PDF/ocr_dataset_1044.jsonl

{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 52, "total_pages": 53, "image_filename": "19930082542_p52.jpg", "text": "60\nNACA TN No. 1867\n\n<!-- Image (124, 109, 905, 823) -->\n\nTreatment\nO Hot-rolled; aged\n$\\sigma$ Hot-rolled; hot-cold-worked\nx Solution-treated; hot-cold-worked\n$\\square$ Solution-treated; aged\n\nFigure 17.- Relationship between 0.02-percent-offset yield strength of low-carbon N-155 alloy at room temperature and Brinell hardness.", "timestamp": "2026-07-22T05:41:15.898429+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 13, "total_pages": 33, "image_filename": "19930085977_p13.jpg", "text": "CONFIDENTIAL\n\n4.53\n45°\n10\n76\n160\n40\n\nBump surface\n\nCenterline of balance\nnormal to bump surface\n\n118 Maximum diameter\n.250\nEnd plate used with\nfloating tail in fuselage\n\nCONFIDENTIAL\n\nNACA\n0 2\nScale, inches\n\nFigure 2.- Details of free-floating tail mounted in fuselage of a model with 0° sweptback wing, aspect\nratio 4, taper ratio 0.6, and NACA 65A006 airfoil section.\n\n12\nNACA RM L9E22", "timestamp": "2026-07-22T05:41:17.865658+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 31, "total_pages": 31, "image_filename": "19930085881_p31.jpg", "text": "NACA Langley - 7-22-49 - 375\n\nCONFIDENTIAL\n\n(a) $\\Lambda = 0^\\circ$\n\n(b) $\\Lambda = 45^\\circ$\n\nFigure 8.— Summary of drag results for each configuration.\n\nCONFIDENTIAL\n\nNACA RM L9D12\n\n29", "timestamp": "2026-07-22T05:41:18.285962+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 25, "total_pages": 51, "image_filename": "19930085962_p25.jpg", "text": "24\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (167, 109, 801, 879) -->\n\nFigure 8. — Concluded.\n(b) $C_L$ vs $C_D$.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:41:21.694229+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 28, "total_pages": 36, "image_filename": "19930085912_p28.jpg", "text": "26\nNACA RM No. E9C16\n\n1105\n\nPlenum-chamber gas pressure, lb/sq ft absolute\nTunnel velocity, ft/sec\n\n[Figure: Graph showing a linear relationship between plenum-chamber gas pressure and tunnel velocity. Data points are plotted at approximately (200, 3200), (275, 3800), (350, 4500), and (410, 5000). A straight line is drawn through these points. The NACA logo is present in the bottom right corner of the graph area.]\n\nFigure 7. - Relation between plenum-chamber gas pressure and tunnel velocity corresponding to optimum temperature distribution in model. Plenum-chamber gas temperature, 1000° F.", "timestamp": "2026-07-22T05:41:22.759419+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 9, "total_pages": 50, "image_filename": "19930086076_p9.jpg", "text": "NACA RM E5F09\n\nThree rows of gutters. - A third row of gutters, similar to the two in flame holder 1, was placed in the middle of that design to compose flame holder 6. This flame holder was investigated with both vaporized isopentane and liquid AN-F-48b fuel and was found to have an efficiency of 50 percent with each fuel at an inlet-air pressure of 55 inches of mercury absolute. Combustion limits were below those of flame holder 1, as shown in figure 4. An indication of combustion efficiency was the condition of the flame holder at the end of approximately 10 minutes of operation. The deterioration of flame holder 6 (fig. 5) showed that higher combustion efficiency had been obtained with this design than with previous designs, which remained intact.\n\nStaggered gutters. - Alternate gutters were staggered 1/4 inch from the row center line in each of the two rows of 11 gutters in flame holder 7 (fig. 6). The results of runs with this flame holder using liquid AN-F-48b fuel showed that the combustion efficiency was higher than that of any previously mentioned two-row design. Figure 7 is a photograph of flame holder 7 after 10 minutes of operation. The deterioration of the Inconel gutters gave evidence of temperatures in excess of $2600^\\circ$ F.\n\nThese results indicated that it was beneficial to have a continuous path of low stream velocity from the downstream end, which is the hottest and the best source of continuous ignition, to the upstream end, which is cooled the most by incoming air. This low-velocity path allows the flame to propagate as far upstream as possible toward the point of fuel injection. Any externally induced pressure fluctuation, which might cause flame blow-out at the upstream end, would have no lasting effect, because this continuous flame path would reestablish burning immediately. The incandescent downstream portions of the flame holder would provide immediate reignition, should a pressure pulse cause the entire flame to blow out.\n\nThe staggering of the gutters in flame holder 7 provided a continuous flow path by sheltering alternate edges of the gutters by the ones immediately upstream.\n\nSimultaneous use of three rows and staggered gutters. - The improved results obtained separately by the addition of the third row of gutters and the staggering of the gutters indicated the design of flame holder 8, which consisted of three rows of staggered gutters (fig. 8). Alternate gutters were staggered 3/32 inch from the row center line in each of the three rows of 11 gutters. Investigation of this flame holder with vaporized isopentane injected at the upstream end of the flame holder yielded 65-percent", "timestamp": "2026-07-22T05:41:23.939565+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 77, "total_pages": 99, "image_filename": "19930082511_p77.jpg", "text": "NACA TN No. 1826\n75\n\n[Figure: Diagram (a) showing a wing section between two curved streamlines with circulation symbols above and below]\n\n(a) Deformation due to image system; no downwash at the wing itself.\n\n[Figure: Diagram (b) showing a wing section between two curved streamlines, with a NACA logo in the lower right]\n\n(b) Undeformed upstream flow; downwash at wing is half of that at infinity.\n\nFigure 5.- Two-dimensional open tunnel of infinite length.", "timestamp": "2026-07-22T05:41:24.651983+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 56, "total_pages": 58, "image_filename": "19930082617_p56.jpg", "text": "NACA TN 1962\n\n55\n\n[Figure: Bottom view of cylinder 77 after buckling.]\n\nFigure 37. - Bottom view of cylinder 77 after buckling.", "timestamp": "2026-07-22T05:41:26.234351+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 17, "total_pages": 114, "image_filename": "19930086061_p17.jpg", "text": "NACA RM L9J07\n\nThe same trends in vortex flow shown for $\\alpha = 14.1^\\circ$ were prevalent at $\\alpha = 24.1^\\circ$. The negative-pressure peaks and dips over the leading semispan became less pronounced in yaw so that at $\\psi = 20^\\circ$ they were not evident. As discussed in the previous section, two apparent reasons are the transformation of the separation vortex to a bound vortex and the large extent of outboard stall. A schematic sketch of the vortex flow as observed by smoke studies at $\\psi = 20^\\circ$ and $\\alpha = 20^\\circ$ is given as figures 30(a) and 30(e). The smoke studies revealed considerable flow of air around the leading edge from the under surface of the trailing semispan into the top of the large vortex and, therefore, indicated that the trailing semispan vortex was transforming into a trailing vortex. This leading-edge flow apparently accounts for the negative values of $P$ on the lower surface of the trailing semispan in the leading-edge region (fig. 19). When the angle of attack was increased to $34.1^\\circ$ for yaw angles of $10^\\circ$ or greater, there was no indication of a vortex on the stalled leading semispan as evidenced by the pressure distributions and the smoke-flow studies. The large trailing-semispan vortex had the characteristics of a trailing vortex with essentially a constant cross-sectional area along the wing leading edge. There was a very strong flow of air around the leading edge into the vortex as mentioned in the previous paragraph for $\\alpha = 24.1^\\circ$. At $\\alpha = 44.1^\\circ$, visible indications of the separation vortex had dissipated and the flow over the wing was completely stalled or unsteady for all yaw angles investigated.\n\nComparison of pressure distributions and flow characteristics of wings 1, 2, and 3 in yaw.-- The effects of yaw on wings 1 and 3 were quite similar to the effects on wing 2. The three principal differences among the pressure distributions and flow characteristics of the related wings noted at zero yaw generally prevailed throughout the yaw range tested. First, for identical stations and equal angles of attack and yaw, the vortex was farther back in percent of the shorter chords of the wings of higher aspect ratio. Second, the highest negative values of the pressure coefficient increased with decreased wing aspect ratio and also were higher at moderate yaw angles than at $\\psi = 0^\\circ$. Thus for station 2 of the trailing semispan at $\\psi = 10^\\circ$ and $\\alpha = 34.1^\\circ$, pressure coefficients of -3.45, -4.10, and -4.22 were measured for wings 1, 2, and 3, respectively. Third, the characteristic more pronounced tip stall with increasing wing aspect ratio observed at zero yaw was also evident for the leading-semispan tip in yaw (figs. 27, 28, and 29). Peculiarities were noted at $\\alpha = 14.1^\\circ$ where wing 3 had more tip stall than wing 2 for $\\psi = 20^\\circ$ and as much or more tip stall than both wings 1 and 2 for $\\psi = 35^\\circ$. Contrary to the usual case, station 6 of the leading semispan of wing 1 actually unstalled in going from $\\psi = 20^\\circ$ to $\\psi = 35^\\circ$ at $\\alpha = 14.1^\\circ$.\n\nEffects of vortex flow on airplane stability and control.-- A thorough understanding of the flow about a highly swept wing has special significance. In particular, if controls were located in the field of", "timestamp": "2026-07-22T05:41:31.925407+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 48, "total_pages": 96, "image_filename": "19930085880_p48.jpg", "text": "46\nNACA RM No. L9C03\n\n18\n16\n14\n12\n10\n8\n6\n4\n2\n0\nTrimming moment, lb-ft\n\nSpeed\n(fps)\n30\n25\n20\n15\n10\n\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n(b) $\\tau = 8^\\circ$.\nFigure 16.- Continued.\nNACA", "timestamp": "2026-07-22T05:41:34.760288+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 65, "total_pages": 66, "image_filename": "19930082914_p65.jpg", "text": "64\nNACA TN No. 1857\n\n<!-- Image (123, 110, 907, 997) -->\n\nFigure 16.- Theoretical velocity distribution for incompressible mixing zone from reference 1.", "timestamp": "2026-07-22T05:41:36.206589+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 6, "total_pages": 28, "image_filename": "19930086092_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM A5F14\n\n$\\Gamma_e$\neffective dihedral, degrees\n\n$\\delta_r$\nrudder angle, measured in a plane perpendicular to the hinge line, degrees\n\n$\\Lambda_{c/4}$\nangle of sweepback of quarter-chord line, degrees\n\n$\\sigma$\nangle of sidewash, degrees\n(Negative when it increases the angle of attack of the vertical tail.)\n\n$d\\sigma/d\\beta$\nrate of change of angle of sidewash with angle of sideslip\n\n$q$\ndynamic pressure, pounds per square foot\n\n$q_t/q$\nratio of the effective dynamic pressure at the vertical tail to the free-stream dynamic pressure\n\n$A_t$\naspect ratio of swept-back vertical tail\n\n$S$\nwing area, square feet\n\n$S_t$\nvertical-tail area, square feet\n\n$b$\nwing span measured perpendicular to the plane of symmetry, feet\n\n$c$\nlocal chord, feet\n\n$\\bar{c}$\nwing mean aerodynamic chord $\\left( \\frac{\\int_0^{b/2} c^2 dy}{\\int_0^{b/2} c \\, dy} \\right)$, feet\n\n$l_t$\ndistance from the quarter-chord point of the mean aerodynamic chord of the wing to the rudder hinge line at the mean aerodynamic chord of the vertical tail\n\n$y$\nspanwise coordinate perpendicular to plane of symmetry, feet\n\nMODEL\n\nThe wing-fuselage combination was the same as that used in reference 2. The wing had the leading edge swept back $63^\\circ$, an aspect ratio of 3.5, a taper ratio of 0.25, no twist, and no dihedral. It had an NACA 64A006 section in a streamwise direction. The fuselage had a fineness ratio of 12.5, an elliptical lengthwise section, and a circular cross section. The wing was mounted on the fuselage center line with zero incidence.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:41:38.285696+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 55, "total_pages": 78, "image_filename": "19930082618_p55.jpg", "text": "```markdown\nNACA TN 1945\n\nSection drag coefficient, $c_d$\nSection lift coefficient, $c_l$\n\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\nSection drag coefficient, $c_d$\nSection lift coefficient, $c_l$\n\nR\n$\\nabla$ 3.0 x $10^6$\n$\\Delta$ 6.0\n$\\triangle$ 9.0\nFlagged symbols denote\nstandard roughness\n\nMoment coefficients, $c_{m_{ac}}$\nSection lift coefficient, $c_l$\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$\\Delta$ 6.0\n$\\triangle$ 9.0\n\na.c. position\nz/c y/c\n.285 -.018\n.286 .002\n.286 .011\n.286 .021\n.250 0\n.250 0\n.250 0\n\n(c) Section drag characteristics and section pitching-moment characteristics about the aerodynamic center of the plain NACA 0012 airfoil section.\n\nFigure 11.— Concluded.\n\n53\n```", "timestamp": "2026-07-22T05:41:39.031298+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 26, "total_pages": 42, "image_filename": "19930085914_p26.jpg", "text": "NACA RM A9D25\n25\n\nMinimum drag coefficient, $C_{D_{min}}$\n.02\n.01\n0\n0 .2 .4 .6 .8 1.0\nMach number, M\n$R=2.0 \\times 10^6$\n\n.02\n.01\n0\n0 2 4 6 8 10\nReynolds number, $R \\times 10^{-6}$\n$M=.20$\nNACA\n\nFigure 6.- The variation of the minimum drag coefficient of the wing-fuselage combination with Mach number and Reynolds number.", "timestamp": "2026-07-22T05:41:40.417712+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 16, "total_pages": 46, "image_filename": "19930085972_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:41:41.906323+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 46, "total_pages": 46, "image_filename": "19930085542_p46.jpg", "text": "```markdown\n44\nNACA RM No. L8L29\n\n$$\n\\frac{\\partial C_{Yp}}{\\partial C_L}\n$$\n[Graph: Y-axis from 0 to 1.6, X-axis from 0 to 4. Curves show variation of $\\partial C_{Yp}/\\partial C_L$ with Aspect ratio.]\n\n$$\n\\frac{\\partial C_{np}}{\\partial C_L}\n$$\n[Graph: Y-axis from -0.4 to 0, X-axis from 0 to 4. Curves show variation of $\\partial C_{np}/\\partial C_L$ with Aspect ratio.]\n\n- Experimental\n- - - - - - Reference 2\n- - - - - Reference 4\n- - - - - Reference 5 (Calculated)\n\n$$\nC_{lp}\n$$\n[Graph: Y-axis from -0.4 to 0, X-axis from 0 to 4. Curves show variation of $C_{lp}$ with Aspect ratio.]\n\nAspect ratio, A\nNACA\n\nFigure 25.- Variation of $\\partial C_{Yp}/\\partial C_L$, $\\partial C_{np}/\\partial C_L$, and $C_{lp}$ with aspect ratio for modified triangular wings. Profile, NACA 0012; $\\Lambda_{c/4} = 36.9^\\circ$; $C_L = 0$.\n\nNACA - Langley Field, Va.\n```", "timestamp": "2026-07-22T05:41:43.632442+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 5, "total_pages": 36, "image_filename": "19930086097_p5.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 3\n\nConsequently, the primary purpose of the present investigation is to determine experimentally if a properly designed airfoil with moderately blunt trailing edge can have lower profile drag at supersonic velocities than a corresponding sharp-trailing-edge airfoil. Two additional purposes of the present report are: (1) to discuss qualitatively some of the more important parameters that are expected to affect the drag of blunt-trailing-edge airfoils, and (2) to make a cursory examination of the lift, lift-drag ratio, and pitching-moment characteristics of these airfoils in supersonic flow.\n\nIn practical applications, the many structural and aerodynamic factors which affect the selection of an airfoil are far too diverse to allow all-inclusive statements to be made about the superiority of one profile shape over another. Even if a selection is made on the basis of drag considerations alone without regard to lift, moment, or lift-drag ratio, then the optimum airfoil will vary with the particular structural criterion governing a given design. Notwithstanding these complications, certain simplified criteria for making drag comparisons can be used which closely represent a few of the numerous practical applications and approximately represent many others. The simplified criterion of equal section modulus is primarily used in this report, although in a few cases comparisons are made on the basis of equal thickness ratio. In most practical cases the actual drag reduction obtainable is believed to be greater than that indicated by the criterion of equal thickness ratio, since the structural properties of blunt-trailing-edge airfoils are generally superior to those of conventional sections.\n\nIt is emphasized that the airfoil shapes investigated in these preliminary tests aim solely at demonstrating certain principles, and do not aim at providing a near optimum airfoil section. The comparisons given herein attempt to illustrate only the approximate magnitude of drag reduction that may be possible in some cases. In considering the general possibilities of blunt-trailing-edge airfoils as practical wing sections, the structural characteristics must always be kept in mind. Moreover, in viewing the experimental results of this investigation it should be remembered that future research undoubtedly will provide means of achieving lower drag while still maintaining the structural advantages of these airfoil sections.\n\nSYMBOLS AND NOTATION\n\n| Symbol | Description | Formula |\n| :--- | :--- | :--- |\n| A | cross-sectional area of airfoil profile | $$\\left[ \\int_{0}^{c} (y_u + y_l) dx \\right]$$ |\n| c | airfoil chord | |\n| $c_d$ | section drag coefficient | |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:41:55.513546+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 26, "total_pages": 51, "image_filename": "19930085962_p26.jpg", "text": "```markdown\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n-1.0\n\n-12 -8 -4 0 4 8 12\nAngle of attack, $\\alpha$, degrees\n\n.08 .04 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28\nPitching-moment coefficient, $C_m$\n\n$\\delta_e$\n(deg)\nO 0\n$\\square$ 2\n$\\diamond$ 4\n$\\nabla$ 6\n$\\triangle$ 10\n$\\blacktriangle$ 20\n$\\blacktriangledown$ 30\n\nCONFIDENTIAL\nLift coefficient, $C_L$\n\nNACA RM A9E05\nCONFIDENTIAL\n\n[Figure: Graph showing two plots. Left plot: $C_L$ vs $\\alpha$ for various $\\delta_e$. Right plot: $C_L$ vs $C_m$ for various $\\delta_e$.]\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\nFigure 9. — The effect of elevator deflection on the aerodynamic characteristics of the tail at a Mach number of 0.90.\n\nNACA\n25\n```", "timestamp": "2026-07-22T05:41:56.528264+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 29, "total_pages": 36, "image_filename": "19930085912_p29.jpg", "text": "NACA RM No. E9C16\n27\n\n<!-- Image (137, 183, 884, 766) -->\n\n(a) Plenum-chamber gas pressure, 3010 pounds\nper square foot absolute; rise in model-air\ntotal temperature, 58° F; bleedback, 3.14\npercent.\n\n(b) Plenum-chamber gas pressure, 3950 pounds\nper square foot absolute; rise in model-air\ntotal temperature, 46° F; bleedback, 4.28\npercent.\n\n(c) Plenum-chamber gas pressure, 5100 pounds\nper square foot absolute; rise in model-air\ntotal temperature, 64° F; bleedback, 6.38\npercent.\n\n(d) Plenum-chamber gas pressure, 5840 pounds\nper square foot absolute; rise in model-air\ntotal temperature, 74° F; bleedback, 7.49\npercent.\n\nFigure 8. - Effect of plenum-chamber gas pressure on temperature ratio $T_x/T_{av}$ at thermocouple\ncross rake. Tunnel velocity, 290 feet per second; plenum-chamber gas temperature, 1000° F;\nangle of attack, 0°.", "timestamp": "2026-07-22T05:41:57.532368+00:00"}
{"citation_id": "19930091987", "source_url": "https://ntrs.nasa.gov/api/citations/19930091987/downloads/19930091987.pdf", "page_number": 1, "total_pages": 12, "image_filename": "19930091987_p1.jpg", "text": "N62 50922\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nREPORT 922\n\nCHARACTERISTICS OF LOW-ASPECT-RATIO WINGS\nAT SUPERCRITICAL MACH NUMBERS\n\nBy JOHN STACK and W. F. LINDSEY\n\n[Figure: Seal of the National Advisory Committee for Aeronautics]\n\nCASE FILE\nCOPY\n\n1949\n\nFor sale by the Superintendent of Documents, U. S. Government Printing Office, Washington 25, D. C. Yearly subscription, $3; foreign, $4.50; single copy price varies according to size. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T05:41:59.195245+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 14, "total_pages": 33, "image_filename": "19930085977_p14.jpg", "text": "CONFIDENTIAL\n\nWing chord plane\nextended at CC = 0°\n\nFloating-tail geometry\nArea (Twice semispan) 0.0178 sq ft\nAspect ratio 4.0\nTaper ratio 0.60\n\nNACA RM L9E22\n\n0.25 c of model\n4.53\n0.25-Chord line\n45°\n0.60\n1.60\nB\n0.80\nB\n0.76\n1.0\nc_t\nBump surface\n1/8 Diameter\nPivot center\nCONFIDENTIAL\n\nSection B-B\n1/16\n\nNACA\n\n0 1 2\nScale, inches\n\nFigure 3.— Details of free-floating tails used in surveys behind model with 0° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section.\n\n13", "timestamp": "2026-07-22T05:42:01.942624+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 11, "total_pages": 98, "image_filename": "19930086073_p11.jpg", "text": "```markdown\nNACA RM A9H04\n\n$$\nC_{n_{\\beta_t}} = C_{N_{\\alpha_t}} \\frac{S_t}{S} \\frac{l}{b} \\left( 1 + \\frac{d\\sigma}{d\\beta} \\right) \\frac{q_t}{q} \\tag{1}\n$$\n\nThe equation to predict rudder effectiveness is\n\n$$\n\\frac{dC_n}{d\\delta_r} = - C_{N_{\\alpha_t}} \\frac{d\\alpha_t}{d\\delta_r} F \\frac{S_t}{S} \\frac{l}{b} \\frac{q_t}{q} \\tag{2}\n$$\n\nThe major problem in applying these two equations is in selecting the effective area and aspect ratio of the vertical tail. For most conventional airplane designs the methods of selection have been fairly well established. These methods do not appear applicable, however, to designs similar to the type under discussion. For this type, it is believed that the effective tail area extends to the fuselage center line. With a complete end-plate effect, as in the case where the wing trailing edge extends beyond the tail trailing edge, the tail area covered by the fuselage should be fully effective. This is indicated by the fact that a similar area of the wing was found to be fully effective. With the present wing-tail arrangement, this area of the tail was probably somewhat less fully effective; that is, the effective aspect ratio was somewhat less than twice the geometric aspect ratio. The actual value could not be established without recourse to the experimental data. The increment of $C_{Y_{\\beta}}$ due to the tail, expressed in terms of the tail lift-curve slope and compared with theoretical values for triangular wings (reference 5), indicates that the effective aspect ratio was 1.3.\n\nWith effective tail area and aspect ratio established, the values of the other factors in the two equations were then selected. The value of the tail length $l$ in equation (1) was taken as the distance from the model moment center to the theoretical center of pressure of the tail (reference 5); for equation (2) the distance was to the rudder hinge line. The value of $d\\sigma/d\\beta$ was assumed zero and $q_t/q$ was assumed to be unity, since the wing was at zero lift and the fuselage effect was considered negligible. The value of $d\\alpha_t/d\\delta_r$ was assumed to be the same as that measured on a triangular wing of aspect ratio 2 (reference 4), which had the same geometrical relation between flap and wing as between rudder and tail in the present case. This value was reduced by the factor $F$ of equation (2) or the ratio of the exposed rudder area to the total rudder area.\n```", "timestamp": "2026-07-22T05:42:03.220958+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 17, "total_pages": 42, "image_filename": "19930085938_p17.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:42:03.694579+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 57, "total_pages": 58, "image_filename": "19930082617_p57.jpg", "text": "56\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:42:04.703474+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 78, "total_pages": 99, "image_filename": "19930082511_p78.jpg", "text": "76\nNACA TN No. 1826\n\n<!-- Image (223, 139, 751, 210) -->\n(a) Two-dimensional vortex, showing current lines between the\ntwo plates.\n\n<!-- Image (194, 293, 745, 359) -->\n(b) Two-dimensional vortex in a perturbation field having a horizontal\nvelocity component.\n\n<!-- Image (223, 436, 791, 578) -->\n(c) Three-dimensional element of lift.\n\n<!-- Image (230, 632, 766, 754) -->\n(d) Horseshoe vortex of finite span.\n\nFigure 6.- Velocity-potential analogies for two- and three-dimensional\nlifting elements.", "timestamp": "2026-07-22T05:42:10.503770+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 49, "total_pages": 96, "image_filename": "19930085880_p49.jpg", "text": "NACA RM No. L9C03\n47\n\n[Figure: A line graph plotting Trimming moment against Wetted area. The y-axis ranges from 0 to 18 lb-ft. The x-axis ranges from 0 to .35 sq ft. There are five curves representing speeds of 10, 15, 20, 25, and 30 fps. The data points are marked with circles, squares, diamonds, triangles, and inverted triangles respectively. A legend box is present in the upper left corner.]\n\nTrimming moment, lb-ft\nSpeed (fps)\n30\n25\n20\n15\n10\nWetted area, sq ft\n(c) $\\tau = 12^\\circ$.\nFigure 16.- Continued.\nNACA", "timestamp": "2026-07-22T05:42:11.181395+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 66, "total_pages": 66, "image_filename": "19930082914_p66.jpg", "text": "```markdown\nNACA TN No. 1857\n65\n\n<!-- Image (85, 110, 856, 936) -->\n\nFigure 17.- Variation of velocity ratio through mixing region. $\\sigma = 15$.\n```", "timestamp": "2026-07-22T05:42:11.680488+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 10, "total_pages": 50, "image_filename": "19930086076_p10.jpg", "text": "8\nNACA RM E9F09\n\ncombustion efficiency at an inlet temperature of $280^\\circ$ F. The flame holder deteriorated slightly during this run.\n\nA second run was made with the isopentane fuel vapor injected 6 inches upstream, but the recorded data were invalidated because of failure of the water spray, which was required for the heat-balance method of determining combustion efficiency. The flame holder was severely deteriorated; only the three end gutters of the original 33 remained intact (fig. 9).\n\nOther configurations. - In an attempt to develop a flame holder with an improved continuous flame path, two designs incorporating baffles instead of gutters were evolved. Flame holder 9 consisted of three rows of staggered plates, 30 plates in each row, with a slit between each plate (fig. 10). These slits permit continuous flame propagation between all regions of low velocity.\n\nA model was constructed of Inconel. This model was investigated with vaporized isopentane injected at the upstream end of the flame holder. The plates glowed intensely and rapidly melted and broke away. The life of the plates in the downstream portion of the combustor was about 1/2 minute and was insufficient to permit data to be taken. Subsequent data taken after approximately the 4 minutes of operation required to set and stabilize conditions showed that combustion efficiency was 57 percent at an inlet-air temperature of $240^\\circ$ F and an inlet-air pressure of 56 inches of mercury absolute. The ratio of the total-pressure loss through the combustor to the inlet dynamic pressure was 3. Figure 11 shows the condition of flame holder 9 immediately after data were obtained. This figure also shows the position of the fuel injector. Inasmuch as the flame holders melted rapidly before data could be recorded, higher combustion efficiencies probably existed at the start of the run than were indicated by the data.\n\nFlame holder 10 consisted of three corrugated Inconel strips with three 1/2-inch holes in each plane of corrugation (fig. 12), which are designed to serve the same function as the slits in flame holder 9.\n\nInvestigation of flame holder 10 was made with vaporized isopentane injected 6 inches upstream of the flame holder. Choking at the exit and failure of the exhaust-nozzle mechanism limited the inlet-air velocity to 150 feet per second. Subsequent runs were made without a variable-area exhaust nozzle. At an inlet-air velocity of 150 feet per second and an inlet temperature of", "timestamp": "2026-07-22T05:42:12.782778+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 27, "total_pages": 42, "image_filename": "19930085914_p27.jpg", "text": "26\nNACA RM A9D25\n\n<!-- Image (196, 98, 782, 801) -->\n\nFigure 7.- The variation of lift-curve slope of\nthe wing-fuselage combination with Mach\nnumber and Reynolds number.", "timestamp": "2026-07-22T05:42:15.536619+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 18, "total_pages": 114, "image_filename": "19930086061_p18.jpg", "text": "14 NACA RM L9J07\n\ninfluence of a vortex, the growth and development of the vortex flow, as on the triangular wings reported herein, would be expected to have first-order effects on the stability and control of an airplane and also on the effectiveness of the controls. As found recently in a low-speed investigation of a small-scale wing having NACA 65-006.5 sections (reference 12), serious discontinuities in the lift, pitching-moment, and damping coefficient curves occurred for particular installations of outboard vertical fins. The significance of these particular results as applied to the full-scale wing is not clear at this time due to inadequacy of large-scale information, but evidently the presence of a vortex on a large-scale wing, as has been observed for wings having sharp-edge sections, would be expected to influence largely the low-speed characteristics of wings having outboard fins, nacelles, or other similar protuberances. The recent investigation of reference 7 for a wing with $47\\frac{1}{2}^\\circ$ leading-edge sweep and with 10-percent-thick biconvex sections has shown the same characteristic pressure distributions as described in this paper. Although the flow was not investigated in detail on the large-scale wing, the presence of a strong vortex was immediately evident in explorations of smaller models of the same plan form.\n\nSECTION LIFT CHARACTERISTICS AT ZERO YAW\n\nAs the angle of attack was increased, the spanwise position of maximum $c_l$ and the extent of reduced tip effectiveness moved inboard on each wing at a rate increasing with increased aspect ratio (figs. 31 to 33). The relatively high tip loading is shown for $\\alpha = 4, 10^\\circ$, as previously discussed, but at $\\alpha = 8, 10^\\circ$ the tips of each wing incurred loss of lift, with the loss being much more severe for the high aspect ratio wing 1.\n\nAs shown again effectively in the curves of $c_l$ against $\\alpha$ of figures 34 to 36, this loss of lift accompanying the collapse or movement off the wing of the negative-pressure peak occurred more rapidly with increased distance from the plane of symmetry of each wing and, as already noted, more rapidly for the wings of higher aspect ratio where the vortex swept behind the wing tips sooner. As noted in the section entitled \"Pressure Distributions and Flow Characteristics at Zero Yaw,\" this primary tip stall occurred apparently from leading-edge separation but did not alter the strong spanwise boundary-layer flow as could be determined visibly by surface tufts. However, with increased wing angle of attack the $c_l$ values for the outboard stations increased again even though the sections became visibly stalled as evidenced by surface tufts.", "timestamp": "2026-07-22T05:42:16.883658+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 53, "total_pages": 53, "image_filename": "19930082542_p53.jpg", "text": "NACA TN No. 1867\n61\n\n<!-- Image (72, 225, 886, 782) -->\n\nFigure 18.- Relationship between rupture strength of low-carbon N-155 alloy\nat 1200° F and Brinell hardness.", "timestamp": "2026-07-22T05:42:17.341340+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 17, "total_pages": 46, "image_filename": "19930085972_p17.jpg", "text": "```markdown\nNACA RM L9B18\n\nSharp-leading-edge section\n7.92\nA\n1.50\n$\\Lambda=0^\\circ$, wing station 15.06\n$\\Lambda=0^\\circ$, wing station 13.24\n0.25 chord line\nWing-panel pivot point\n0.40 chord line\nSee Table I for area of cutouts\n40°\n5.00\n6.44\nC-1411 (NACA 0015)\n55°\nExternal airfoil flap\nWing-fuselage intersection at $\\Lambda=0^\\circ$\nWing-fuselage intersection at $\\Lambda=15^\\circ$\nWing-fuselage intersection at $\\Lambda=30^\\circ$\nWing-fuselage intersection at $\\Lambda=45^\\circ$\n$\\Lambda=0^\\circ$, wing station 17.40\n$\\Lambda=0^\\circ$, wing station 6.50\n\nPortion of wing panel removed by faired and unfaired cutout.\nPortion of wing panel removed by large unfaired cutout.\n\nSection A-A, Faired cutout\nSection A-A, Unfaired cutout\nSection A-A, Large unfaired cutout\n\nSharp-leading-edge section\n1.50\nVane\n2.00\n.500\n.453\nHinge\n.200\n.765\n$\\delta_f=40^\\circ$\nExternal airfoil flap\nNACA\n\n2 1 0 1\nScale, inches\n\nFigure 4.- Details of wing modifications.\n\n15\n```", "timestamp": "2026-07-22T05:42:19.869569+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 27, "total_pages": 51, "image_filename": "19930085962_p27.jpg", "text": "26\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (206, 110, 733, 934) -->\n\nFigure 9. — Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:42:34.338955+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 7, "total_pages": 28, "image_filename": "19930086092_p7.jpg", "text": "```markdown\nNACA RM A9F14 CONFIDENTIAL 5\n\nThe vertical tail used had the same plan form¹ as a half-span of the wing: the leading edge swept back 63°, an aspect ratio of 1.75, a taper ratio of 0.25, and an NACA 64A006 section in a streamwise direction. Certain geometric characteristics of the tail pertinent to computations are given in the following tabulation:\n\n| Item | Symbol | Value |\n| :--- | :--- | :--- |\n| Relative tail area | $S_t/S$ | 0.168 |\n| Relative tail length | $l_t/\\bar{c}$ | 2.41 |\n| Relative tail volume | $S_t \\times l_t / S \\times \\bar{c}$ | 0.404 |\n\nThe vertical tail was equipped with a rudder having a ratio of flap chord to total chord of 0.25. The rudder had a minimum of balance, the nose of the rudder being an arc with a radius equal to half the airfoil thickness at the hinge line. The gap between rudder and fin was 1 percent of the local chord perpendicular to the leading edge and was unsealed. The rudder was equipped with a strain gage to provide hinge-moment data.\n\nPhotographs of the model mounted in the wind tunnel are shown in figure 2. The dimensions of the model are shown in figure 3. Based on a wing loading of 50 pounds per square foot and a design weight of 40,000 pounds, the model tested in the Ames 40- by 80-foot wind tunnel was about half scale.\n\nTESTS\n\nSix-component force data were obtained for the model with tail off, with tail on and rudder undeflected, and with tail on and rudder deflected to various set angles. Rudder hinge-moment data were also obtained. The data were obtained by varying the angle of sideslip while maintaining a constant angle of attack. All data were obtained at a tunnel speed of approximately 100 miles per hour, corresponding to a Reynolds number of $8.0 \\times 10^6$ based on the mean-aerodynamic-chord length of 8.639 feet.\n\nStandard tunnel-wall corrections for a straight wing of the same area and span as the swept-back wing have been applied to angle-of-attack and drag-coefficient data. This procedure was followed, since a brief analysis indicated that tunnel-wall corrections were approximately the same for straight and swept wings of the size under consideration. The corrections applied are as follows:\n\n¹The tail area and span used herein were measured to the center line of the fuselage.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:42:36.653032+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 15, "total_pages": 33, "image_filename": "19930085977_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:42:39.412227+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 18, "total_pages": 42, "image_filename": "19930085938_p18.jpg", "text": "NACA RM No. L9B04\n\n[Figure: Profile view of a model aircraft with single boom, mounted on a float-like structure, against a dark background. A scale ruler is visible near the front of the float. In the lower right corner of the image, a label reads “NACA L-56520.1”]\n\n(a) Profile view.\n\nFigure 2.— Model with single boom. Langley tank model 237-7B.\n\n17", "timestamp": "2026-07-22T05:42:40.484736+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 58, "total_pages": 58, "image_filename": "19930082617_p58.jpg", "text": "NACA TN 1962\n57\n\n[Figure: Side view of a large cylindrical test specimen labeled \"CYL 79\" showing buckling deformation after testing.]\n\nFigure 38. - Side view of cylinder 79 after buckling.\n\nNACA-Langley - 10-20-49 - 850", "timestamp": "2026-07-22T05:42:43.958107+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 30, "total_pages": 36, "image_filename": "19930085912_p30.jpg", "text": "28\n\nModel mass flow, lb/sec\n35\n30\n25\n20\n15\n10\n\nTunnel\nvelocity\n(ft/sec)\nO 217\n□ 290\n◇ 370\n△ 470\n\n0 10 20 30 40 50 60 70 80 90\nRise in model-air total temperature, °F\n\nNACA\n\nFigure 9. - Variation of mass air flow through model with rise in model-air total temperature\nfor various tunnel velocities. Angle of attack, 0°; free-stream total temperature, 0° F.\n\nNACA RM No. E9C16", "timestamp": "2026-07-22T05:42:47.191203+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 1, "total_pages": 39, "image_filename": "19930093769_p1.jpg", "text": "RM E8L10a\nCONFIDENTIAL\nCOPY NO.\nR.M. No. E8L10a\n\nNACA RM No. E8L10a\n\nNACA\nRESEARCH MEMORANDUM\n\nCOMPARISON OF PERFORMANCE OF AN-F-58 FUEL AND GASOLINE\nIN J34-WE-22 TURBOJET ENGINE\nBy Harry W. Dowman and George G. Younger\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\nTECHNICAL LIBRARY\nAIRESEARCH MANUFACTURING CO.\n9851-9951 SEPULVEDA BLVD.\nCLASSIFIED DOCUMENT\n\n[Stamp: CANCELLED]\n[Stamp: Classification changed by authority of NACA Office of the Assistant Director for Research 2-16-52]\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\nWASHINGTON\nApril 7, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:42:48.935049+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 56, "total_pages": 78, "image_filename": "19930082618_p56.jpg", "text": "```markdown\n54\n\nSection lift coefficient, $c_l$\nMoment coefficient, $c_{m_{c/4}}$\nSection angle of attack, $\\alpha_{abs}$, deg\n\nR\n0 $0.7 \\times 10^6$\n$\\circ$ 1.0\n$\\diamond$ 1.5\n$\\triangle$ 2.0\n$\\square$ 3.0\n$\\nabla$ 6.0\n$\\blacktriangle$ 9.0\n\nFlagged symbols denote standard roughness\n\nNACA\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\nFigure 12.- Aerodynamic characteristics of the NACA 4412 airfoil section, 24-inch chord.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:42:54.559735+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 28, "total_pages": 42, "image_filename": "19930085914_p28.jpg", "text": "C\nNACA RM A59E5\n\n20\nFlagged symbols denote\nrerun.\n16\nLift-drag ratio, $L/D$\n12\n8\n4\n0\n0 .2 .4 .6 .8 1.0 for M=.20\nLift coefficient, $C_L$\nM=.20 M=.40 M=.60 M=.70 M=.80 M=.85 M=.89 M=.92 M=.93\nNACA\n\nFigure 8.- The variation of lift-drag ratio with lift coefficient of the wing-fuselage\ncombination for several Mach numbers at a Reynolds number of 800,000.\n\n27", "timestamp": "2026-07-22T05:42:54.790938+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 18, "total_pages": 46, "image_filename": "19930085972_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:42:58.455576+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 1, "total_pages": 29, "image_filename": "19930093789_p1.jpg", "text": "RM-E8I21\nCONFIDENTIAL\nCOPY NO. 112\nRM No. E8I21\n\nNACA RM NO. E8I21\n\nNACA\nRESEARCH MEMORANDUM\n\nSOME EFFECTS OF STATOR CONE ANGLE AND BLADE-TIP LEAKAGE\nON 40-PERCENT-REACTION TURBINE HAVING ROTOR-BLADE CAPS\n\nBy Robert E. English, Robert J. McCready\nand John S. McCarthy\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nTECHNICAL LIBRARY\nAIRESEARCH MANUFACTURING CO.\n9001-9951 SEPULVEDA BLVD.\nINGLEWOOD,\nCALIFORNIA\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nMarch 23, 1949\n\nCONFIDENTIAL\n\nClassification\nCANCELLED\nCHANGED TO Unclassified\nBy authority of Tech Abstract, Jan 29, 1954\nDate Pub 4, 1954\nChanged by GN", "timestamp": "2026-07-22T05:42:59.735999+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 1, "total_pages": 34, "image_filename": "19930082472_p1.jpg", "text": "TN-1797\n\nNACA TN No. 1797\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\nNo. 1797\n\nA STUDY OF STALL PHENOMENA ON A $45^\\circ$\nSWEPT-FORWARD WING\n\nBy Gerald M. McCormack and Woodrow L. Cook\n\nAmes Aeronautical Laboratory,\nMoffett Field, Calif.\n\nTECHNICAL LIBRARY\nAIRESEARCH MANUFACTURING CO.\n9851-9951 SE. ULVEDA BLVD.\nLOS ANGELES 45, CALIF.\nCALIFORNIA\n\n[Figure: NACA logo]\n\nWashington\nJanuary 1949", "timestamp": "2026-07-22T05:43:03.542892+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 11, "total_pages": 50, "image_filename": "19930086076_p11.jpg", "text": "NACA RM E9F09\n\n290° F, combustion efficiency was 80 percent. These data were taken near the end of 5 minutes of operation. The condition of this flame holder after 5 minutes running time is shown in figure 13. The ratio of the total-pressure loss through the combustor to the inlet dynamic pressure was 2.\n\nUse of molybdenum. - The flame-holder designs that gave improved combustion efficiencies reached temperatures high enough to rapidly melt Inconel, thus preventing accurate determination of combustor performance. In an attempt to find a more heat-resistant material, one of the downstream gutters of flame holder 7 was replaced by a silicon-coated molybdenum prism to compose flame holder 11. In addition, the two rows of gutters were made to converge on each other slightly in order to utilize more of the central air stream for combustion (fig. 14). Figure 15 shows the molybdenum prism intact amid the remains of several melted Inconel gutters at the end of the 5-minute run. At an inlet-air pressure of 55 inches of mercury absolute and with the use of liquid AN-F-48b fuel, the combustion efficiency was 45 percent.\n\nIn order to substantiate further the applicability of silicon-coated molybdenum to flame-holder components, flame holder 12 was fabricated with the same dimensions as flame holder 11 except that the last five gutters in each row were replaced by silicon-coated molybdenum prisms. The flame holder was operated for 47 minutes at the standard operating conditions with only slight oxidation of the molybdenum in local spots where the coating apparently failed (fig. 16). The fuel used with this flame holder was liquid AN-F-48b.\n\nFlame holder 13 duplicated flame holder 9 except that silicon-coated molybdenum plates were used instead of Inconel plates (fig. 7). Only 1/8-inch Inconel was available for the large retaining shell required to hold the plates. The shell rapidly melted during the investigation although the plates showed no deterioration. In this run, liquid isopentane was sprayed through the hollow-cone spray-nozzle fuel injector. Figure 18 shows the condition of the flame holder after 5 minutes of operation. The rough surface on some of the molybdenum plates is Inconel that melted and deposited there during operation. At the standard operating condition, combustion efficiency was 65 percent. The ratio of the total-pressure loss through the combustor to the inlet dynamic pressure was 3.\n\nMagnitude of pressure fluctuations. - Pressure fluctuations in the combustor were recorded during the investigation of flame holder 13 and were compared to fluctuations recorded during isothermal flow (fig. 19). Combustor pressure varied less than", "timestamp": "2026-07-22T05:43:05.527233+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 6, "total_pages": 36, "image_filename": "19930086097_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM A9H11\n\n$c_{d_{min}}$ minimum section drag coefficient (profile drag at zero lift)\n\n$c_{d_f}$ section friction drag coefficient\n\n$c_{d_w}$ pressure drag coefficient at zero lift for section forward of base (wave drag)\n\n$c_l$ section lift coefficient\n\n$c_m$ section pitching-moment coefficient taken about midchord position\n\n$(\\Delta c_{d_{min}})_t$ increment in profile drag for a given thickness ratio\n\n$(\\Delta c_{d_{min}})_s$ increment in profile drag for a given section modulus\n\n$C_D$ wing drag coefficient\n\n$C_{D_{min}}$ minimum wing drag coefficient\n\n$C_L$ wing lift coefficient\n\n$C_1, C_2$ constants appearing in Busemann second-order airfoil theory\n\n$h$ trailing-edge thickness\n\n$(L/D)_{max}$ maximum lift-drag ratio\n\n$M$ Mach number\n\n$p$ local static pressure\n\n$P$ pressure coefficient $\\left( \\frac{p - p_\\infty}{\\frac{1}{2} \\rho_\\infty U_\\infty^2} \\right)$\n\n$P_{b_v}$ base pressure coefficient for vacuum $\\left( \\frac{-2}{\\gamma M_\\infty^2} \\right)$\n\n$R$ Reynolds number\n\n$t$ maximum thickness of airfoil\n\n$U$ velocity\n\n$x$ airfoil abscissa\n\n$y$ airfoil ordinate\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:43:06.916320+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 16, "total_pages": 33, "image_filename": "19930085977_p16.jpg", "text": "NACA RM L9H22 CONFIDENTIAL 15\n\n[Figure: A pictorial view of a wing with four vertical fins mounted on its surface. The wing has a sweptback design, and the fins are positioned along the span. One fin is larger and located near the root, while three smaller fins are spaced further out toward the tip. The image includes a label “NACA L-61940” in the lower right corner.]\n\nFigure 4.— A pictorial view of a $0^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section showing free-floating tails.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:43:07.869544+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 19, "total_pages": 42, "image_filename": "19930085938_p19.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:43:08.751921+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 12, "total_pages": 98, "image_filename": "19930086073_p12.jpg", "text": "10\nNACA RM A9H04\n\nThe following values were thus substituted in the two equations:\n\n| | |\n| :--- | :--- |\n| $C_{N_{\\alpha_t}}$ | 0.027 |\n| $\\frac{d\\alpha_t}{d\\delta_r}$ | .62 |\n| F | .845 |\n| $S_t/S$ | .218 |\n| $l/b$ (equation (1)) | .525 |\n| $l/b$ (equation (2)) | .741 |\n| $\\frac{d\\sigma}{d\\beta}$ | 0 |\n| $q_t/q$ | 1.00 |\n\nThe computed and experimental values compare as follows:\n\n| | $C_{n_{\\beta_t}}$ | $dC_n/d\\delta_r$ |\n| :--- | :--- | :--- |\n| Computed | 0.0031 | -0.0023 |\n| Experimental | .0032 | -.0025 |\n\nThe agreement between the computed and experimental values is thus satisfactory for zero angle of attack and would probably remain satisfactory until the angle of attack is reached at which the flow due to the separation vortices begins to have a strong influence upon the tail characteristics.\n\nCONCLUDING REMARKS\n\nThe results of this investigation show that the body combined with the triangular plan-form wing caused no changes in the lift characteristics of the wing and caused only a 1-percent decrease in the static margin. Flap lift and pitching-moment effectiveness decreased proportional to the decrease in flap area caused by the addition of the body. The wing with body and vertical tail exhibited positive dihedral effect", "timestamp": "2026-07-22T05:43:13.929463+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 1, "total_pages": 21, "image_filename": "19930082474_p1.jpg", "text": "NACA TN No. 1799\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\nNo. 1799\n\nON THE FLYING QUALITIES OF HELICOPTERS\nBy John P. Reeder and F. B. Gustafson\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\n[Figure: NACA logo with wings]\n\nWashington\nJanuary 1949", "timestamp": "2026-07-22T05:43:15.973623+00:00"}