PDF/ocr_dataset_1035.jsonl

{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 55, "total_pages": 99, "image_filename": "19930082511_p55.jpg", "text": "NACA TN No. 1826\n53\n\nAn extrapolation from the present results indicates that the induced upflow at the lifting element, for the lifting element in the plane of the entrance lip, is approximately zero, or the average between the completely open and completely closed cases. The same result (that the effect in the plane of the entrance lip is the average of the effects for the completely open and the completely closed tunnels) was also obtained for the two-dimensional tunnel (fig. 18).\n\nCONCLUSION\n\nFor an open wind tunnel, the corrections corresponding to an infinitely long open jet will usually be adequately accurate if the region of interest (where the lift is located and where the boundary-induced flow is being considered) is at least half the jet height from the jet entrance and exit sections. As the distance of the lifting element from the entrance is decreased below this limit, the boundary-induced flow decreases rapidly and, when the lifting element is in the entrance plane, the induced angle at the lifting element is about the average of that for an open tunnel and that for a closed tunnel.\n\nIn the theoretical studies of these flows, the usual boundary conditions of pressure uniformity on the free surface and of zero normal velocity on the closed surface must be supplemented with the conditions that the velocity be continuous at the entrance lip and that the velocities far downstream and far upstream in the closed sections be equal. For the two-dimensional open tunnel, a convenient general mathematical approach is to transform the infinite strip (representing the tunnel) to the upper half-plane by the logarithmic transformation and then to develop the desired complex velocity in this transformed plane. For the circular open tunnel the solution may be effected by expressing the potential by a finite series of Bessel functions, satisfying the boundary condition on the free surface at a finite number of points, and solving for the coefficients by simultaneous linear equations.\n\nFor noncircular open wind tunnels, solutions in terms of available functions will be very inconvenient. For such cases, the trends indicated by the present results may suffice, when applied to the presumably known corrections for the infinitely long open and closed configurations, to provide adequate corrections. Solutions for the general three-dimensional configuration may also be possible by electrical-analogy methods, in which either the perturbation velocity potential or the acceleration potential is analogous to the electrical potential in an electrolyte solution. Such analogies may be characterized, however, by considerable technical difficulty.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va., December 20, 1948", "timestamp": "2026-07-22T05:21:59.026180+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 38, "total_pages": 46, "image_filename": "19930085519_p38.jpg", "text": "```markdown\nNACA RM No. L8K19\n37\n\n$\\delta_p$\n(percent\nlocal wing\nchord)\n$\\circ$ - 1/2\n$\\triangle$ - 1\n$\\square$ - 2\n$\\diamond$ - 3\n$\\blacktriangle$ - 5\n$\\blacktriangledown$ - 7\n\nRolling-moment coefficient, $C_l$\nYawing-moment coefficient, $C_n$\n\n| Angle of attack, $\\alpha$, deg | -12 | -8 | -4 | 0 | 4 | 8 | 12 | 16 | 20 | 24 | 28 |\n| :--- | :---: | :---: | :---: | :---: | :---: | :---: | :---: | :---: | :---: | :---: | :---: |\n| $C_l$ | | | | | | | | | | | |\n| $C_n$ | | | | | | | | | | | |\n\n[Figure: Graph showing variation of rolling-moment coefficient and yawing-moment coefficient with angle of attack for various projections of the plug aileron on the 42° sweptback wing. Flap retracted. NACA logo present.]\n\n(a) Sharp plug-slot lower lip.\n\nFigure 17.- Variation of rolling-moment coefficient and yawing-moment coefficient with angle of attack for various projections of the plug aileron on the 42° sweptback wing. Flap retracted.\n```", "timestamp": "2026-07-22T05:22:02.024196+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 4, "total_pages": 42, "image_filename": "19930085914_p4.jpg", "text": "NACA RM A9D25\n\n$C_L$ lift coefficient $\\left(\\frac{\\text{lift}}{qS}\\right)$\n\n$C_m$ pitching-moment coefficient about the quarter-chord point of the wing M.A.C. $\\left(\\frac{\\text{pitching moment}}{qSc}\\right)$\n\n$\\frac{dC_L}{d\\alpha}$ lift-curve slope, per degree\n\nM Mach number $\\left(\\frac{V}{a}\\right)$\n\nq dynamic pressure $\\left(\\frac{1}{2}\\rho V^2\\right)$, pounds per square foot\n\nR Reynolds number $\\left(\\frac{\\rho Vc}{\\mu}\\right)$\n\nS wing area, square feet\n\nt maximum thickness of wing section, feet\n\nV free-stream velocity, feet per second\n\ny lateral distance, feet\n\n$\\alpha$ angle of attack of root chord line, degrees\n\n$\\alpha_t$ angle of twist with reference to root chord (positive for washin), degrees\n\n$\\mu$ coefficient of viscosity of air, slugs per foot-second\n\n$\\rho$ mass density of air, slugs per cubic foot\n\nMODEL AND APPARATUS\n\nPhotographs of the model used in this investigation are presented in figure 1, and dimensions are given in figures 2 and 3. The wing was constructed of solid steel, and the fuselage of steel and aluminum.\n\nThe wing had a leading-edge sweepback of $63^\\circ$, a tip-chord-to-root-chord ratio of 0.25, and an aspect ratio of 3.5. The streamwise airfoil sections had the NACA 64A005 thickness distribution combined with a=1 mean-camber lines. The wing, as developed theoretically for a lift", "timestamp": "2026-07-22T05:22:02.202800+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 44, "total_pages": 66, "image_filename": "19930082914_p44.jpg", "text": "NACA TN No. 1857\n43\n\n10 Megohms\nTo 16000 V-d-c\n2 Microfarads\n6V\nSwitch\n1000\nresistor\nMagnesium\nelectrodes\n\nLight source\nEntrance slit\nL3\nL4\nL5\nExit slit\nPrism\nL1\n\nInterferometer\nM2\nTest\nSection\nS2\nL2\nC1\nS1\nC2\nM1\n\nCamera\nShutter\nFilm\n\nFigure 8.- Complete optical diagram of interferometer used in this investigation. L1, parabolic mirror f/9.0, 36-inch focal length; L2, Eastman lens f/3.5, 13.5-inch focal length; L3, Eastman lens f/1.6, 50-millimeter focal length; L4, Bausch and Lomb f/2.3, 2-inch focal length; L5, Bausch and Lomb f/8.0, 6-inch focal length; C1 and C2, compensating plates.", "timestamp": "2026-07-22T05:22:08.426829+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 17, "total_pages": 96, "image_filename": "19930085880_p17.jpg", "text": "NACA RM No. L9C03\n15\n\n<!-- Image (162, 93, 882, 446) -->\n\n(a) $\\tau = 4^\\circ$.\n\n<!-- Image (162, 488, 882, 841) -->\n\n(b) $\\tau = 8^\\circ$.\n\nFigure 6.- Variation of wetted lengths at chine and at model centerline\nwith wetted area for model 250B.", "timestamp": "2026-07-22T05:22:10.168590+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 26, "total_pages": 33, "image_filename": "19930085544_p26.jpg", "text": "NACA RM No. L8K26\n25\n\nCalculated\nExperimental (reference 3)\n\n| $\\frac{dC_T}{dx}$ | $\\omega t$ (deg) | $C_T$ |\n| :--- | :--- | :--- |\n| .08 | 120 | 0.280 |\n| .06 | 90 | 0.240 |\n| .04 | 180 | 0.110 |\n| .02 | 270 | -.0014 |\n| 0 | 300 | -.0018 |\n| -.02 | | |\n| -.04 | | |\n| -.06 | | |\n\nFuselage surface for wake surveys\n(survey rake located 1/8\" behind propeller)\n\n[Figure: Graph showing curves for Calculated and Experimental data. The x-axis ranges from 0 to 10. The y-axis ranges from -.06 to .08. There are multiple curves plotted with labels 120, 90, 180, 270, 300 corresponding to $\\omega t$ (deg) and 0.280, 0.240, 0.110, -.0014, -.0018 corresponding to $C_T$.]\n\nFigure 8.- Thrust gradient curves for three blade positions. B.2; $\\beta_{.26}^\\circ$; J.1.2; $\\alpha_T, 14^\\circ$.", "timestamp": "2026-07-22T05:22:10.476788+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 21, "total_pages": 36, "image_filename": "19930085869_p21.jpg", "text": "NACA RM L9D15\n\nCONFIDENTIAL\n\n[Figure: Test setup showing a large aircraft model mounted on a support structure inside a wind tunnel or test facility. The model is positioned horizontally, with visible internal framework and instrumentation. A person is seated in the cockpit area of the model. Various cables, sensors, and structural supports are visible around the model. The background shows the interior of a large industrial testing chamber.]\n\nFigure 5.—Test setup.\n\nCONFIDENTIAL\n\n19\n\nNACA\nI-60571", "timestamp": "2026-07-22T05:22:14.666398+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 6, "total_pages": 23, "image_filename": "19930085906_p6.jpg", "text": "NACA RM ESF20 CONFIDENTIAL 5\n\nthe curve near the origin indicates the presence of some higher order exponential. Other preheater designs might be expected to follow the same general variation but with a different time constant.\n\nThe data in figure 6 indicate that with given operating conditions an appreciable rise in the final fuel temperature can be achieved by increasing the distance between the flame holder and the preheater from 0 to 8 inches. A further increase from 8 to 12 inches results in little or no gain in the final fuel temperature. A similar result is indicated in figure 8, in which the time to reach a fuel temperature of $200^\\circ$ F from an initial fuel temperature of $63^\\circ \\pm 7^\\circ$ is presented as a function of preheater position. (From the results obtained in reference 1, $200^\\circ$ F may be considered a useful preheat temperature.) The data of figure 8 indicate a sharp decrease in the time required to reach $200^\\circ$ F as the distance between the flame holder and the preheater is increased from 0 to 8 inches. The mean time was reduced from 2.4 minutes at zero distance to 1.2 minutes at 8 inches between the flame holder and the preheater. A further increase in this distance from 8 to 12 inches results in an additional reduction in the time of only 0.1 minute. Thus, for all operating conditions, useful preheat temperatures were approached within the first 2 minutes of ignition and the temperature rise attained 99 percent of its final value within 5 minutes. Also from the trends obtained, it is probable that little improvement can be expected from further increase in the distance between the flame holder and the preheater. Further improvement can probably be obtained, however, if the fuel-preheating path is increased by adding more coils to the preheater.\n\nThe general range of final fuel temperatures obtained at each preheater position is indicated in figure 9. In addition to being affected by the preheater position, the final fuel temperature is influenced by the engine operating condition. Variation in ram-jet operating conditions resulted in a maximum spread of $\\pm 15$ percent in the fuel-temperature rise. At the same engine conditions, the final fuel temperature was approximately the same for the preheater in either the 8- or 12-inch position. The difference in spread of the data in figure 9 for these positions is due to differences in ram-jet operating conditions for the data presented.\n\nThe heat absorbed per pound of fuel is presented in figure 10 as a function of preheater position. Because the heat absorbed is a function of the final fuel temperature, the values of figure 9 and 10 exhibit the same trend. The amount of heat absorbed per pound of fuel increases sharply as the distance between the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:22:15.421953+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 14, "total_pages": 37, "image_filename": "19930085889_p14.jpg", "text": "NACA RM L9F14 CONFIDENTIAL 13\n\nAileron Characteristics\n\nThe effects of sweepback on the aileron rolling-moment-effectiveness parameter $C_{l_\\delta}$ and on the rolling-effectiveness parameter $(pb/2V)_\\delta$ are shown in figure 18. The values of $C_{l_\\delta}$ were determined in straight flow and the values of $(pb/2V)_\\delta$ were determined from the relation\n\n$$\n\\left(\\frac{pb}{2V}\\right)_\\delta = \\frac{C_{l_\\delta}}{C_{l_p}}\n$$\n\nwhere $C_{l_p}$ is obtained from figure 13 and represents the damping of the wing with ailerons neutral. The values of $(pb/2V)_\\delta$ presented in figure 18, therefore, neglect any possible effect of aileron deflection on the damping in roll or of rolling on aileron effectiveness. Previous experience has indicated, however, that such effects are negligible, except for very large aileron deflections or for angles of attack near the maximum lift coefficient.\n\nResults obtained for the parameters $C_{l_\\delta}$ and $(pb/2V)_\\delta$ depend, of course, on the particular convention used in defining the aileron deflection $\\delta$. In the present paper, $\\delta$ is measured in a plane perpendicular to the aileron hinge axis and, therefore, a given value of $\\delta$ represents a constant angular rotation of the aileron about its hinge axis regardless of the sweep angle of the wing. With this convention, an increase in sweep angle is found to produce large reductions in both $C_{l_\\delta}$ and $(pb/2V)_\\delta$. (See fig. 18.) According to an alternate convention, the deflection $\\delta$ is measured in the plane of symmetry and, therefore, a constant value of $\\delta$ corresponds to an increasing angular rotation of the aileron about the hinge line as the wing sweep angle is increased. When the latter convention is used, the effect of sweepback on the parameters $C_{l_\\delta}$ and $(pb/2V)_\\delta$ is found to be considerably smaller than that indicated in figure 18.\n\nA comparison of figures 18 and 13 shows that for the three wings investigated the variation of $C_{l_\\delta}$ with lift coefficient is small relative to the variation of $C_{l_p}$ with lift coefficient. The resulting variation of the rolling-effectiveness parameter $(pb/2V)_\\delta$,\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:22:28.127903+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 12, "total_pages": 31, "image_filename": "19930085881_p12.jpg", "text": "CONFIDENTIAL\n10\n\nRocket motor\n5.00D\nAileron hinge line\n0.71\n7.07\n1.41\n1.00\n13.11\n56.00\n10.00\nAileron hinge line\n\n0.71\n7.07\n0.020 steel plates\n1.41\nSection A-A (typical)\n\n0.71\n7.07\n1.41\nAileron hinge line\nA\nA\n1.00\n13.11\n2.00\n\nDimensions are in inches\nCenter of gravity at station 35.0\n\nFuselage Ordinates\n| Station | Diameter |\n| :--- | :--- |\n| 0 | 0 |\n| 2.50 | 1.22 |\n| 5.00 | 2.30 |\n| 7.50 | 3.16 |\n| 10.00 | 3.92 |\n| 12.50 | 4.52 |\n| 15.00 | 4.88 |\n| 17.50 | 5.00 |\n\nNACA\n\n(a) Geometric details.\nFigure 1.- General arrangement of test vehicles.\nCONFIDENTIAL\n\nNACA RM L9D12", "timestamp": "2026-07-22T05:22:28.444274+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 1, "total_pages": 42, "image_filename": "19930085938_p1.jpg", "text": "NACA RM No. L9B04\nRM No. L9B04\n\n[Figure: NACA logo with wings]\n\nRESEARCH MEMORANDUM\n\nHYDRODYNAMIC CHARACTERISTICS OF AERODYNAMICALLY\nREFINED PLANING-TAIL HULLS\n\nBy\nRobert McKann and Claude W. Coffee\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nMarch 28, 1949", "timestamp": "2026-07-22T05:22:29.086668+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 33, "total_pages": 53, "image_filename": "19930082542_p33.jpg", "text": "NACA TN No. 1867\n\n33\n\naged 24 hr\nno age\n1350° F\n1500° F\n1750° F\n\nHot-rolled\nhot-cold-rolled\n80° F - 10 percent\n1200° F\n5 percent\n10 percent\n15 percent; aged 24 hr at 1400° F\n20 percent; reduced 85 percent from 1800° to 1400° F\n\n2 hr at 1800° F, cool to 1400° F, rolled 20 percent, reheated to 1800° F for 2 hr, and then repeated six times\n\n1800° F; 2 hr W.Q.\n2050° F; 2 hr W.Q.\n2100° F\n2150° F\n\n1 hr W.Q.\n2 hr W.Q.\n5 hr W.Q.\n2 hr A.C.\n\naged\n1350° F; 24 hr\n1500° F; 24 hr\n1600° F; 24 hr\n1750° F; 24 hr\n\n2 hr\n8 hr\n24 hr\n50 hr\n\nrolled 15 percent at 1200° F\n\nrolled 10 percent at 1200° F\n\n5 percent at 1200° F\n10 percent at\n15 percent at\n20 percent at\n25 percent at\n\n80° F\n1200° F\n\n1000° F\n1200° F\n1400° F\n1600° F\n1700° F\n1800° F\n\naged 24 hr at 1400° F\n\nSolution-treated\n2200° F\n2250° F\n1/2 hr W.Q.\n1 hr W.Q.\n1 hr A.C.\n\n1350° F\n1400° F\n1500° F\n1600° F\n1750° F\n\naged\n24 hr\n50 hr\n\n24 hr\n24 hr\n24 hr\n\nrolled 10 percent at 1200° F\n\n5 percent at 1200° F\n10 percent at 1200° F\n15 percent at\n\n1000° F\n1200° F\n1400° F\n1600° F\n\naged 24 hr at 1400° F\n\n[Figure: NACA logo]\n\nFigure 2. - Diagram showing treatments used on the low-carbon N-155 bar stock.", "timestamp": "2026-07-22T05:22:30.115287+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 36, "total_pages": 58, "image_filename": "19930082617_p36.jpg", "text": "NACA TN 1962\n35\n\nStringers\n$\\circ$ 1 to 9\n$\\times$ 10 to 16\n\nMoment\n(in. - lb)\n1 $36.0 \\times 10^3$\n2 $72.0 \\times 10^3$\n3 $108.0 \\times 10^3$\n4 $144.0 \\times 10^3$\n5 $180.0 \\times 10^3$\n\n[Figure: Cross-section diagram showing dimensions 2.57\", Band B, arrows labeled A, and a circular cross-section A-A with 45° angles]\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n10 8 6 4 2 0 -2 -4 -6 -8 -10 $\\times 10^{-4}$\nStrain\n\n[Figure: NACA logo]\n\n1 2 3 4 5\n\nFigure 24.- Strain diagram of cylinder 79. Band B.", "timestamp": "2026-07-22T05:22:32.465846+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 20, "total_pages": 29, "image_filename": "19930085879_p20.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:22:35.601509+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 40, "total_pages": 78, "image_filename": "19930082618_p40.jpg", "text": "```markdown\n38\n\nSection drag coefficient, $c_d$\nSection lift coefficient, $c_l$\nR\n$\\circ$ $0.7 \\times 10^6$\n$\\square$ $1.0$\n$\\diamond$ $1.5$\n$\\triangle$ $2.0$\nFlagged symbols denote standard roughness\n\nSection drag coefficient, $c_d$\nSection lift coefficient, $c_l$\nR\n$\\nabla$ $3.0 \\times 10^6$\n$\\square$ $6.0$\n$\\triangleleft$ $9.0$\nFlagged symbols denote standard roughness\n\nNACA\n\nMoment coefficient, $c_m$\nSection lift coefficient, $c_l$\nR\n$\\circ$ $0.7 \\times 10^6$\n$\\square$ $1.0$\n$\\diamond$ $1.5$\n$\\triangle$ $2.0$\n$\\nabla$ $3.0$\n$\\square$ $6.0$\n$\\triangleleft$ $9.0$\n\na.c. position\nx/c y/c\n.248 -.099\n.255 -.048\n.257 -.050\n.259 -.082\n.259 -.048\n.253 -.030\n.252 -.021\n\n(b) Section drag characteristics and section pitching-moment characteristics about the aerodynamic center of the plain NACA $64_1A212$ airfoil section.\n\nFigure 6.— Concluded.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:22:38.836919+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 56, "total_pages": 99, "image_filename": "19930082511_p56.jpg", "text": "54\nNACA TN No. 1826\n\nAPPENDIX A\n\nEVALUATION OF $P_{mn}(\\xi, \\rho)$\n\nEvaluation for $m \\neq 0$\n\nContour integration.- For $n \\neq 0$,\n\n$$P_{mn}(\\xi, \\rho) = \\frac{1}{\\pi} \\int_0^\\infty \\frac{J_m(i\\rho q)}{iq J_m'(iq)} dq \\int_a^b \\sin n\\pi \\frac{\\beta - a}{b - a} \\cos q(\\beta - \\xi) d\\beta \\quad \\text{(A1)}$$\n\nThe inner integral may be found directly:\n\n$$\\int_a^b \\sin n\\pi \\frac{\\beta - a}{b - a} \\cos q(\\beta - \\xi) d\\beta$$\n\n$$= \\frac{1}{2} \\int_a^b \\left\\{ \\sin \\left[ n\\pi \\frac{\\beta - a}{b - a} + q(\\beta - \\xi) \\right] + \\sin \\left[ n\\pi \\frac{\\beta - a}{b - a} - q(\\beta - \\xi) \\right] \\right\\} d\\beta$$\n\n$$= \\frac{\\frac{n\\pi}{b - a}}{q^2 - \\left( \\frac{n\\pi}{b - a} \\right)^2} \\left[ (-1)^n \\cos q(b - \\xi) - \\cos q(\\xi - a) \\right] \\quad \\text{(A2)}$$\n\nThe problem of evaluating $P_{mn}(\\xi, \\rho)$ thus reduces to that of evaluating integrals of the form\n\n$$\\int_0^\\infty \\frac{J_m(i\\rho q) \\cos kq \\, dq}{iq J_m'(iq) (q^2 - h^2)}$$\n\nConsider the integral in the complex $z$-plane\n\n$$\\frac{1}{2\\pi i} \\oint \\frac{J_m(i\\rho z) e^{ikz}}{iz J_m'(iz) (z^2 - h^2)} dz$$", "timestamp": "2026-07-22T05:22:44.580143+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 39, "total_pages": 46, "image_filename": "19930085519_p39.jpg", "text": "38\nNACA RM No. L8K19\n\n$\\delta p$\n(percent\nlocal wing\nchord)\n$\\circ$ -1/2\n$\\triangle$ -1\n$\\square$ -2\n$\\diamond$ -3\n$\\triangleleft$ -5\n$\\triangleright$ -7\n\nRolling-moment coefficient, $C_l$\nYawing-moment coefficient, $C_n$\n\n<!-- Image (112, 355, 825, 780) -->\n\n(b) Faired plug-slot lower lip.\nFigure 17.- Concluded.", "timestamp": "2026-07-22T05:22:45.295274+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 6, "total_pages": 36, "image_filename": "19930085912_p6.jpg", "text": "4\nNACA RM No. E9C16\n\nL latent heat of vaporization of water, Btu per pound\n\nl depth of jet penetration into air stream at distance s\ndownstream of orifice center line, inches\n\nm liquid-water content, pounds water per pound air\n\nPf total pressure at front rakes, pounds per square\nfoot absolute\n\nPj jet total pressure, inches of mercury absolute\n\nP0 free-stream total pressure, pounds per square\nfoot absolute\n\np local surface static pressure, pounds per square foot\nabsolute\n\nP0 free-stream static pressure, pounds per square foot absolute\n\nΔp static-pressure drop across screen, pounds per square foot\n\nq dynamic pressure ahead of screen, pounds per square foot\n\nq0 free-stream dynamic pressure, pounds per square foot\n\nS pressure coefficient, $1 - \\left( \\frac{p-P_0}{q_0} \\right)$\n\ns distance downstream of orifice center line or mixing distance,\ninches\n\nTav model-air total temperature (area weighted), °F\n\nTg plenum-chamber gas temperature, °F\n\nTj total temperature of jet $(T_g + 460)$, °R\n\nTx local air total temperature in model, °F\n\nT0 free-stream total temperature, °F\n\nVi inlet-stream velocity at orifices, feet per second\n\nVj velocity of jet at vena contracta, feet per second\n\nVs velocity through screen, feet per second", "timestamp": "2026-07-22T05:22:45.699799+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 2, "total_pages": 23, "image_filename": "19930085934_p2.jpg", "text": "NACA RM E9G12\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nMETHOD OF DETERMINING CENTRIFUGAL-FLOW-COMPRESSOR PERFORMANCE\n\nWITH WATER INJECTION\n\nBy Joseph T. Hamrick and William L. Beede\n\nSUMMARY\n\nA method is presented for computing the isentropic and actual enthalpy change between the inlet and the outlet of a centrifugal-flow compressor when water injection is used. The change in state of the water in the working fluid complicates the use of specific heats in accounting for property changes in the fluid. Consequently, the method of calculation is based on fluid property values given in steam and air tables. Functional charts are included to show the effect of water-air ratio on isentropic outlet temperature, isentropic enthalpy rise, and adiabatic efficiency with various pressure ratios for a fixed set of inlet conditions.\n\nFor a given isentropic enthalpy change or adiabatic efficiency, the rate of increase in the compressor pressure ratio decreased as the water-air ratio was increased. For compressor-inlet conditions of pressure, 14 inches of mercury absolute; temperature, $77^\\circ$ F; and specific humidity, 0, water-air ratios greater than 0.05 were relatively ineffective for pressure ratios of less than 8. On the basis of the large difference in flow conditions between wet and dry compression, the design of a compressor that would have good efficiency for both wet and dry compression might be difficult.\n\nINTRODUCTION\n\nWater-methanol injection at the compressor inlet is a standard thrust-augmentation method for turbojet engines. The large increase in thrust obtainable with this method has resulted in investigations of the performance effects of water injection on the engine components to determine possible sources of losses. Difficulty in evaluating the properties of the vapor-air mixture complicates the analysis of the turbojet cycle and especially the computation of compressor efficiency.\n\nA method of computing the efficiency of a centrifugal compressor with water injection at the inlet was derived at the NACA Lewis", "timestamp": "2026-07-22T05:22:46.845637+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 27, "total_pages": 33, "image_filename": "19930085544_p27.jpg", "text": "26\nNACA RM No. L8K26\n\n<!-- Image (119, 136, 857, 935) -->\n\nFigure 9.-- Variation of force coefficient with blade position. $\\epsilon, 0.1; \\frac{\\alpha}{\\alpha_{20}}, 0$.", "timestamp": "2026-07-22T05:22:50.493639+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 22, "total_pages": 36, "image_filename": "19930085869_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:22:55.336944+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 45, "total_pages": 66, "image_filename": "19930082914_p45.jpg", "text": "44\nNACA TN No. 1857\n\n<!-- Image (95, 133, 874, 827) -->\n\no Fringe positions with no flow\n• Fringe positions with flow\n\nFigure 9.- Production of fringe shifts.", "timestamp": "2026-07-22T05:22:55.531288+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 18, "total_pages": 96, "image_filename": "19930085880_p18.jpg", "text": "16\nNACA RM No. L9C03\n\n<!-- Image (119, 99, 855, 437) -->\n\n(c) $\\tau = 12^\\circ$.\n\n<!-- Image (119, 486, 855, 824) -->\n\n(d) $\\tau = 16^\\circ$.\n\nFigure 6.- Continued.", "timestamp": "2026-07-22T05:22:56.394539+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 5, "total_pages": 42, "image_filename": "19930085914_p5.jpg", "text": "4\nNACA RM A9D25\n\ncoefficient of 0.25 at a Mach number of 1.5, was cambered and twisted to support a uniform distribution of lift over its surface. This development was described in reference 2. The model of reference 2 was constructed with less twist than was indicated by theory, the theoretical twist being reduced by the amount expected from wing deflection at the design lift coefficient and at the test dynamic pressure. Since the range of aerodynamic forces encountered in this series of tests was so wide, it was impossible to design the model to compensate for the effects of aerodynamic loading on wing twist. Consequently, the model was designed with the same twist variation under the no-load condition as the model of reference 2. Spanwise variation of camber and twist is shown in figure 3, and section coordinates are given in table I.\n\nThe fuselage shape used in this investigation has been determined by Haack (reference 3) to have minimum pressure drag at supersonic speeds for a given length and volume, assuming closure at the tail as indicated by the dashed lines in figure 2. The after 21 percent of the model fuselage length was cut off to permit installation on the sting support. The resultant fineness ratio of the fuselage was 9.9; whereas the fineness ratio of the basic closed body was 12.5. The equation defining the coordinates of the fuselage is given in figure 2.\n\nThe model was equipped with constant-chord plain flaps extending over the outer 50 percent of the span. The flap chord was 25 percent of the wing chord at midsemispan. The flap had a radius nose and the unsealed gap was approximately 3/64 inch. This large gap was necessary to permit the desired angular deflection since the flap had considerable spanwise curvature. For the tests reported herein, the flap was undeflected, and was restrained near its inner extremity.\n\nThe model was mounted on a sting-type support, and the angle of attack was continuously controllable from a remote station during wind-tunnel operation. All forces and moments were measured by means of a wire-resistance strain-gage balance enclosed by the model.\n\nTESTS\n\nLift, drag, and pitching-moment data have been obtained throughout an angle-of-attack range for the wing-fuselage combination. The angle-of-attack range for the tests was from -8° to +19°, except at high Mach numbers and the highest Reynolds numbers where the angle was limited by vibration of either the model or its support, or by wind-tunnel power. At Reynolds numbers of 0.8 million and 2.0 million, data were obtained over a range of Mach numbers up to a maximum of 0.93. At a Mach number of 0.20, data were obtained over a range of Reynolds numbers from 0.8 million to 9.0 million. Lift, drag, and pitching-moment data have been obtained for the fuselage alone throughout the same range of angle of", "timestamp": "2026-07-22T05:22:56.743795+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 13, "total_pages": 31, "image_filename": "19930085881_p13.jpg", "text": "NACA RM L9D12\n\nCONFIDENTIAL Relative size of angles exaggerated\n\nAileron hinge line\n\nWing camber line\n\n$i_w$\n\n$\\delta_a$\n\nAileron camber line\n\np\n\nPositive values are indicated on figure\n\n(b) Sign conventions.\n\nFigure 1.— Concluded.\n\nCONFIDENTIAL\n\n11", "timestamp": "2026-07-22T05:23:05.243885+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 7, "total_pages": 23, "image_filename": "19930085906_p7.jpg", "text": "6 CONFIDENTIAL NACA RM E5F20\n\nflame holder and the preheater is extended. From a mean of 80 Btu per pound of fuel, the amount of heat absorbed by the fuel can be increased to a mean of 120 Btu per pound of fuel, a rise of 50 percent, if the distance between flame holder and preheater is changed from 0 to 12 inches.\n\nAn indication of the significance of the heat absorbed by the fuel may be obtained from a comparison with the fuel latent heat of vaporization. The fuel used had a latent heat of vaporization of 145 Btu per pound. The data in figure 10 therefore indicate that a mean of 55 percent of the latent heat of vaporization can be supplied to the fuel at the zero position between the flame holder and the preheater and a mean of 83 percent of the latent heat of vaporization can be supplied to the fuel by the preheater if the distance is increased to 12 inches.\n\nBecause the performance of a ram jet is adversely affected by the total-pressure losses the air undergoes in flowing through the engine, the internal pressure losses should be kept as low as possible. The introduction of a preheater in the combustion chamber would contribute toward increasing these pressure losses. An evaluation was therefore made of the effect of the preheater used in this investigation on the combustion-chamber pressure loss. The total-pressure loss across the flame holder alone was 1.5 times the combustion-chamber-inlet dynamic pressure. Because it was believed that the combined pressure loss of the flame holder and the fuel preheater would be a maximum with the preheater 12 inches downstream of the flame holder, the combined loss was determined for only that preheater position. With the preheater located 12 inches downstream of the flame holder, the combined loss across both was only 1.6 times the combustion-chamber-inlet dynamic pressure, an increase of 6.7 percent. This slight increase is negligible when compared with the performance gains possible by preheating the ram-jet fuel (references 1 and 2).\n\nNo definite variation of combustion efficiency with preheater position could be established. In general, the combustion efficiency ranged from 60 to 80 percent. Approximately the same range of efficiencies was obtained when this flame holder was used with no internal preheater installed, in which case the fuel was preheated with an external steam-heat exchanger (reference 4).\n\nSUMMARY OF RESULTS\n\nFrom a subsonic sea-level investigation undertaken to evaluate the effectiveness of a simple internal regenerative fuel preheater for a 20-inch ram jet, the following results were obtained:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:23:07.287802+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 34, "total_pages": 53, "image_filename": "19930082542_p34.jpg", "text": "```markdown\n34\nNACA TN No. 1867\n\nProperties at room temperature\n\nBrinell hardness\n350\n250\n150\n\nTensile strength\n160,000\n140,000\n120,000\n100,000\n\nStress, psi\n0.02-percent-offset yield strength\n80,000\n60,000\n40,000\n\nElongation, percent\n40\n20\n\nRupture properties at 1200° F\n\nRupture strength\n60,000\n50,000\n40,000\n30,000\n\nStress, psi\n100 hr\n1000 hr\n\nRupture elongation\n40\n20\n0\n\nElongation, percent\n100 hr\n\nHot-rolled\n1350 1500 1750\nTemperature, °F\nAged 24 hr\n\n5 10 15 20\nReduction, percent\nHot-cold-rolled at 1200° F\n\n20% red. at 1200° F\n65% red. 1800° to 1400° F\n1800° 2 hr A.C.; 20% reduction at 1400° F repeated six times; 1800° A.C.\n\nx 10-percent reduction at room temperature\no 1400° F 24 hr after 15 percent hot-cold-work\n\nNACA\n\nFigure 3.- Effect of aging and rolling on properties of hot-rolled bar stock of low-carbon N-155 alloy.\n```", "timestamp": "2026-07-22T05:23:08.745382+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 21, "total_pages": 29, "image_filename": "19930085879_p21.jpg", "text": "NACA RM L9D11\n\n[Figure: Graph with three curves labeled Longitudinal, Transverse, Normal; vertical axis labeled \"Acceleration, g\" ranging from -6 to 2; horizontal axis labeled \"Time from launching, sec\" ranging from 7 to 22; NACA logo in bottom right corner of graph]\n\nFigure 7.— Portion of accelerometer curves, RM-11A.\n\n19", "timestamp": "2026-07-22T05:23:08.929283+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 37, "total_pages": 58, "image_filename": "19930082617_p37.jpg", "text": "36\nNACA TN 1962\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X 10^3\n2 72.0 X 10^3\n3 108.0 X 10^3\n4 144.0 X 10^3\n5 180.0 X 10^3\n\n2.57\"\nBand L\nA\nA-A\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n10 8 6 4 2 0 -2 -4 -6 -8 -10 X 10^-4\nStrain\n\n1\n2\n3\n4\n5\n\nNACA\n\nFigure 25.- Strain diagram of cylinder 79. Band L.", "timestamp": "2026-07-22T05:23:09.903906+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 23, "total_pages": 24, "image_filename": "19930085626_p23.jpg", "text": "22\n\nCONFIDENTIAL\n\n$$\n\\frac{pb}{2V}\n$$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:23:11.866070+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 15, "total_pages": 37, "image_filename": "19930085889_p15.jpg", "text": "14 CONFIDENTIAL NACA RM L9F14\n\ntherefore, is determined primarily by $C_{l_p}$. All of the wings show reductions in rolling effectiveness as the lift coefficient is increased up to about 0.5. In the case of the $46.7^\\circ$ sweptback wing, this reduction amounts to about 40 percent of the value at zero lift. At higher lift coefficients $(pb/2V)_\\delta$ increases for all of the wings because $C_{l_p}$ decreases more rapidly than $C_{l_\\delta}$. The values of $(pb/2V)_\\delta$ presented in figure 18, however, (as previously mentioned) neglect any possible effect of aileron deflection on the damping in roll or of rolling on aileron effectiveness.\n\nCONCLUSIONS\n\nAn investigation made in the Langley stability tunnel of a series of thin sweptback wings of aspect ratio 4, each tested in combination with a fuselage, indicates the following conclusions:\n\n1. The maximum lift coefficient of the wing-fuselage combinations increased as the angle of sweepback increased. At lift coefficients below 0.8, the lift curves were very nearly the same for all three models. The usual effect of sweepback in reducing the lift-curve slope appeared to be confined to the lift-coefficient range between about -0.2 and 0.2 but was less than was expected, probably because the usual effect of sweepback was masked by a variable influence of the fuselage.\n\n2. The aerodynamic center at low lift coefficients moved rearward from 17.6 percent to 27.0 percent of the mean aerodynamic chord as the sweep angle was increased from $3.6^\\circ$ to $46.7^\\circ$. This rearward shift was considerably larger than that indicated by theory for plain wings and appears to have been caused by a variable contribution of the fuselage. For the $46.7^\\circ$ sweptback wing, tests showed that the fuselage had almost no effect on the aerodynamic center; but for the $3.6^\\circ$ sweptback wing, the fuselage is believed to have a destabilizing effect, as is usually expected.\n\n3. At low lift coefficients the derivative of rolling moment due to yaw varied linearly with lift coefficient, and the rate of variation increased with an increase in sweep angle in very much the manner that is predicted by theory. The linear variations were maintained over only very small ranges of lift coefficient for the more highly swept wings; however as a result, the maximum positive values of the derivative of rolling moment due to yaw for the $32.6^\\circ$ and $46.7^\\circ$ sweptback wings were smaller than the values of this derivative for the $3.6^\\circ$ sweptback wing at lift coefficients greater than 0.6.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:23:13.072361+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 40, "total_pages": 46, "image_filename": "19930085519_p40.jpg", "text": "NACA RM No. L8K19\n39\n\nPlug-slot\nlower lip\n$\\alpha$\n(deg)\nSharp\n0\nSharp\n8\nFaired\n0\nFaired\n8\n\nSpoiler 18\n(Reference 1)\n0.1\n6.9\n\nRolling-moment\ncoefficient, $C_l$\n.03\n.02\n.01\n0\n-.01\n\n-8\n-6\n-4\n-2\n0\n\nPlug aileron or spoiler aileron\nprojection, $\\delta_p$, percent c\n\nNACA\n\nFigure 18.- Variation of rolling-moment coefficient with control\nprojection for various plug- and spoiler-aileron configurations\non the 42° sweptback wing. Flap retracted.", "timestamp": "2026-07-22T05:23:19.799597+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 28, "total_pages": 33, "image_filename": "19930085544_p28.jpg", "text": "NACA RM No. L8K26\n27\n\n<!-- Image (186, 120, 822, 848) -->\n\nFigure 10.- Variation of force coefficient with blade position. k,0.1.", "timestamp": "2026-07-22T05:23:27.001248+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 23, "total_pages": 36, "image_filename": "19930085869_p23.jpg", "text": "CONFIDENTIAL\nSpray in propellers\nGross load coefficient for stability\nand resistance tests.\nNACA RM L59D15\n\nGross load coefficient, $C_{\\Delta L_0}$\n4\n3\n2\n1\n0\n\nPropellers\nclear\nSpray strip similar to that on unswept model.\nFinal spray strip used throughout investigation.\n\nSpeed coefficient, $C_V$\n0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0\n\nNACA\n\nFigure 6.— Gross load coefficients at which spray entered propellers for two spray-strip arrangements.\nCONFIDENTIAL\n21", "timestamp": "2026-07-22T05:23:30.737017+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 3, "total_pages": 23, "image_filename": "19930085934_p3.jpg", "text": "2\nNACA RM E9G12\n\nlaboratory and is presented herein. Inasmuch as the isentropic-\nadiabatic compression process is used in computing efficiency with\ndry compression, the same process is used in this method. Conditions\nof the water-air mixture at the inlet are arbitrarily set in order\nto simplify the method. Although this method can be applied to\nmixtures that include injection liquids for which extensive thermo-\ndynamic data are available, water alone is used in the numerical\nexample in order to simplify computation. Because the water con-\ntained in the air changes to vapor during the compression process,\nthe latent heat of vaporization must be accurately evaluated.\nAnother complication is the variation in the specific heat of the\nvapor, which precludes the use of constant specific heats in cal-\nculating efficiency. The use of variable specific heats would be\ncumbersome and impractical. For computing changes in enthalpy\ntrial-and-error methods based on the steam and air tables of refer-\nences 1 and 2 were used to eliminate the use of variable-specific-\nheat formulas.\n\nThe method outlined herein was used to compute theoretical\nvalues needed for evaluating performance. Functional charts are\ngiven for isentropic outlet temperature, isentropic enthalpy change,\nand adiabatic efficiency for a range of pressure ratios and fixed\ninlet conditions. Water-air ratio is plotted as a parameter in all\nthe charts.\n\nSYMBOLS\n\nThe following symbols are used in the calculations and the\nfigures:\n\n| | |\n| :--- | :--- |\n| $c_p$ | specific heat at constant pressure, (Btu/(lb)($^\\circ$F)) |\n| h | enthalpy, (Btu/lb) |\n| N | ratio of actual pressure to base pressure |\n| P | total pressure, (lb/sq in.) or (in. Hg absolute) |\n| q | specific humidity, (lb water/lb air) |\n| R | universal gas constant for air, 53.35 (ft-lb/(lb)($^\\circ$F)) |\n| s | entropy, (Btu/(lb)($^\\circ$F)) |\n| T | temperature, (459.7 + t) ($^\\circ$R) |\n| t | temperature, ($^\\circ$F) |", "timestamp": "2026-07-22T05:23:31.390631+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 46, "total_pages": 66, "image_filename": "19930082914_p46.jpg", "text": "NACA TN No. 1857\n45\n\n[Figure: A black and white image showing a pattern of horizontal, parallel dark and light bands (fringes). On the left side, two short, dark, rectangular objects protrude into the fringe pattern. In the bottom corners, there are small triangular markers.]\n\nFigure 10.- Undisturbed fringes.\nNACA", "timestamp": "2026-07-22T05:23:32.613961+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 19, "total_pages": 96, "image_filename": "19930085880_p19.jpg", "text": "NACA RM No. L9C03\n17\n\n[Figure: A graph plotting Wetted length, ft (y-axis) against Total wetted area, sq ft (x-axis). The y-axis ranges from 0 to 2.0. The x-axis ranges from 0 to .4. Two lines are plotted, labeled \"At center line\" and \"At chine\". A small diagram of a hull cross-section is shown in the upper left. The NACA logo is in the lower right.]\n\n(e) $\\tau = 20^\\circ$.\n\nFigure 6 .- Concluded.", "timestamp": "2026-07-22T05:23:38.653485+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 14, "total_pages": 31, "image_filename": "19930085881_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:23:42.103549+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 41, "total_pages": 78, "image_filename": "19930082618_p41.jpg", "text": "NACA TN 1945\n\nSection lift coefficient, $c_l$\n\nMoment coefficient, $c_{m_{c/4}}$\n\nSection angle of attack, $\\alpha_s$, 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$\\triangleright$ 6.0\n$\\triangleleft$ 9.0\n\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\n\nFigure 7.- Aerodynamic characteristics of the NACA 64$_1$-612 airfoil section, 24-inch chord.\n\n39", "timestamp": "2026-07-22T05:23:43.972829+00:00"}
{"citation_id": "19930085906", "source_url": "https://ntrs.nasa.gov/api/citations/19930085906/downloads/19930085906.pdf", "page_number": 8, "total_pages": 23, "image_filename": "19930085906_p8.jpg", "text": "NACA RM E5F20 CONFIDENTIAL 7\n\n1. Between 55 and 83 percent of the latent heat of vaporization could be supplied to the fuel by the preheater.\n\n2. At a given engine condition, the amount of heat added to the fuel and therefore the final fuel temperature increased as the distance between flame holder and preheater was increased from 0 to 8 inches. A further increase to 12 inches resulted in little improvement.\n\n3. For the configuration investigated, approximately 5 minutes of operation after ignition was required before a stable fuel temperature was attained (initial temperature, $63^\\circ \\pm 7^\\circ$). This time did not appear dependent on preheater position or ram-jet operating condition. The rate of fuel-temperature rise was maximum immediately after ignition and a useful preheat temperature of $200^\\circ$ F was approached within 2 minutes.\n\n4. The combustion efficiencies obtained did not appear to be affected by the presence of the preheater nor by the fuel-preheater position.\n\n5. The total-pressure loss across the combustion chamber was increased only 6.7 percent when the preheater was added.\n\nCONCLUSIONS\n\nAs a result of this investigation at subsonic sea-level conditions, the following conclusions are indicated: An internal regenerative preheater of the type investigated can be used to preheat fuels to temperatures that result in improved combustion efficiencies. Internal regenerative preheaters may be designed for which the additional internal pressure loss resulting from the introduction of the preheater into the combustion chamber may be considered as negligible. In order to permit operation over a wide range of engine conditions and to prevent fuel vapor lock, the system should employ a variable-orifice fuel injector to permit the fuel pressure to remain above the fuel vapor pressure over a wide range of fuel flows. Because a period of 1 to 2 minutes is required to reach useful preheat temperatures, the application of such a preheater is limited to ram jets having a flight duration greater than this period of time.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:23:49.904030+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 22, "total_pages": 29, "image_filename": "19930085879_p22.jpg", "text": "```markdown\n20\n\n[Figure: Graph showing Accelerometer (Gs) vs Time from launching, sec. Annotations include: \"Main rocket burned out\", \"Peak acceleration during ejection, 10.8 gs\", \"Transverse\", \"Normal\", \"Longitudinal\", \"Flaps start to open\". NACA logo present.]\n\nFigure 8.- Portion of accelerometer curves, RM-11B.\n\nNACA RM L9D11\n```", "timestamp": "2026-07-22T05:23:50.654680+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 2, "total_pages": 42, "image_filename": "19930085938_p2.jpg", "text": "D\n\nNACA RM No. L9B04\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nHYDRODYNAMIC CHARACTERISTICS OF AERODYNAMICALLY REFINED PLANING-TAIL HULLS\n\nBy Robert McKann and Claude W. Coffee\n\nSUMMARY\n\nInvestigations were made in order to determine the hydrodynamic characteristics of two aerodynamically refined planing-tail hulls. One hull had an afterbody that was a tapered boom and the other had two afterbodies consisting of tapered booms fairing out of the engine nacelles. Over a wide range of center-of-gravity location, both models had a large range of elevator deflection for stable take-offs. The lower trim limit of stability peak was high for both configurations but trims obtainable were great enough to permit operation above the lower trim limit. No upper-limit porpoising was encountered with either configuration. Stable landings could be made over a wide range of trim and location of the center of gravity, provided the vertical chine strips were not extended to the point of the step. Extension of the vertical chine strips to the point of the step resulted in unstable landings. The relatively high trims and the vertical chine strips were effective in reducing the propeller spray. The hump load-resistance ratios for the aerodynamically refined hulls were low (2.9 to 3.6). Directional instability was noticed over a short range of speed with the single-boom configuration. The twin booms provide a substantial amount of transverse righting moment.\n\nINTRODUCTION\n\nIn order to obtain flying-boat forms that will permit increased range and speed over those in present-day use, several refinements of the planing-tail type of hull have been investigated in the Langley 300 MPH 7- by 10-foot tunnel and in Langley tank no. 2. The air drag of the planing-tail flying-boat hull employing a deep step and full-step fairing has been shown in reference 1 to be considerably less than that of a comparable conventional-type hull. In reference 2, the hydrodynamic characteristics of this planing-tail-hull configuration were shown to be an improvement over those of a conventional hull. The aerodynamic characteristics of several modifications of the planing-tail type of hull embodying an airfoil-section forebody plan form and slender \"boom like\" afterbodies have been reported in reference 3. This aerodynamic refinement resulted in a decrease in hull", "timestamp": "2026-07-22T05:23:50.831862+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 35, "total_pages": 53, "image_filename": "19930082542_p35.jpg", "text": "NACA TN No. 1867\n35\n\nProperties at room temperature\nBrinell hardness\nTensile strength\n0.02-percent-offset yield strength\nElongation\nRupture properties at 1200° F\nRupture strength\nRupture elongation\n\n| Solution time, hr | 1 | 2 | 2 | 5 |\n| :--- | :--- | :--- | :--- | :--- |\n| Cooling | W.Q. | W.Q. | A.C. | W.Q. |\n| Solution temperature, °F | 2050 | | | |\n\nProperties at room temperature\nBrinell hardness\nTensile strength\n0.02-percent-offset yield strength\nElongation\nRupture properties at 1200° F\nRupture strength\nRupture elongation\n\n| Solution temperature, °F | 1800 | 1900 | 2000 | 2100 | 2200 | 2300 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Hot-rolled | | | | | | |\n| Solution time, hr | 2 | 2 | 2 | 1 | 1 | 1/2 |\n| Cooling | Water-quenched | | | | | |\n\nFigure 4.- Effect of time and cooling rate on properties of solution-treated bar stock of low-carbon N-155 alloy.\nFigure 5.- Effect of solution-treating temperature on properties of low-carbon N-155 alloy bar stock.", "timestamp": "2026-07-22T05:23:51.372390+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 38, "total_pages": 58, "image_filename": "19930082617_p38.jpg", "text": "NACA TN 1962\n37\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X $10^3$\n2 72.0 X $10^3$\n3 108.0 X $10^3$\n4 144.0 X $10^3$\n5 180.0 X $10^3$\n\n[Figure: Cross-section diagram labeled A-A with dimensions 2.57\" and 45° angles, and label \"Band V\"]\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\nStrain\n10 8 6 4 2 -2 -4 -6 -8 -10 X $10^{-4}$\n\n[Graph showing multiple curves labeled 1 through 5, with data points marked as O and X corresponding to stringers]\n\n[NACA logo]\n\nFigure 26.- Strain diagram of cylinder 79. Band V.", "timestamp": "2026-07-22T05:23:53.157244+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 6, "total_pages": 42, "image_filename": "19930085914_p6.jpg", "text": "NACA RM A9D25\n\nattack and Mach number at a Reynolds number of 2.0 million.\n\nCORRECTIONS\n\nThe data have been corrected for the effects of tunnel-wall interference, constriction due to the tunnel walls, base pressure, and static tares due to the weight of the model. No correction has been applied to account for the effect of flap deflection under load upon the force and moment coefficients presented. At the highest loading condition, this deflection was of the order of $1^\\circ$. The angle of attack of the model was measured visually by means of a cathetometer, hence no corrections were necessary to account for deflection of the support equipment.\n\nTunnel-Wall Interference\n\nCorrections to the data due to induced tunnel-wall interference have been evaluated by the method of Glauert (reference 4). Since the ratio of model span to tunnel diameter was small, the total corrections were small, and no account was taken of the sweepback of the wing. The following corrections were added:\n\n$$\n\\Delta \\alpha = 0.26 \\, C_L\n$$\n\n$$\n\\Delta C_L = 0.0046 \\, C_L^2\n$$\n\nNo correction was applied to the pitching moment.\n\nConstriction Effects\n\nThe constriction effects of the tunnel walls have been evaluated by the method of reference 5. This method has not been modified to account for the effects of sweepback. The magnitude of the corrections applied to the Mach number and to the dynamic pressure is illustrated by the following table:\n\n| Corrected Mach number | Uncorrected Mach number | | q, corrected / q, uncorrected | |\n|---|---|---|---|---|\n| | Wing and fuselage | Fuselage alone | Wing and fuselage | Fuselage alone |\n| 0.930 | 0.919 | 0.921 | 1.012 | 1.012 |\n| .920 | .911 | .912 | 1.010 | 1.010 |\n| .890 | .884 | .885 | 1.007 | 1.007 |\n| .850 | .846 | .847 | 1.005 | 1.005 |\n| .800 | .798 | .798 | 1.003 | 1.003 |\n| .700 | .698 | .699 | 1.002 | 1.002 |\n| .600 | .599 | .599 | 1.002 | 1.002 |\n| .400 | .399 | .400 | 1.001 | 1.001 |\n| .200 | .200 | .200 | 1.001 | 1.001 |", "timestamp": "2026-07-22T05:23:53.812953+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 57, "total_pages": 99, "image_filename": "19930082511_p57.jpg", "text": "NACA TN No. 1826\n55\n\naround the contour indicated in figure 27. Its value is the sum of the residues of the integrand at its poles inside the contour. These poles are the values of $z$ for which $J_m'(iz) = 0$. The zeros of $J_m'$ will be designated $y_{sm}$; they are real and may be obtained from the formula in appendix III of reference 15. The poles of the integrand then occur at $iz = y_{sm}$, that is, at $z = -iy_{sm}$. Since only the poles within the contour are desired, only the negative zeros of $J_m'$ are considered.\n\nThe residue of $\\frac{1}{J_m'(iz)}$ at $z = -iy_{sm}$ is,\n\n$$\n\\lim_{z \\to -iy_{sm}} \\frac{z + iy_{sm}}{J_m'(iz)} = \\lim_{z \\to -iy_{sm}} \\frac{z + iy_{sm}}{J_m'(iz) - J_m'(y_{sm})}\n$$\n\nsince $J_m'(y_{sm}) = 0$; by the definition of the derivative this expression reduces to\n\n$$\n- \\frac{i}{J_m''(y_{sm})}\n$$\n\nThe residue of the integrand at $z = -iy_{sm}$ is thus\n\n$$\ni \\frac{J_m(\\rho y_{sm}) e^{ky_{sm}}}{y_{sm}(y_{sm}^2 + h^2) J_m''(y_{sm})}\n$$\n\nBut the Bessel functions satisfy the relation\n\n$$\nJ_m'' + \\frac{1}{x} J_m' + \\left(1 - \\frac{m^2}{x^2}\\right) J_m = 0\n$$\n\nso that at $x = y_{sm}$, where $J_m' = 0$,\n\n$$\nJ_m''(y_{sm}) = \\frac{m^2 - y_{sm}^2}{y_{sm}^2} J_m(y_{sm})\n$$\n\nwhence, finally,\n\n$$\n\\frac{1}{2\\pi i} \\oint \\frac{J_m(i\\rho z) e^{ikz} dz}{iz J_m'(iz) (z^2 - h^2)} = i \\sum_s \\frac{J_m(\\rho y_{sm}) e^{ky_{sm}} y_{sm}}{(y_{sm}^2 + h^2) (m^2 - y_{sm}^2) J_m(y_{sm})}\n$$", "timestamp": "2026-07-22T05:23:54.138989+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 24, "total_pages": 24, "image_filename": "19930085626_p24.jpg", "text": "```markdown\nCONFIDENTIAL\n\nNACA RM No. L8K23\n\n$$ \\frac{pb}{2V} $$\n\n| Aileron | $\\delta_a$, deg |\n| :--- | :--- |\n| Outboard | 5 |\n| -do- | 10 |\n| Inboard | 5 |\n| Full span | 5\n\n.12\n.08\n.04\n0\n-.04\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\nM\n\n[Figure: NACA logo]\n\nFigure 7.— Summary of results for true-contour ailerons. Curves for outboard ailerons obtained by averaging results shown in figure 4(a).\nCONFIDENTIAL\n\n23\n```", "timestamp": "2026-07-22T05:23:55.261071+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 49, "total_pages": 49, "image_filename": "19930082498_p49.jpg", "text": "```markdown\n48\n\nSound-pressure level, db\n90\nOver-all\nF\n80\n3F\n70\n2F\n60\n50\n0\n10\n20\n30\n40\n50\n60\n70\n80\n90\nTail-pipe length, in.\n\n[Figure: A line graph showing the effect of tail-pipe length on sound level. The y-axis represents \"Sound-pressure level, db\" ranging from 50 to 90. The x-axis represents \"Tail-pipe length, in.\" ranging from 0 to 90. There are five data series plotted: \"Over-all\" (solid line with circles), \"F\" (dashed line with squares), \"3F\" (solid line with diamonds), \"2F\" (solid line with triangles), and an unlabeled dashed line with inverted triangles. The NACA logo is visible in the bottom right corner of the plot area.]\n\nFigure 8.- Effect of tail-pipe length on sound level. Muffler configurations 22 to 28; engine speed, 2000 rpm; fundamental frequency F, 100 cps.\n\nNACA TN No. 1838\n```", "timestamp": "2026-07-22T05:23:58.777206+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 16, "total_pages": 37, "image_filename": "19930085889_p16.jpg", "text": "NACA RM L9F14 CONFIDENTIAL 15\n\n4. The derivative of yawing moment due to rolling was either zero or positive through most of the lift-coefficient range for each of the wings tested.\n\n5. At zero lift coefficient there is a decrease in damping in roll with an increase of sweepback. The values obtained in the Langley stability tunnel by the rolling-flow technique show good agreement throughout the sweep range with those obtained by free rotation of the models in the Langley 7- by 10-foot tunnel and with those calculated by Weissinger's theory.\n\n6. An increase in sweepback caused large reductions in the rolling moment and in the wing-tip helix angle resulting from a unit angular deflection of the ailerons about their hinge axes. For the $46.7^\\circ$ swept-back wing, the rate of variation of wing-tip helix angle with aileron deflection decreased by about 40 percent as the lift coefficient increased from 0 to 0.5 but then increased slightly with a further increase in lift coefficient.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:23:59.641125+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 29, "total_pages": 33, "image_filename": "19930085544_p29.jpg", "text": "28\nNACA RM No. L8K26\n\n<!-- Image (122, 129, 822, 866) -->\n\nFigure 11.- Variation of force coefficient with k. $\\frac{a}{\\alpha p_0} = 0$.", "timestamp": "2026-07-22T05:24:03.451045+00:00"}