^\\circ$ |\n| 53e | $5.3^\\circ$ |\n| 53f | $5.3^\\circ$ |\n| 53m | $3.7^\\circ$ |\n| 53n | $3.8^\\circ$ |\n\n$\\phi = 10.6^\\circ$\nNACA 63A009\n\n| Model number | $\\delta_a$ |\n| :--- | :--- |\n| 120b | $3.5^\\circ$ |\n| 120e | $4.7^\\circ$ |\n| 120f | $4.7^\\circ$ |\n\n$\\phi = 10.2^\\circ$\nDouble-wedge\n\n| Model number | $\\delta_a$ |\n| :--- | :--- |\n| 116a | $3.5^\\circ$ |\n| 116c | $5.0^\\circ$ |\n| 116d | $5.0^\\circ$ |\n| 116e | $3.6^\\circ$ |\n| 116f | $3.9^\\circ$ |\n\n$\\phi = 21.0^\\circ$\nNACA 16-009\n\n| Model number | $\\delta_a$ |\n| :--- | :--- |\n| 118b | $3.5$ |\n| 118c | $5.1$ |\n| 118d | $5.5$ |\n| 118e | $5.5$ |\n\n$\\phi = 20.4^\\circ$\nCircular-arc\n\nM\nNACA\n\n(b) $\\Lambda = 45^\\circ$.\nFigure 7.— Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:40:01.562566+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 54, "total_pages": 58, "image_filename": "19930082617_p54.jpg", "text": "NACA TN 1962\n53\n\n[Figure: Side view of a buckled cylindrical structure with circumferential stiffeners and rivets. The cylinder shows significant deformation and wrinkling. A small NACA logo is visible on the right side of the image.]\n\nFigure 36.- Side view of cylinder 77 after buckling.", "timestamp": "2026-07-22T05:40:11.592110+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 53, "total_pages": 78, "image_filename": "19930082618_p53.jpg", "text": "```markdown\nNACA TN 1945\n\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-.1\n0\n-.1\n-.2\n-.24\n-16 -8 0 8 16 24\nSection lift coefficient, $c_l$\nMoment coefficient, $c_{m,c/4}$\nSection angle of attack, $\\alpha_0$, deg\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\nabla$ 3.0\n$\\blacktriangle$ 4.0\n$\\blacktriangledown$ 5.0\n$\\blacklozenge$ 6.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\nFigure 11.- Aerodynamic characteristics of the NACA 0012 airfoil section, 24-inch chord.\n\n51\n```", "timestamp": "2026-07-22T05:40:14.640876+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 75, "total_pages": 99, "image_filename": "19930082511_p75.jpg", "text": "NACA TN No. 1826\n73\n\n[Figure: A black and white photograph showing a two-dimensional jet with different pressures on the two free surfaces. An arrow points to the left side of the jet.]\n\nNACA\n\nFigure 4.- Two-dimensional jet with different pressures on the two free surfaces.", "timestamp": "2026-07-22T05:40:14.899993+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 46, "total_pages": 96, "image_filename": "19930085880_p46.jpg", "text": "44\nNACA RM No. L9C03\n\nResistance, lb\nSpeed (fps)\nWetted area, sq ft\n(e) $\\tau = 20^\\circ$.\nFigure 15.- Concluded.", "timestamp": "2026-07-22T05:40:15.060645+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 44, "total_pages": 46, "image_filename": "19930085542_p44.jpg", "text": "42\nNACA RM No. J8L29\n\n50\n40\na.c., %c\n30\n20\n\n1.6\n1.4\n$C_{L_{max}}$\n1.2\n1.0\n.8\n\nExperimental\nReference 2\nReference 3\nReference 5 (Calculated)\n\n.06\n.04\n$C_{L_{\\alpha}}$\n.02\n0\n0 1 2 3 4\nAspect ratio, A\nNACA\n\nFigure 23.- Variation of aerodynamic center, $C_{L_{max}}$, and $C_{L_{\\alpha}}$ with aspect ratio for modified triangular wings. Profile, NACA 0012; $\\Lambda_{c/4} = 36.9^\\circ$; $C_L = 0$.", "timestamp": "2026-07-22T05:40:15.858153+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 16, "total_pages": 42, "image_filename": "19930085938_p16.jpg", "text": "NACA RM No. L9B04\n\n[Figure: A model of a floatplane with a single boom and tail float.]\n\n(c) With single boom and tail float. (Langley tank model 237-7F1.)\n\nFigure 1.- Concluded.\n\nNACA\nL-54129.1\n\n15", "timestamp": "2026-07-22T05:40:18.153038+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 24, "total_pages": 42, "image_filename": "19930085914_p24.jpg", "text": "NACA RM A9D25\n\nLift coefficient, $C_L$\n\n|       | -8   | -4   | 0    | 4    | 8    | 12   | 16   | 20   |\n|-------|------|------|------|------|------|------|------|------|\n| .8    |      |      |      |      |      |      |      |      |\n| .6    |      |      |      |      |      |      |      |      |\n| .4    |      |      |      |      |      |      |      |      |\n| .2    |      |      |      |      |      |      |      |      |\n| 0     |      |      |      |      |      |      |      |      |\n| -.2   |      |      |      |      |      |      |      |      |\n| -.4   |      |      |      |      |      |      |      |      |\n| -.6   |      |      |      |      |      |      |      |      |\n\nAngle of attack, $\\alpha$, deg\n\nfor M=20\n\nM=20  M=40  M=60  M=70  M=80  M=85  M=89  M=92  M=93\n\n(b) $C_L$ vs $\\alpha$\n\nFigure 5.- Continued.\n\n[NACA logo]\n\n23", "timestamp": "2026-07-22T05:40:28.098973+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 3, "total_pages": 36, "image_filename": "19930086097_p3.jpg", "text": "NACA RM A9H11\n\nCONFIDENTIAL\nUNCLASSIFIED\n\nClassification Changed to\nUNCLASSIFIED\nAuthority\n100 Dir 5200.10\nDate\n10/31/64\nBy\nS. P. Anderson\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nREDUCTION OF PROFILE DRAG AT SUPERSONIC VELOCITIES\n\nBY THE USE OF AIRFOIL SECTIONS HAVING\n\nA BLUNT TRAILING EDGE\n\nBy Dean R. Chapman\n\nSUMMARY\n\nA preliminary theoretical and experimental investigation has been made of the supersonic aerodynamic characteristics of blunt-trailing-edge airfoils. Calculations of the drag of a family of airfoils with finite trailing-edge thickness are presented for various values of the base pressure. Theoretical expressions for the lift, pitching moment, and maximum lift-drag ratio are developed using the Busemann second-order theory for two-dimensional supersonic flow. In order to compare the theoretical estimates with experimental data, measurements were taken of the lift and drag on wings of various airfoil sections at Mach numbers of 1.5 and 2.0 and at Reynolds numbers varying from 0.2 to 1.2 million. Rectangular plan forms with an aspect ratio of 4 and a thickness ratio of either 10 or 9.1 percent were used throughout the experiments.\n\nThe experimental findings are in accord with the theoretical considerations in indicating a decrease in profile drag and an increase in lift-curve slope for properly designed airfoils with moderately blunt trailing edges. As compared to a 10-percent-thick double-wedge airfoil of equal section modulus, reductions in profile drag of 15 to 31 percent have been measured in the Mach number and Reynolds number range investigated. As compared to sharp-trailing-edge airfoils in general, the experimental results showed an increase in lift-curve slope of 17 percent for a 10-percent-thick airfoil with the maximum thickness located at the trailing edge.\n\nThe minimum drag of blunt-trailing-edge airfoils depends to a large extent on the profile shape near the trailing edge. As a result, the improper design of a blunt-trailing-edge airfoil may lead to an increase in minimum drag coefficient. It is shown that, in such cases, the maximum lift-drag ratio is not necessarily reduced since the lift-curve slope may be sufficiently improved to more than compensate for a small increase in minimum drag.\n\nUNCLASSIFIED", "timestamp": "2026-07-22T05:40:34.667790+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 63, "total_pages": 66, "image_filename": "19930082914_p63.jpg", "text": "62\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:40:35.100766+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 10, "total_pages": 98, "image_filename": "19930086073_p10.jpg", "text": "8\nNACA RM A9H04\n\nConsideration is now given to the velocity vectors on two sections\nof the vertical tail; one near the top of the tail (fig. 26(c)) and\none near the base (fig. 26(d)). The angle of attack of the section near\nthe top of the tail increases with increases in the angle of attack of\nthe wing. It is quite likely, therefore, that this section stalls, if\nit were not stalled initially. (It was shown in reference 3 that the\ntips of triangular plan-form wings stall at a very low angle of attack.)\nOn the section near the base of the tail, the component of velocity\ncontributed by the vortices reverses direction with increase in angle of\nattack of the wing. As a result the angle of attack of this section of\nthe tail decreases with a consequent loss in side force produced by the\ntail. The influence of the separation vortices appears, therefore, to\naccount for the loss in tail effectiveness with increasing lift coeffi-\ncient.\n\nRudder effectiveness.- The increments of $C_l$, $C_n$, and $C_y$ per\ndegree of rudder deflection were found, on the basis of a $10^\\circ$ rudder\ndeflection, to be essentially independent of sideslip up to a $C_L$ of\n0.7 (fig. 27). Above this value of $C_L$, the curves for constant values\nof the sideslip angle are no longer coincident, particularly at the\nlarger angles of sideslip. It is of interest to note that there was no\nloss in rudder effectiveness. This is in contrast to the loss of\neffectiveness of the vertical tail when the model was at high lift coef-\nficients and would indicate that there was apparently no serious loss\nin dynamic pressure at the tail.\n\nAileron effectiveness.- Although adding the body to the wing\nreduced the flap area considerably, the moment of the flap area about\nthe fuselage center line decreased only negligibly. Thus, the increment\nof rolling moment per degree of aileron travel was nearly the same for\nthe wing plus body as for the wing alone. (See fig. 28.) In both cases\nthe ailerons were deflected approximately equal amounts in the direction\nto give positive roll. Rolling effectiveness decreased with both\nincreasing $C_L$ and $\\beta$. The yawing-moment curves of the same figure\nindicate the existence of a small amount of adverse yawing moment which\nincreased with lift coefficient, but was little affected by sideslip\nbelow 0.9 $C_L$. Certain of the curves of figure 28 exhibit nonlinearities\nnear the stall, a characteristic similar to that reported in reference 1.\n\nEstimation of Tail and Rudder Effectivenesses\n\nIt has already been pointed out that the tail on this model did not\nprovide directional stability at high lift coefficients. It is of\ninterest, however, to determine if the directional stability and rudder\neffectiveness can be predicted when the model is at zero lift.\n\nThe contribution of the vertical tail to the directional stability\nof the model can be expressed as follows:", "timestamp": "2026-07-22T05:40:37.494449+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 14, "total_pages": 46, "image_filename": "19930085972_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:40:38.407421+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 51, "total_pages": 53, "image_filename": "19930082542_p51.jpg", "text": "```markdown\nNACA TN No. 1867\n59\n\n<!-- Image (138, 116, 874, 853) -->\n\nTreatment\n$\\sigma$ Hot-rolled; hot-cold-worked\nO Hot-rolled; aged\nx Solution-treated; hot-cold-worked\n$\\square$ Solution-treated; aged\n\nFigure 16.- Relationship between 0.2-percent-offset yield strength of low-carbon N-155 alloy at room temperature and Brinell hardness.\n```", "timestamp": "2026-07-22T05:40:38.627390+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 12, "total_pages": 33, "image_filename": "19930085977_p12.jpg", "text": ".1591\nCONFIDENTIAL\n\n0.25-Chord line\n4.243\n2.166\n$\\bar{c}$\n1.944 Reference\ncenterline\n90°\n2.652\nBump surface\n7.07\nCenterline of balance\nnormal to bump surface\n11.8 Maximum diameter\n11.8\n.250\n.56\nWing-alone end plate\nWing-fuselage end plate\nCONFIDENTIAL\n\nTabulated Wing Data\nArea (Twice semispan) 0.125 sq ft\nMean aerodynamic chord 0.1805 ft\nAspect ratio 4\nTaper ratio 0.6\nIncidence 0.0°\nDihedral 0.0°\nAirfoil section parallel to\nfree stream NACA 65A006\n\nNACA RM L9H22\n\nNACA\n0 2\nScale, inches\n\nFigure 1.— General arrangement of a model with 0° sweptback wing, aspect ratio 4, taper ratio 0.6,\nand NACA 65A006 airfoil section.\n11", "timestamp": "2026-07-22T05:40:40.193204+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 8, "total_pages": 50, "image_filename": "19930086076_p8.jpg", "text": "6\nNACA RM E9F09\n\nA similar effect was sought with the exhaust nozzle by running a flame holder both with and without choking at the nozzle, but no pronounced effect could be detected. Inasmuch as these runs were conducted with air flow choked at the diffuser inlet, it is probable that air-flow pulsations were minimized at this point; therefore any possible effect of exhaust-nozzle choking was minimized. The early part of the investigation showed that stability was much improved by the use of gutters immersed in the flame, as also shown in reference 1.\n\nTwo parallel rows of gutters. - Flame holder 1 (fig. 3) consisted of two rows of gutters spaced with their centers 5 inches apart and with 11 gutters in each row. The gutters, made of 1/8-inch Inconel, were 4 inches long, 1 inch from tip to base, 3/4-inch wide, and were mounted 1 1/2 inches between tips. Combustion limits are given in figure 4. Combustion was stable for inlet-air velocities up to 325 feet per second with liquid AN-F-48b fuel at a fuel-air ratio of 0.07. At a velocity of 200 feet per second combustion was stable at fuel-air ratios as lean as 0.018. Combustion efficiency was below 30 percent at the standard operating condition.\n\nThe stability of this flame holder was adequate for application at the specified flight conditions, but combustion efficiency was very unsatisfactory. In order to determine whether any of the dimensions of this flame holder can markedly affect combustion efficiency, several flame holders were made that involved various systematic changes from flame holder 1, which was taken as a standard. These flame holders were investigated with liquid AN-F-48b fuel.\n\nFlame holder 2 contained 7 gutters in each row with the spacing between gutters increased to 2 1/2 inches to maintain the same combustor length. Flame holder 3 contained 15 gutters in each row with the spacing between gutters reduced to again give the same combustor length. Flame holder 4 contained 11 gutters in each row but had the rows closer together (4 in. between center lines of rows). Flame holder 5 contained 11 gutters in each row but had the gutter width increased to 1 inch. With all these flame holders the combustion limits were below those of flame holder 1, as shown in figure 4. Combustion efficiency was below 30 percent, indicating that combustion efficiency was not affected to a marked degree by changes in those dimensions that were varied in this series of flame holders.", "timestamp": "2026-07-22T05:40:40.827189+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 27, "total_pages": 36, "image_filename": "19930085912_p27.jpg", "text": "NACA RM No. E9C16\n25\n\n1105\n\nCold-gas\nbleedback\n(percent)\n0\n2.75\n4.12\n5.21\n6.51\n7.32\n\nPressure coefficient, S\n\nNACA\n\nFigure 6. - Effect of cold-gas bleedback on lip-pressure\ndistribution. Free-stream total temperature, 0° F; angle\nof attack, 0°.", "timestamp": "2026-07-22T05:40:43.932296+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 16, "total_pages": 114, "image_filename": "19930086061_p16.jpg", "text": "12\nNACA RM L9J07\n\nConcept of correlation between pressure distributions and flow.-\nAs observed by smoke-flow studies, the vortex increased in size on the\ntrailing semispan but became obscure on the leading semispan as the wings\nwere yawed at moderate and high angles of attack. As for zero yaw, the\npressure distributions on the trailing semispan had negative-pressure\npeaks and dips up to yaw angles of about $20^\\circ$, which indicated that the\nvortex had the characteristics of the separation vortex as discussed\nfor $\\psi = 0^\\circ$. Nevertheless, increasing the sweep of the leading edge in\nyaw gradually transformed the trailing semispan vortex into more of a\ntrailing vortex of approximately constant cross-sectional area. Thus\nat $\\psi = 35^\\circ$ there was little evidence of negative-pressure peaks or dips\non the upper surface of the trailing semispan, which indicated that the\nnature of the vortex was different from that at zero yaw.\n\nApparently as the leading edge of the leading semispan was losing\nsweep with increased yaw, the vortex became more clearly defined as\nmerely part of the bound (or lifting) vortex system. However, earlier\ntip stall of the leading semispan also occurred in yaw with the result\nthat visible indications of the vortex, as evidenced by the smoke-flow\nstudies and the pressure distributions, became unnoticeable over the\noutboard sections.\n\nPressure distributions and flow characteristics of wing 2 in yaw.-\nAs evidenced by the negative-pressure peaks and dips, the pressure\ndistributions for the low angle of attack of $4.1^\\circ$ (fig. 17) indicate\nthat the vortex generally moved increasingly forward on the leading\nsemispan and rearward on the trailing semispan as the yaw angle increased.\nThe pressure distributions over the leading semispan at the highest yaw\nangles approached those indicated by two-dimensional theory for low\nangles of attack. (See fig. 8 for $\\alpha = 4.1^\\circ$ at $\\psi = 0^\\circ$.) The extreme\noutboard stations of the left semispan were even more highly loaded than\nat zero yaw. The loading on the leading semispan increased and that on\nthe trailing semispan decreased in yaw. The airfoil sections parallel\nto the air stream changed with increasing yaw so that at $35^\\circ$ of yaw the\nleft semispan leading edge with only $25^\\circ$ of sweep was the leading edge\nof the entire wing, and the region of greatest lift over the forward\npart of these altered airfoil sections was mostly on the leading semi-\nspan. With increasing sweep of the right semispan in yaw, the peak-\nnegative-pressure region at the leading edge became smaller and did not\nexist in the extreme case of $\\psi = 35^\\circ$ when the leading edge had $95^\\circ$ of\nsweep.\n\nIn a reverse manner than that at $\\alpha = 4.1^\\circ$, the vortex moved rear-\nward on the leading semispan and slightly forward on the trailing semi-\nspan as the angle of attack was increased to $14.1^\\circ$ (fig. 18). (Data\nnot presented indicated that at $\\alpha = 8.1^\\circ$ the yaw range investigated had\npractically no effect on the vortex location.) An increase in the angle\nof yaw to $10^\\circ$ or more caused tip stall of the leading semispan as shown\nby the boundary-layer-flow diagrams of figure 28.", "timestamp": "2026-07-22T05:40:45.255923+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 24, "total_pages": 51, "image_filename": "19930085962_p24.jpg", "text": "NACA RM A52D05\nCONFIDENTIAL\n\nLift coefficient, $C_L$\n14\n12\n10\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n-10\n\n$\\delta_e$\n(deg)\n$\\circ$ 0\n$\\square$ 2\n$\\diamond$ 4\n$\\triangle$ 6\n$\\nabla$ 10\n$\\square$ 20\n$\\triangle$ 30\n\nAngle of attack, $\\alpha$, deg\n-16 -12 -8 -4 0 4 8 12 16\n\nPitching-moment coefficient, $C_m$\n.12 .08 .04 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 8. — The effect of elevator deflection on the aerodynamic characteristics of the tail at a Mach number of 0.87.\n\nCONFIDENTIAL\n23", "timestamp": "2026-07-22T05:40:45.482817+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 76, "total_pages": 99, "image_filename": "19930082511_p76.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:40:47.543356+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 55, "total_pages": 58, "image_filename": "19930082617_p55.jpg", "text": "54\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:40:47.744630+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 5, "total_pages": 28, "image_filename": "19930086092_p5.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 3\n\n$C_n$ yawing-moment coefficient $\\left( \\frac{\\text{yawing moment}}{qSb} \\right)$\n\n$C_Y$ side-force coefficient $\\left( \\frac{\\text{side force}}{qS} \\right)$\n\n$C_{l_\\beta}$ rate of change of rolling-moment coefficient with angle of sideslip, per degree\n\n$C_{n_\\beta}$ rate of change of yawing-moment coefficient with angle of sideslip, per degree\n\n$C_{Y_\\beta}$ rate of change of side-force coefficient with angle of sideslip, per degree\n\n$\\Delta C_{n_\\beta t}$ increment of the rate of change of yawing-moment coefficient with angle of sideslip due to adding the vertical tail, per degree\n\n$C_{n_{\\delta_r}}$ rate of change of yawing-moment coefficient with angle of rudder deflection, per degree\n\n$C_{h_{\\delta_r}}$ rate of change of rudder hinge-moment coefficient with angle of rudder deflection, per degree\n\n$c_{h_\\delta}$ value in two-dimensional flow of the rate of change of flap hinge-moment coefficient with angle of flap deflection, per degree\n\n$C_{h_{\\alpha_t}}$ rate of change of rudder hinge-moment coefficient with angle of attack of vertical tail, per degree\n\n$c_{h_\\alpha}$ value in two-dimensional flow of the rate of change of flap hinge-moment coefficient with angle of attack of airfoil, per degree\n\n$dC_N/d\\delta_r$ rate of change of the vertical-tail normal-force coefficient with angle of rudder deflection, per degree\n\n$(dC_N/d\\alpha)_t$ rate of change of the vertical-tail normal-force coefficient with angle of attack of the vertical tail, per degree\n\n$L/D$ ratio of lift to drag\n\n$\\alpha_\\delta$ relative rudder effectiveness $\\left[ \\frac{dC_N/d\\delta_r}{(dC_N/d\\alpha)_t} \\right]$\n\n$\\alpha$ angle of attack, degrees\n\n$\\beta$ angle of sideslip, degrees\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:40:50.289349+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 45, "total_pages": 46, "image_filename": "19930085542_p45.jpg", "text": "NACA RM No. L8L29\n43\n\n$$ \\frac{\\partial C_{n\\psi}}{\\partial C_L^2} $$\n0\n-.004\n-.008\n-.012\n\nExperimental\nReference 2\nReference 5 (Calculated)\n\n$$ \\frac{\\partial C_{l\\psi}}{\\partial C_L} $$\n.024\n.020\n.016\n.012\n.008\n.004\n0\n0 1 2 3 4\nAspect ratio A\nNACA\n\nFigure 24.- Variation of $\\partial C_{n\\psi}/\\partial C_L^2$ and $\\partial C_{l\\psi}/\\partial C_L$ with aspect ratio for modified triangular wings. Profile, NACA 0012; $\\Lambda_{c/4} = 36.9^\\circ$; $C_L = 0$.", "timestamp": "2026-07-22T05:40:52.420741+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 54, "total_pages": 78, "image_filename": "19930082618_p54.jpg", "text": "```markdown\n52\n\nSection lift coefficient, $c_l$\n2.8\n2.4\n2.0\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\n-1.2\n-1.6\n-2.0\n-2.4\n-2.8\n\nMoment coefficient, $c_{m_{c/4}}$\n.1\n0\n-.1\n-.2\n-.3\n-.4\n\nSection angle of attack, $\\alpha_{0c}$, deg\n-16\n-8\n0\n8\n16\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\nabla$ 6.0\nFlagged symbols denote\nstandard roughness\n\n[Figure: Graph showing section lift coefficient and moment coefficient versus section angle of attack for various Reynolds numbers (R). The graph contains multiple curves with different symbols representing different R values. A NACA logo is present on the graph.]\n\n(b) Section lift and pitching-moment characteristics of the NACA 0012 airfoil section with a\n0.20c simulated split flap deflected 60°.\n\nFigure 11.- Continued.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:40:58.116347+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 47, "total_pages": 96, "image_filename": "19930085880_p47.jpg", "text": "NACA RM No. L9C03\n45\n\n[Figure: A line graph plotting Trimming moment against Wetted area. The vertical axis is labeled \"Trimming moment, lb-ft\" and ranges from 0 to 18. The horizontal axis is labeled \"Wetted area, sq ft\" and ranges from 0 to .35. The graph contains five data series represented by different symbols (inverted triangles, triangles, diamonds, squares, and circles), each corresponding to a specific speed labeled on the right side: 30, 25, 20, 15, and 10 (fps). A small rectangular box is drawn in the upper left corner of the plot area. The NACA logo is present near the bottom right of the plot area.]\n\nFigure 16.- Variation of moment with wetted area, Model 250A.\n(a) $\\tau = 4^\\circ$.", "timestamp": "2026-07-22T05:40:58.899917+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 64, "total_pages": 66, "image_filename": "19930082914_p64.jpg", "text": "NACA TN No. 1857\n63\n\n$$\\rho/\\rho_{atm}$$\n\nDistance from\nnozzle, inches\n2\n2 1/2\n3 1/2\n4 1/2\n5 1/2\n6\n7 1/2\n\n$$\\sigma y/x$$\n\nFigure 15.— Variation of density ratio through mixing region. $$\\sigma = 15$$.", "timestamp": "2026-07-22T05:41:06.003786+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 25, "total_pages": 42, "image_filename": "19930085914_p25.jpg", "text": "24\nNACA RM A9D25\n\n<!-- Image (79, 110, 868, 997) -->\n\nLift coefficient, $C_L$\nPitching-moment coefficient, $C_m$\n(c) $C_L$ vs $C_m$\nfor M=.20\nFigure 5. - Concluded.", "timestamp": "2026-07-22T05:41:08.594918+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 4, "total_pages": 36, "image_filename": "19930086097_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM A9H11\n\nThe trends to be followed in designing airfoils with lower drag and improved structural characteristics are briefly discussed in light of the present results and existing knowledge about base pressure in two-dimensional flow. It is concluded that in many cases the combined structural and aerodynamic advantages offered by blunt-trailing-edge airfoils are sufficient to warrant their use as a practical wing section.\n\nINTRODUCTION\n\nThe first experimental measurements at supersonic velocities of the aerodynamic characteristics of a blunt-trailing-edge airfoil appear to have been made in 1933 by Busemann and Walchner (reference 1). In this supersonic wind-tunnel investigation a wedge airfoil was included among the various profiles tested. Since a symmetrical sharp-trailing-edge airfoil of comparable thickness was not included among the profiles investigated, very little information about the relative drag of sharp- and blunt-trailing-edge airfoils can be obtained from these early experiments. Both the theoretical and experimental results of this investigation showed, however, that the wedge airfoil produces a greater lift-curve slope than sharp-trailing-edge airfoils.\n\nSubsequent to the work of Busemann and Walchner, and prior to the relatively recent investigation of Eggers (reference 2), practically no experimental data have been published on the characteristics of airfoils with blunt trailing edges. The investigation of reference 2 was concerned with the behavior of such airfoils at subsonic, rather than supersonic, free-stream velocities. The results showed that airfoils with maximum thickness located close to the trailing edge have remarkably good lift characteristics at subsonic supercritical velocities, but have undesirably high drag coefficients throughout most of the subsonic speed range. The high drag at low subsonic speeds has been known for many years and explains why very little attention has been paid in the past to the possibilities of blunt-trailing-edge airfoils.\n\nAt supersonic speeds there is no reason to presume that an airfoil with moderately blunt trailing edge will have higher drag than an airfoil with a sharp trailing edge. On the basis of an estimate made in reference 3 of the base pressure in two-dimensional flow, it has been concluded that the opposite, in fact, is probably more often closer to the truth. In this reference it was pointed out that the use of properly chosen airfoil sections having a blunt trailing edge would substantially decrease the pressure drag of the airfoil contour forward of the base, but would not necessarily introduce excessive base drag if the boundary layer near the trailing edge were relatively thick compared to the base height. The approximate numerical calculations given therein indicated that in some cases a properly designed blunt-trailing-edge airfoil could have from 20- to 30-percent lower profile drag than a corresponding airfoil with a sharp trailing edge. The present experimental investigation has been conducted in view of the possibilities suggested by these calculations.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:41:12.823380+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 15, "total_pages": 46, "image_filename": "19930085972_p15.jpg", "text": "NACA RM L59L18\n\n[Figure: Variable-sweep model mounted on single-support strut in 300 MPH 7- by 10-foot tunnel. A = 45°; rear view.]\n\nFigure 3.- Variable-sweep model mounted on single-support strut in 300 MPH 7- by 10-foot tunnel.  \nA = 45°; rear view.\n\nNACA  \nL-53593\n\n13", "timestamp": "2026-07-22T05:41:14.262099+00:00"}