PDF/ocr_dataset_1028.jsonl

{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 4, "total_pages": 24, "image_filename": "19930085626_p4.jpg", "text": "```markdown\nNACA RM No. L8K23 CONFIDENTIAL 3\n\n$\\frac{\\theta}{m}$ wing torsional-stiffness parameter\n\nm concentrated couple applied near wing tip in plane parallel\nto model center line and normal to wing chord plane,\ninch-pounds\n\n$\\theta$ angle of twist produced by m at any section along wing\nspan in plane parallel to that of m, radians\n\n### TEST VEHICLES AND TESTS\n\nThe general arrangement of the test vehicles is shown in figures 1\nand 2. Further pertinent information is contained in table I. The test\nvehicles, which were relatively simple, inexpensive, and expendable,\nconsisted of a pointed cylindrical wooden body to which the particular\nwing-aileron configuration under investigation was attached in a three-\npanel arrangement. Unpublished tests of 3- and 4-panel arrangements\nindicate that, with regard to the rolling-effectiveness characteristics,\nthe interference effects between the wings were negligible. A small\nradio transmitter, designated spinsonde, which produced a plane polarized\nsignal was enclosed in the pointed nose of the body for the measurement\nof rolling velocity. The wings, which were constructed mainly of wood,\nwere stiffened by means of steel plates cycle-welded into the upper and\nlower wing surfaces as shown in figure 1. The measured torsional-\nstiffness characteristics of two typical wings are shown in figure 3.\nThe degree of torsional stiffness indicated by the curves of figure 3\nhas been shown by tests reported in reference 4 to be sufficient to make\nthe effects of wing twisting negligible.\n\nThe geometric details of the wing-aileron configurations tested are\nshown in figure 4. The configuration shown in figure 4(a) is that of\nreference 1. The ailerons were formed by deflecting the chord line of\nthe basic section at the indicated hinge line. This method of construction\nsimulates plain, sealed ailerons in actual aircraft construction. For\nthe majority of the test flights the aileron deflection was $5^\\circ$.\n\nThe test vehicles were propelled by a two-stage rocket-propulsion\nsystem to a Mach number of about 1.9. During coasting flight following\nburnout of the rocket motor, time histories of the rolling velocity\nproduced by the ailerons (obtained with spinsonde radio equipment) and\nthe flight-path velocity (obtained with Doppler radar) were recorded.\nThese data, in conjunction with atmospheric data obtained with radio-\nsondes, permitted the evaluation of the rolling-effectiveness param-\neter $pb/2V$ as a function of Mach number. Also the drag coefficient\nof the test vehicles was obtained by a process involving the graphic\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T05:06:29.952705+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 30, "total_pages": 30, "image_filename": "19930082585_p30.jpg", "text": "NACA TN 1907\n29\n\n[Figure: Graph plotting Descending velocity, V, ft/sec (y-axis, 0 to 40) against Altitude lost, ft (x-axis, 0 to 200). The graph contains three curves representing different Moments of inertia, $I_1$ (slug-ft$^2$): a dashed line for 100, a solid line for 200, and a dash-dot line for 400. The NACA logo is visible in the bottom right corner of the plot area.]\n\nFigure 10.- Effect of blade moment of inertia on descending velocity against altitude lost. Slow exponential pitch change.", "timestamp": "2026-07-22T05:06:33.035961+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 35, "total_pages": 37, "image_filename": "19930082450_p35.jpg", "text": "34\nNACA TN No. 1778\n\n45\n40\n35\n30\n25\n20\n15\n10\n40\n\n$$ \\frac{H}{t_W} = 21 $$\n$$ \\frac{D_W}{t_W} = 20 $$\n\n$$ \\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S} $$\n\n$$ \\sigma_{cr}, \\text{ksi} $$\n13.8\n8.9\n\n$$ \\frac{P_1}{L\\sqrt{c}}, \\text{ksi} $$\n\n.05\n.075\n.10\n.125\n.15\n.20\n.25\n.30\n.40\n.50\n.60\n.75\n1.00\n\nColors indicate minimum weight proportions for $$ \\frac{t_W}{t_S} = 1.00 $$.\nRed means some other, blue means no other value of $$ \\frac{t_W}{t_S} $$ gives less weight.\n\n35\n30\n25\n20\n15\n35\n\n31\n(30)\n\n$$ \\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S} $$\n\n$$ \\sigma_{cr}, \\text{ksi} $$\n19.8\n13.8\n8.9\n\n$$ \\frac{P_1}{L\\sqrt{c}}, \\text{ksi} $$\n\n.05\n.075\n.10\n.125\n.15\n.20\n.25\n.30\n.40\n.50\n.60\n\n30\n25\n20\n15\n\n41\n(40)\n\n$$ \\frac{S}{t_S} \\text{ or } \\frac{b_S}{t_S} $$\n\n$$ \\sigma_{cr}, \\text{ksi} $$\n18.4\n13.8\n8.9\n\n$$ \\frac{P_1}{L\\sqrt{c}}, \\text{ksi} $$\n\n.05\n.075\n.10\n.125\n.15\n.20\n.25\n.30\n.40\n.50\n.60\n\nNACA\n\n20 30 40 50 60 70 80 90 100 110 120\n\n$$ \\frac{P_1}{t_S}, \\text{ksi} $$\n\nFigure 9-Direct-reading design chart (alternate form) for 24S-T aluminum-alloy Z-stiffened panels. $$ \\frac{t_W}{t_S} = 1.00 $$.", "timestamp": "2026-07-22T05:06:33.481132+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 18, "total_pages": 46, "image_filename": "19930085519_p18.jpg", "text": "NACA RM No. L8K19\n17\n\n[Figure: A photograph showing the lower surface of a sweptback wing mounted in a wind tunnel. The wing has a full-span slotted flap deflected downwards. A label in the bottom right corner of the image reads \"NACA L-54920\".]\n\n(a) Wing lower surface.\n\nFigure 2.- The $42^\\circ$ sweptback wing mounted in the Langley 300 MPH 7- by 10-foot tunnel. Full-span slotted flap deflected $50^\\circ$.", "timestamp": "2026-07-22T05:06:39.105175+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 33, "total_pages": 50, "image_filename": "19930082496_p33.jpg", "text": "32\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:06:46.695068+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 21, "total_pages": 36, "image_filename": "19930085487_p21.jpg", "text": "NACA RM No. E8J22\n\nINCHES\n0 1 2\n\nNACA\nC-14381\n2-27-46\n\nFigure 6. - Sand pattern for fourth bending-mode vibration on first-stage blade. Frequency of vibration, 9600 cycles per second.\n\n19", "timestamp": "2026-07-22T05:06:47.203349+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 17, "total_pages": 20, "image_filename": "19930085536_p17.jpg", "text": ".40\n90°\n.32\n.24\n.16\n.08\n0\n-.08\n-.16\n-.24\n\nPressure coefficient, $C_p$\n\nAngle of yaw, $\\psi$ (deg)\n3\n6\n\nAngle of yaw $\\psi$ (deg)\n0\n3\n6\n9\n12\n16\n\nExperimental data\nLinearized theory (equation (8))\nLinearized theory (equation (9))\n\nYaw axis\n180°\n\n$\\theta$, deg\n0\n30\n60\n90\n120\n150\n180\n\nNACA RM No. E8K05\n\nFigure 4. - Variation of pressure coefficient with position on model at an angle of attack of 0°.\n\nNACA\n15", "timestamp": "2026-07-22T05:06:47.367980+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 47, "total_pages": 66, "image_filename": "19930082245_p47.jpg", "text": "```markdown\n12\n10\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n-10\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\n\nSection angle of attack, $\\alpha$, deg\n\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\n.24\n.20\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\n\nSection pitching-moment coefficient, $C_m$\n\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\n(e) $c_n = 0.4$.\nFigure 9.—Continued.\n\n46\nNACA\nNACA TN No. 1596\n```", "timestamp": "2026-07-22T05:06:48.678758+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 10, "total_pages": 46, "image_filename": "19930085542_p10.jpg", "text": "longitudinal stability and in the lift-curve slope. Increased longitudinal stability is noted throughout the lift-coefficient range as the aspect ratio is decreased. (See fig. 11.) As the aspect ratio is reduced the aerodynamic center moves rearward toward the 50-percent-chord point (fig. 20) as is indicated for very low aspect ratios by the theory of reference 2. The theory of reference 3 indicates a more gradual movement of the aerodynamic center starting from a more forward position. It should be noted that reference 3 neglects any possible changes in the chordwise pressure distributions of the wing. The empirical curve taken from reference 8 indicates the same trend as the experimental; however, the aerodynamic center is in a more forward position than the results of the present investigation indicate. It should be noted that all the triangular wings tested in reference 8 had flat-plate airfoil sections, whereas those tested herein had an NACA 0012 profile with a larger trailing-edge angle and a blunt leading edge. Large-scale tests of a triangular wing with a double-wedge airfoil (5 percent thick at 20 percent chord) indicate about the same position of the aerodynamic center as does the present investigation for an aspect ratio of 2.0. (See reference 9.)\n\nThe trend of $C_{L_{\\max}}$ in figure 20 agrees with the trends of reference 8 in that the peak value of $C_{L_{\\max}}$ was reached at about the same aspect ratio. As the aspect ratio is decreased, the lift-curve slope is decreased as can be seen in figure 20. The swept-wing theory of references 3 and 5 shows fair agreement with the experimental data for the aspect-ratio range considered. The theory of reference 2 approaches the experimental values only as the aspect ratio approaches zero as would be expected.\n\nIt should be remembered that all the profiles for the models for which the data are presented in figure 12 are of NACA 0012 sections parallel to the plane of symmetry. With a highly swept model (as model 4) there is a very large area in the plane of symmetry forward of the quarter chord of the mean aerodynamic chord which, when the model is yawed or rolled, acts in the manner of a fin. Model 4, because of this area, has positive values of the directional-stability parameter $C_{n_\\psi}$ below $C_L = 0.73$. The model does have increasing directional stability at the stall while models 2 and 7 do not. If model 4 was equipped with a high-aspect-ratio fin, the objectionable characteristics below $C_L = 0.73$ might be overcome, resulting in a model having better over-all characteristics than models 2 or 7. The values of $\\partial C_{n_\\psi} / \\partial C_L^2$ presented in figure 21 (obtained by plotting $C_{n_\\psi}$ against $C_L^2$ and taking slopes at $C_L = 0$) indicate increasing directional stability as the aspect ratio is decreased. (The values are not presented for model 4 because of the erratic nature of the curve of $C_{n_\\psi}$.) The theory of reference 5 is in qualitative agreement with the experimental results. Very high maximum values of the", "timestamp": "2026-07-22T05:06:52.880491+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 3, "total_pages": 36, "image_filename": "19930085869_p3.jpg", "text": "NACA RM L9D15\nCONFIDENTIAL\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nRESEARCH MEMORANDUM\nHYDRODYNAMIC CHARACTERISTICS OF\nA SWEPT PLANING-TAIL HULL\nBy Robert E. McKann, Claude W. Coffee,\nand Donald D. Arabian\n\nSUMMARY\n\nThe hydrodynamic characteristics of a swept planing-tail hull were\ndetermined by tests in Langley tank no. 2. The hull was derived from an\naerodynamically refined planing-tail hull by sweeping aft the water\nplanes above the chines. This procedure resulted in an aft movement of\nthe hull volume which produced a more favorable volume distribution about\nthe center of gravity. With vertical spray strips, only light spray\nstruck the propellers over a short speed range before the hump. No spray\ncame over the bow. Heavy spray struck the tail surfaces near hump speed.\nA large range of elevator deflection was available for take-offs over a\nwide range of center-of-gravity location. The minimum trim of $2^\\circ$ at high\nspeed rather than lower-limit porpoising determined the minimum elevator\ndeflection for take-off. Upper-limit porpoising occurred over a short\nspeed range near take-off. Landings at locations of the center of\ngravity from 0.20$\\bar{c}$ to 0.40$\\bar{c}$ were stable. The hump load-resistance ratio\nof 3.1 was lower than ratios obtained for conventional hulls.\n\nINTRODUCTION\n\nSeveral refinements of the planing-tail-type flying-boat hull have\nbeen made to decrease its drag. The refinements include the use of\nsymmetrical airfoil sections for the forebody plan form and slender boom-\nlike afterbodies. Tests of the hulls in the Langley 300 MPH 7- by\n10-foot tunnel (see reference 1) indicated drag approaching that of the\nfuselage of a modern transport airplane. Tank investigations, described\nin reference 2, showed that the hulls had acceptable hydrodynamic\nperformance.\n\nThe problem of airplane balance may limit the application of the\nhulls to special-purpose, high-performance airplanes, because of the\nlarge portion of the total volume forward of the center of gravity.\nSince the center-of-gravity position was fixed by aerodynamic and hydro-\ndynamic requirements, a possible solution to the balance problem was to\nmove the volume aft. A new hull, the volume of which was shifted aft\nwith respect to the center of gravity by sweeping aft the water planes\nabove the chines, was derived. Wind-tunnel tests of the new hull on a\nswept wing at low speeds (see reference 3) and high subsonic speeds\n(see reference 4) were made in the Langley 300 MPH 7- by 10-foot tunnel.\nThese tests indicated drag similar to that of the unswept hull.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:06:53.063333+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 24, "total_pages": 66, "image_filename": "19930082914_p24.jpg", "text": "NACA TN No. 1857\n\n10-inch length was covered by taking a series of seven interferograms, each of which covered a portion of the jet and its mixing region that was about $2\\frac{1}{2}$ inches high and about $1\\frac{1}{2}$ inches wide. Only the first six of these interferograms were used in obtaining the density and the velocity distributions that are given in the present paper. These six interferograms covered the first $6\\frac{1}{2}$ inches of the mixing region. The interferograms are shown as figures 11(a) to 11(f). A composite of the six interferograms is shown as figure 14.\n\nThree regions are distinguishable in these pictures. The lower portion of each picture, where the fringes have retained essentially the same spacing that they had with no air flow, was taken with light that passed through nearly undisturbed room air. The upper portion of each picture shows the free-stream region of the jet. The center portion of each picture is the mixing zone between the free room air and the free-stream portion of the jet.\n\nIf the irregularities in the fringes in the lower and the upper portions of the interferograms are averaged out, then the spacing of the fringes is found to remain the same as it was with no air flow. This shows that there is uniform density in each of the regions. In the mixing zone, however, the fringes are crowded together and there is a density gradient in that region. The density varies, in fact, from atmospheric outside the jet to about $1\\frac{1}{2}$ times atmospheric inside the jet.\n\nAs can be seen in figure 11(a), at the end of the nozzle the boundary layer that has built up along the nozzle and that emerges from the nozzle is much thicker than the beginning of the jet mixing region. At a distance of about 2 inches downstream from the nozzle, however, this boundary layer has lost its identity and there is only the mixing zone, the undisturbed jet, and the room air. Only data taken from the region between 2 and $7\\frac{1}{2}$ inches downstream from the nozzle are included in the present paper. In that region seven vertical cross sections were chosen, at 2, $2\\frac{1}{2}$, $3\\frac{1}{2}$, $4\\frac{1}{2}$, $5\\frac{1}{2}$, 6, and $7\\frac{1}{2}$ inches from the nozzle.\n\nDensity Distribution\n\nThe density variation along each cross section was obtained. The method of obtaining the density variation was first to measure the variation of fringe shift along each vertical cross section. This was done by the method described in the section entitled \"Evaluation of Density Fields.\" Figure 15 shows a plot of the density variation across the mixing zone at the seven cross sections. On the vertical axis is plotted the ratio of", "timestamp": "2026-07-22T05:07:04.083309+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 8, "total_pages": 33, "image_filename": "19930085544_p8.jpg", "text": "NACA RM No. L8K26\n\nblade at each position is determined from\n\n$$\n\\frac{T_{wt}}{B} = \\frac{\\rho n^2 D^4}{B} \\int_{x_0}^{1.0} \\left( \\frac{dC_T}{dx} \\right)_{wt} dx \\tag{4}\n$$\n\nwhere\n\n$$\n\\left( \\frac{dC_T}{dx} \\right)_{wt} = \\kappa \\pi 3 x^3 \\frac{\\alpha_1}{57.3} \\frac{\\cot \\phi - \\tan \\gamma}{\\left( \\cot \\phi + \\frac{\\alpha_1}{57.3} \\right)^2} \\left( 1 + \\frac{J}{\\pi x} \\sin \\alpha_T \\sin \\omega t \\right)^2 \\tag{5}\n$$\n\nEquation (5), except for the factor $\\left( 1 + \\frac{J}{\\pi x} \\sin \\alpha_T \\sin \\omega t \\right)^2$, is from reference 4. The quantities $\\phi$ and $\\alpha_1$ are determined by the same method as in reference 4 using the quantity $J_{wt}$ (equation (3)) in place of $J$. The additional factor $\\left( 1 + \\frac{J}{\\pi x} \\sin \\alpha_T \\sin \\omega t \\right)^2$ in equation (5) is needed to put the element thrust gradient $\\frac{dC_T}{dx}$ in terms of $\\rho n^2 D^4$ rather than in terms of the apparently varying $n$ in $J_{wt}$ (equation (3)).\n\nThe turning moment (yawing moment of a pitched propeller on the propeller shaft) is the difference in bending moments from the highly loaded side to the lightly loaded side of the inclined propeller. For the steady-force calculations this bending moment reaches a maximum on the two-blade propeller when the blades are in the horizontal plane. The maximum turning moment from the steady-flow calculations is found by graphically integrating\n\n$$\nx R \\rho n^2 D^4 \\left[ \\left( \\frac{dC_T}{dx} \\right)_{90} - \\left( \\frac{dC_T}{dx} \\right)_{270} \\right] dx \\tag{6}\n$$\n\nfrom the spinner surface to the propeller tip.", "timestamp": "2026-07-22T05:07:05.570972+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 1, "total_pages": 29, "image_filename": "19930085879_p1.jpg", "text": "NACA RM L9D11\n\nNACA\n\nRESEARCH MEMORANDUM\n\nFLIGHT INVESTIGATION OF THE JETTISONABLE-NOSE METHOD\nOF PILOT ESCAPE USING ROCKET-PROPELLED MODELS\n\nBy\nReginald R. Lundstrom and Burke R. O'Kelly\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nJune 2, 1949", "timestamp": "2026-07-22T05:07:09.651680+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 19, "total_pages": 46, "image_filename": "19930085519_p19.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:07:14.039845+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 38, "total_pages": 99, "image_filename": "19930082511_p38.jpg", "text": "36\nNACA TN No. 1826\n\nNormal velocity and normal displacement at the free surface.- The normal velocity on the free surface may be written in the following form:\n\n$$\n\\begin{aligned}\n\\text{I.P.}Q(x) &= \\frac{2i}{\\pi} (K' - E') \\int_{1}^{x} \\frac{dx}{\\sqrt{(x^2 - 1)(1 - \\frac{x^2}{a^2})}} \\\\\n&\\quad - \\frac{2iK'}{\\pi} \\int_{1}^{x} \\sqrt{\\frac{1 - \\frac{x^2}{a^2}}{x^2 - 1}} \\, dx - \\frac{2iK'}{a^2\\pi} x \\sqrt{\\frac{x^2 - 1}{1 - \\frac{x^2}{a^2}}}\n\\end{aligned}\n$$\n\nBy the substitution of $x^2 = a^2 - (a^2 - 1)t^2$, the integrals are readily reduced to standard forms of incomplete elliptic integrals, and the equation takes the form\n\n$$\n\\text{I.P.}Q(x) = \\frac{2i}{\\pi} \\left[ E' F'\\left(\\frac{a^2 - x^2}{a^2 - 1}\\right) - K' E'\\left(\\frac{a^2 - x^2}{a^2 - 1}\\right) \\right] - \\frac{2iK'}{a^2\\pi} x \\sqrt{\\frac{x^2 - 1}{1 - \\frac{x^2}{a^2}}} \\quad (25)\n$$\n\nThe normal displacement, or distortion, of the free surface is found by integrating this expression along the free surface in the physical plane:\n\n$$\n\\begin{aligned}\n\\text{Normal displacement at } x &= \\int_{1}^{x} \\text{I.P.}Q(x) \\, d\\xi \\\\\n&= \\int_{1}^{x} \\text{I.P.}Q(x) \\, \\frac{dx}{\\pi x}\n\\end{aligned}\n$$\n\nThe integral may be evaluated numerically; however, the third term of $\\text{I.P.}Q(x)$ is amenable to analytical treatment:\n\n$$\n\\begin{aligned}\n- \\frac{2iK'}{a^2\\pi} \\int_{1}^{x} x \\sqrt{\\frac{x^2 - 1}{1 - \\frac{x^2}{a^2}}} \\frac{dx}{\\pi x} &= - \\frac{2iK'}{a^2\\pi^2} \\int_{1}^{x} \\sqrt{\\frac{x^2 - 1}{1 - \\frac{x^2}{a^2}}} \\, dx \\\\\n&= - \\frac{2iK'}{\\pi^2} \\left\\{ \\left[ E' - E'\\left(\\sqrt{\\frac{a^2 - x^2}{a^2 - 1}}\\right) \\right] \\right. \\\\\n&\\quad \\left. - \\frac{1}{a^2} \\left[ K' - F'\\left(\\sqrt{\\frac{a^2 - x^2}{a^2 - 1}}\\right) \\right] \\right\\}\n\\end{aligned}\n$$", "timestamp": "2026-07-22T05:07:16.418465+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 22, "total_pages": 36, "image_filename": "19930085487_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:07:17.463181+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 34, "total_pages": 50, "image_filename": "19930082496_p34.jpg", "text": "NACA TN No. 1836\n33\n\n[Figure: A photograph of a turbine-wheel assembly. The image shows a circular metal wheel with numerous blades attached around its circumference. Labels point to different sections: \"Metal control blades\" at the top left, top right, and bottom right; \"Ceramal blades\" in the center. Handwritten markings on the wheel include \"WHEEL #19\", \"20\", \"40\", \"60\", \"80\", and \"100\". A scale bar labeled \"INCHES\" with markings for 0, 1, and 2 is visible in the bottom left corner.]\n\nNACA\nC.20621\n2-9-48\n\nFigure 7. - Turbine-wheel assembly with ceramal blades before phase 2 of investigation. Upstream side.", "timestamp": "2026-07-22T05:07:19.686516+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 36, "total_pages": 37, "image_filename": "19930082450_p36.jpg", "text": "NACA TN No. 1778\n35\n\n$$ \\frac{H}{t_w} = .26 $$\n$$ \\left( \\frac{b_w}{t_w} = 25 \\right) $$\n\n$$ \\frac{b_s}{t_s} \\text{ or } \\frac{b_s}{t_s} $$\n\n$$ \\frac{P_1}{L/\\sqrt{c}}, \\text{ ksi} $$\n\n$$ \\sigma_{cr}, \\text{ ksi} $$\n13.8\n8.9\n\n$$ \\sigma_{cr}, \\text{ ksi} $$\n19.2\n13.8\n8.9-75\n\n$$ \\sigma_{cr}, \\text{ ksi} $$\n19.7\n18.4\n17.2\n13.6\n8.9\n\n$$ \\frac{S}{t_s} \\text{ or } \\frac{b_s}{t_s} $$\n\n$$ \\frac{P_1}{L/\\sqrt{c}}, \\text{ ksi} $$\n\n$$ \\frac{S}{t_s} \\text{ or } \\frac{b_s}{t_s} $$\n\n$$ \\frac{P_1}{L/\\sqrt{c}}, \\text{ ksi} $$\n\n$$ \\sigma_{cy} = 44 \\text{ ksi} $$\n\n$$ \\frac{P_1}{t_s}, \\text{ ksi} $$\n\n$$ \\frac{t_w}{t_s} = 1.00 $$\n\nFigure 9.- Concluded.\n\nNACA", "timestamp": "2026-07-22T05:07:20.162155+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 31, "total_pages": 49, "image_filename": "19930082498_p31.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:07:21.234346+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 18, "total_pages": 20, "image_filename": "19930085536_p18.jpg", "text": ".36 Angle of\nattack, $\\alpha$\n(deg)\n10\n\n.32\n\n.28\n\n.24\n\n.20\n\n.16\n\n.12\n\n.08\n\n.04\n\n-90 -60 -30 0 30 60 90\n$\\theta$, deg\n\nAngle-of-attack\naxis\n\n90°\n\n-90°\n\nAngle of\nattack\n$\\alpha$\n(deg)\n0\n3\n6\n10\nExperimental data\n\nLinearized theory\n(equation (8))\n\nAngle of\nattack, $\\alpha$\n(deg)\n0\n3\n10\n\nNACA\n\nFigure 5. - Variation of pressure coefficient with position on model at an angle of yaw of 0°.\n\n16\n\nNACA RM NO. E8K05", "timestamp": "2026-07-22T05:07:21.404158+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 23, "total_pages": 78, "image_filename": "19930082618_p23.jpg", "text": "NACA TN 1945\n21\n\nTABLE XI\nORDINATES OF THE\nNACA 0012 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | 0 |\n| .5 | ----- | .5 | -3. |\n| 1.25 | 1.984 | 1.25 | -1.894 |\n| 2.5 | 2.613 | 2.5 | -2.613 |\n| 5 | 3.555 | 5 | -3.555 |\n| 7.5 | 4.280 | 7.5 | -4.280 |\n| 10 | 4.803 | 10 | -4.803 |\n| 15 | 5.545 | 15 | -5.545 |\n| 20 | 6.049 | 20 | -6.049 |\n| 25 | 6.341 | 25 | -6.341 |\n| 30 | 6.402 | 30 | -6.402 |\n| 40 | 6.009 | 40 | -6.009 |\n| 50 | 5.292 | 50 | -5.292 |\n| 60 | 4.350 | 60 | -4.350 |\n| 70 | 3.264 | 70 | -3.264 |\n| 80 | 2.023 | 80 | -2.023 |\n| 90 | 1.118 | 90 | -1.118 |\n| 95 | .587 | 95 | -.587 |\n| 100 | .126 | 100 | -.126 |\n\nL.E. radius: 1.58\n\nTABLE XII\nORDINATES OF THE\nNACA 1412 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | 0 |\n| 1.25 | 2.44 | 1.25 | -1.43 |\n| 2.5 | 3.39 | 2.5 | -1.95 |\n| 5 | 4.72 | 5 | -2.49 |\n| 7.5 | 5.76 | 7.5 | -2.74 |\n| 10 | 6.59 | 10 | -2.82 |\n| 15 | 7.89 | 15 | -2.88 |\n| 20 | 8.84 | 20 | -2.74 |\n| 25 | 9.41 | 25 | -2.50 |\n| 30 | 9.76 | 30 | -2.26 |\n| 40 | 9.80 | 40 | -1.80 |\n| 50 | 9.19 | 50 | -1.40 |\n| 60 | 8.14 | 60 | -1.00 |\n| 70 | 6.69 | 70 | -.65 |\n| 80 | 4.89 | 80 | -.39 |\n| 90 | 2.71 | 90 | -.22 |\n| 95 | 1.47 | 95 | -.16 |\n| 100 | (.13) | 100 | (-.13) |\n| 100 | ----- | 100 | 0 |\n\nL.E. radius: 1.58\nSlope of radius through L.E.: 0.20\n\nTABLE XIII\nORDINATES OF THE\nNACA 4415 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | ----- | 0 | 0 |\n| 1.25 | 3.07 | 1.25 | -1.79 |\n| 2.5 | 4.17 | 2.5 | -2.42 |\n| 5 | 5.74 | 5 | -3.27 |\n| 7.5 | 6.91 | 7.5 | -3.71 |\n| 10 | 7.84 | 10 | -3.98 |\n| 15 | 9.27 | 15 | -4.18 |\n| 20 | 10.27 | 20 | -4.13 |\n| 25 | 10.92 | 25 | -3.99 |\n| 30 | 11.29 | 30 | -3.75 |\n| 40 | 11.23 | 40 | -3.25 |\n| 50 | 10.53 | 50 | -2.72 |\n| 60 | 9.29 | 60 | -2.14 |\n| 70 | 7.63 | 70 | -1.55 |\n| 80 | 5.52 | 80 | -1.03 |\n| 90 | 3.08 | 90 | -.57 |\n| 95 | 1.67 | 95 | -.36 |\n| 100 | (.14) | 100 | (-.14) |\n| 100 | ----- | 100 | 0 |\n\nL.E. radius: 2.48\nSlope of radius through L.E.: 0.20\n\n[Figure: NACA logo]\n\nTABLE XIV\nORDINATES OF THE\nNACA 23012 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | ----- | 0 | 0 |\n| 1.25 | 2.67 | 1.25 | -1.23 |\n| 2.5 | 3.61 | 2.5 | -1.71 |\n| 5 | 4.91 | 5 | -2.26 |\n| 7.5 | 5.80 | 7.5 | -2.61 |\n| 10 | 6.43 | 10 | -2.82 |\n| 15 | 7.19 | 15 | -3.00 |\n| 20 | 7.50 | 20 | -2.97 |\n| 25 | 7.60 | 25 | -2.88 |\n| 30 | 7.55 | 30 | -2.66 |\n| 40 | 7.11 | 40 | -2.18 |\n| 50 | 6.41 | 50 | -1.71 |\n| 60 | 5.47 | 60 | -1.27 |\n| 70 | 4.36 | 70 | -0.90 |\n| 80 | 3.08 | 80 | -.56 |\n| 90 | 1.68 | 90 | -.29 |\n| 95 | .92 | 95 | -.17 |\n| 100 | (.13) | 100 | (-.13) |\n| 100 | ----- | 100 | 0 |\n\nL.E. radius: 1.58\nSlope of radius through L.E.: 0.305\n\nTABLE XV\nORDINATES OF THE\nNACA 23015 AIRFOIL SECTION\n[Stations and ordinates given in\npercent of airfoil chord]\n\n| Upper surface | | Lower surface | |\n| :--- | :--- | :--- | :--- |\n| Station | Ordinate | Station | Ordinate |\n| 0 | ----- | 0 | 0 |\n| 1.25 | 3.34 | 1.25 | -1.54 |\n| 2.5 | 4.48 | 2.5 | -2.25 |\n| 5 | 6.09 | 5 | -2.94 |\n| 7.5 | 7.20 | 7.5 | -3.61 |\n| 10 | 7.94 | 10 | -4.09 |\n| 15 | 8.92 | 15 | -4.74 |\n| 20 | 9.32 | 20 | -5.11 |\n| 25 | 9.08 | 25 | -5.30 |\n| 30 | 9.05 | 30 | -5.36 |\n| 40 | 8.39 | 40 | -5.32 |\n| 50 | 7.34 | 50 | -5.10 |\n| 60 | 6.21 | 60 | -4.81 |\n| 70 | 4.83 | 70 | -4.31 |\n| 80 | 3.35 | 80 | -2.85 |\n| 90 | 1.74 | 90 | -1.55 |\n| 95 | 1.12 | 95 | -.90 |\n| 100 | (.14) | 100 | (-.14) |\n| 100 | ----- | 100 | 0 |\n\nL.E. radius: 2.48\nSlope of radius through L.E.: 0.305", "timestamp": "2026-07-22T05:07:21.588342+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 48, "total_pages": 66, "image_filename": "19930082245_p48.jpg", "text": "```markdown\n14\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\nSection angle of attack, $\\alpha$, deg\n.28\n.24\n.20\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\nSection pitching-moment coefficient, $c_m$\nNACA TN No. 1596\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n(f) $c_n = 0.6$.\nFigure 9. - Continued.\n47\n[Figure: NACA logo]\n```", "timestamp": "2026-07-22T05:07:24.738731+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 22, "total_pages": 28, "image_filename": "19930085471_p22.jpg", "text": "20\nNACA RM No. L8J11\n\n[Stamp: UNCLASSIFIED]\n[Stamp: CONFIDENTIAL]\n[Stamp: RESTRICTED]\n\n$$V/b\\alpha_{\\alpha} \\text{ exp}$$\n\n[Figure: Scatter plot with diagonal line. Data points are plotted with various symbols. A legend is provided on the right side.]\n\nWing\ndesignation\nA-1\nB-1\nB-2\nB-3\nB-4\nB-5\nC-1\nC-2\nD-1\nE-1\nF-1\nG-1\n\n[Logo: NACA]\n\n$$V/b\\alpha_{\\alpha} \\text{ theor}$$\n\n[Stamp: RESTRICTED]\n[Stamp: CONFIDENTIAL]\n[Stamp: UNCLASSIFIED]\n\nFigure 6.- Comparison of experimental values of $V/b\\alpha_{\\alpha}$ to theoretical values of $V/b\\alpha_{\\alpha}$ at M = 1.31.", "timestamp": "2026-07-22T05:07:25.760987+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 2, "total_pages": 29, "image_filename": "19930085879_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:07:32.868742+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 4, "total_pages": 36, "image_filename": "19930085869_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM L9D15\n\nA model representative of a full-scale flying boat embodying the swept hull, was tested in Langley tank no. 2. The gross weight of the assumed flying boat was 65,000 pounds but its volume was 60 percent less than that of the Boeing XPBB-1, a conventional flying boat of the same gross weight. The results of the tank tests of the swept hull are given in this paper.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| $C_{\\Delta_o}$ | gross load coefficient $(\\Delta_o/wb^3)$ |\n| $C_{\\Delta}$ | load coefficient $(\\Delta/wb^3)$ |\n| $C_V$ | speed coefficient $(V/\\sqrt{gb})$ |\n| $C_R$ | resistance coefficient $(R/wb^3)$ |\n| $\\Delta/R$ | load-resistance ratio |\n| $\\Delta$ | load on water, pounds |\n| $\\Delta_o$ | gross load, pounds |\n| $R$ | resistance, pounds |\n| $V$ | speed, feet per second |\n| $g$ | acceleration of gravity, feet per second per second |\n| $b$ | maximum beam of hull (6.43 ft, full-size) |\n| $w$ | specific weight of water (63.0 lb/cu ft in these tests) |\n| $\\bar{c}$ | mean aerodynamic chord |\n| $\\tau$ | trim measured between forebody keel and horizontal, degrees |\n| $\\delta_e$ | elevator deflection, degrees |\n\nMODEL AND APPARATUS\n\nA powered dynamic model of the swept-hull configuration, designated Langley tank model 237-6SB, was used for the tank tests. Photographs of the model are shown in figure 1. The general arrangement and hull lines\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:07:33.508643+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 20, "total_pages": 46, "image_filename": "19930085519_p20.jpg", "text": "NACA RM No. L8K19\n\n19\n\n[Figure: A black-and-white photograph showing a large, white, swept-back aircraft wing mounted inside what appears to be a wind tunnel or test facility. The wing is angled upward and to the right, with visible structural seams and small vertical markers along its surface. In the background, dark walls and ceiling structures are visible, including some pipes or conduits on the left. In the bottom-right corner of the image, there is a small rectangular label with “NACA” above “I-54921”.]\n\n(b) Wing upper surface.\n\nFigure 2.- Concluded.", "timestamp": "2026-07-22T05:07:36.598251+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 35, "total_pages": 50, "image_filename": "19930082496_p35.jpg", "text": "34\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:07:37.425561+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 11, "total_pages": 46, "image_filename": "19930085542_p11.jpg", "text": "2E NACA RM No. L8L29 9\n\neffective dihedral parameter $C_{l_\\psi}$ are obtained as the aspect ratio is decreased. (See fig. 12.) The rate of change of $C_{l_\\psi}$ with $C_L$ increases as the aspect ratio decreases. (See fig. 21.) Although the values obtained from the theories of references 2 and 5 predict the trends, they are low in magnitude.\n\nThe slopes of the curves of $C_{Y_p}$ in figure 13 were taken at $C_L = 0$ and are presented in figure 22. The curve from the theory of reference 2 shows good agreement with experiment down to an aspect ratio of 2.31 after which the experimental curve falls to zero at $A = 1.07$. The curve from the theory of reference 5 (derived for untapered swept wings) gives consistently high values of $\\partial C_{Y_p}/\\partial C_L$. The theories of references 2 and 5 predict negative values of $C_{n_p}$ at positive lift coefficients; however, the experimental values of $C_{n_p}$ are negative only at moderate lift coefficients. (See fig. 13.) The available theory is, therefore, extremely limited in the range of applicability to triangular plan forms. As the aspect ratio is decreased, $\\partial C_{n_p}/\\partial C_L$ at $C_L = 0$ increases negatively and the theory of reference 5 predicts the results with fair accuracy. (See fig. 22.) The curve from the theory of reference 2 indicates the proper trend, but the correlation with experiment is good only at the lowest test aspect ratio $A = 1.07$.\n\nThe results of reference 8 for flat-plate triangles indicate positive values of $C_{l_p}$ at moderate lift coefficients for aspect ratios below the range considered herein. The opposite trend is noted herein for model 4 ($A = 1.07$) which shows an increase in $C_{l_p}$ starting at a lift coefficient of about 0.1. (See fig. 13.) This increase is believed to be caused by the sudden increase in $C_{L_\\alpha}$ (fig. 11) and the vertical-fin effect at the wing nose resulting from the use of a blunt-nose airfoil section in combination with a large sweep angle. At high angles of attack the vertical displacement of the nose from the axis of rotation causes an increase in damping in roll. The values of $C_{l_p}$ (taken at $C_L = 0$ from fig. 13) presented in figure 22 show an almost linear decrease in damping as the aspect ratio is reduced. The theory of reference 4 is in good agreement with the experimental curve. The theory of reference 5 predicts the trend of the experimental curve but the magnitude is about 15 percent too high. The theory of reference 2 shows fair agreement with experiment only at the lowest test aspect ratio ($A = 1.07$).", "timestamp": "2026-07-22T05:07:37.590299+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 37, "total_pages": 37, "image_filename": "19930082450_p37.jpg", "text": "36\nNACA TN No. 1778\n\n$$ \\frac{H}{t_w} $$\n46\n41\n36\n31\n26\n35\n\n$$ \\bar{\\sigma}_f, \\text{ ksi} $$\nor\n$$ \\sigma_{cr}, \\text{ ksi} $$\n30\n25\n20\n15\n\n$$ \\bar{\\sigma}_f $$\n$$ \\sigma_{cr} $$\n\n25\n30\n35\n40\n\n$$ \\frac{S}{t_s} $$\n\n[Figure: NACA logo]\n\nFigure 10.-Plot for obtaining design from design charts.", "timestamp": "2026-07-22T05:07:41.999295+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 5, "total_pages": 24, "image_filename": "19930085626_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM No. L8K23\n\ndifferentiation of the curve of flight-path velocity against time. The\nscale of the tests is indicated by the curve of Reynolds number against\nMach number shown in figure 5. A more complete description of the\ntechnique is given in references 3 and 4.\n\nACCURACY\n\nThe accuracy of the test results is estimated to be within the\nfollowing limits:\n\n$$\n\\begin{array}{lr}\n\\frac{pb}{2V} \\text{ (due to limitations on model constructional accuracy)} & \\pm 0.005 \\\\\n\\frac{pb}{2V} \\text{ (due to limitations on instrumentation)} & \\pm 0.0005 \\\\\nC_D & \\pm 0.002 \\\\\nM & \\pm 0.005\n\\end{array}\n$$\n\nIn figure 6 is shown the effect of the moment of inertia about the\nroll axis on the measured variation of $pb/2V$ with Mach number. The\ncorrection was made by the method described in reference 3 using an arbi-\ntrarily estimated value of -0.2 for the damping-in-roll derivative over\nthe entire Mach number range. The value of -0.2 is probably very approxi-\nmate; it was simply chosen to show that the magnitude of the correction\nis small for any reasonable negative value of the damping-in-roll deriva-\ntive. The data presented herein have not been corrected for inertia\neffects.\n\nRESULTS AND DISCUSSION\n\nThe results of the present investigation are shown in figure 4 as\ncurves of the wing-tip helix angle $pb/2V$ and total-drag coefficient $C_D$\nagainst Mach number. In each part of figure 4 is shown a drawing of the\nparticular wing-aileron configuration for which the experimental results\nare presented.\n\nTrue-contour ailerons.— The experimental results for the outboard,\ninboard, and full-span true-contour ailerons are shown in figures 4(a),\n4(b), and 4(c), respectively, and are summarized in figure 7. The results\nshown in figure 4(a) are from reference 1 and were obtained with two pairs\nof models. The pairs differed nominally only in the aileron deflection.\n\nThe effectiveness of the outboard ailerons was reversed in the Mach\nnumber range from about 0.94 to 1.0 for $\\delta_a = 5^\\circ$; no reversal was obtained\nwith $\\delta_a = 10^\\circ$. The effectiveness of the inboard ailerons also reversed\nbut at slightly lower Mach numbers than for the outboard ailerons. No\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:07:42.258508+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 49, "total_pages": 66, "image_filename": "19930082245_p49.jpg", "text": "```markdown\n16\n14\n12\n10\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nSection angle of attack, $\\alpha$, deg\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\n\n.28\n.24\n.20\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\nSection pitching-moment coefficient, $c_m$\n$\\delta_a$\n(deg)\n-12\n-6\n-4\n-2\n0\n2\n4\n6\n12\n18\n30\nNACA\n\n(g) $c_n=0.7$.\nFigure 9.—Continued.\n\n48\nNACA TN No. 1596\n```", "timestamp": "2026-07-22T05:07:47.746671+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 24, "total_pages": 78, "image_filename": "19930082618_p24.jpg", "text": "```markdown\n22\n\n[Figure: A graph plotting aerodynamic characteristics. The top plot shows Section lift coefficient ($c_l$) vs Section angle of attack ($\\alpha_0$). The bottom plot shows Moment coefficients ($c_{m_{c/4}}$) vs Section angle of attack ($\\alpha_0$). An airfoil cross-section is shown in the top left corner.]\n\nSection lift coefficient, $c_l$\nMoment coefficients, $c_{m_{c/4}}$\nSection angle of attack, $\\alpha_0$, deg\n\nR\n0.7 x $10^6$\n1.0\n2.0\n3.0\n4.0\n6.0\n9.0\n\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(a) Section lift and pitching-moment characteristics of the plain airfoil section.\nFigure 1.- Aerodynamic characteristics of the NACA 64-409 airfoil section, 24-inch chord.\n\nNACA TN 1945\n```", "timestamp": "2026-07-22T05:07:48.299260+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 25, "total_pages": 66, "image_filename": "19930082914_p25.jpg", "text": "```markdown\n24\nNACA TN No. 1857\n\ndensity to atmospheric density $\\rho/\\rho_{atm}$. On the horizontal axis is plotted the nondimensional parameter $\\sigma y/x$. The variables $y$ and $x$ are position coordinates. The $y$-axis is vertical. The $x$-axis is not quite horizontal but has been so chosen that it coincides with the line along which the density ratio is 1.1. This is the line along which the velocity is 0.5 free-stream velocity. (It has been customary in the past to place the $x$-axis along the 0.5-velocity-ratio contour. For the jet under discussion, this contour was at an angle of $-1\\frac{1}{4}^\\circ$ to the horizontal.) The parameter $\\sigma$ is an experimentally determined scale factor. Its value is obtained by comparing the experimentally determined velocity distribution with the theoretical distribution. The parameter $\\sigma$ will be discussed in a subsequent section. Figure 15 shows that there is fairly good similarity in the density distributions. The lack of scatter of the experimental points at the outside portion of the mixing region is explained by the fact that atmospheric density was used as the reference in the quantity $\\rho/\\rho_{atm}$ that was plotted. The scatter that occurs at the inside part of the mixing region can be attributed partly to variations in atmospheric density and partly to variations in the stagnation temperature of the jet, both of which varied from picture to picture.\n\n### Velocity Distribution\n\nFrom the density distributions the velocity distributions were calculated with the aid of several assumptions and approximations. It was assumed that the static pressure in the jet and the mixing region was the same as the pressure in the room air outside the jet. The temperature distributions through the jet were then obtained from the density distributions by the general gas law.\n\nFor the different interferograms the room-air temperature varied between $79.6^\\circ$ F and $80.1^\\circ$ F. The stagnation temperature of the jet air varied between $71.6^\\circ$ F and $73.6^\\circ$ F. (The stagnation temperature was measured with a thermocouple that was installed in the 6-inch pipe ahead of the nozzle where the air velocity was low, about 150 feet per second, and the temperature recovery factor was very nearly equal to unity.) The assumption was made that there was no heat transfer in the mixing region. The jet and the mixing region were therefore considered to be isothermal from the standpoint of stagnation temperature. For the calculations of velocity distribution, the assumption was made that the stagnation temperature in the mixing zone was the same as the temperature of the room air. Then, from the constant stagnation temperature and the static temperature distributions, the velocity distributions were calculated from the conservation-of-energy equation\n\n$$u^2 = 2c_p(T_{stag} - T)$$\n\nwhere $u$ is the velocity and $c_p$ is the specific heat at constant pressure.\n```", "timestamp": "2026-07-22T05:07:49.536484+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 23, "total_pages": 28, "image_filename": "19930085471_p23.jpg", "text": "NACA RM No. L8J11\n21\n\n[Figure: A scatter plot comparing experimental and theoretical values. The x-axis is labeled $\\omega_p/\\omega_\\alpha$ theor and ranges from 0 to 1.2. The y-axis is labeled $\\omega_p/\\omega_\\alpha$ exp and ranges from 0 to 1.2. A diagonal line represents perfect agreement. A legend box titled \"Wing designation\" lists symbols for A-1 through G-1. There are \"UNCLASSIFIED\" and \"CONFIDENTIAL\" stamps on the chart.]\n\nFigure 7.- Comparison of experimental values of $\\omega_p/\\omega_\\alpha$ with theoretical values of $\\omega_p/\\omega_\\alpha$ at M = 1.31.", "timestamp": "2026-07-22T05:07:52.135198+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 39, "total_pages": 99, "image_filename": "19930082511_p39.jpg", "text": "NACA TN No. 1826\n37\n\nAt the edge of the exit lip, where $x = a$, this expression reduces to\n\n$$\n- \\frac{2iK'}{\\pi^2} \\left( E' - \\frac{K'}{a^2} \\right)\n$$\n\nNUMERICAL RESULTS\n\nIn the following sections are described some numerical results that were computed by the preceding equations in order to show the magnitudes of the entrance and exit effects. It will be noted that, since the complex velocity has been made nondimensional by dividing by $V$, the component $v$ is identical with the tunnel-induced angle $\\epsilon$, in radians, and the component $u$ is the fractional increase in the horizontal velocity. The equivalence of the two ordinate scales indicated in the plots of the results follows from the equation\n\n$$\n\\frac{\\Gamma}{V} = \\frac{ccl}{2}\n$$\n\nClosed-open tunnel.- In figure 17 are shown calculated values of the induced downwash angle along the tunnel axis for various positions of the lifting vortex along the axis. The figure shows that for $\\xi_1 = 1.0$ and $1.5$, the induced angles at the vortex itself ($\\xi = 1.0$ and $1.5$, respectively) are almost exactly $\\frac{\\Gamma'}{2}$, which is the value for an infinitely long open tunnel; and, furthermore, the two curves are symmetrical about the point $\\xi = \\xi_1$. In fact, within the accuracy of the plot, these two curves are identical with the curve for an infinitely long open tunnel. It may be concluded that the closed entrance has no effect if the vortex is more than one tunnel height from the entrance. For $\\xi_1 = 0.5$, which is a more likely location of the wing, the induced velocity at $\\xi = \\xi_1$ is $0.48\\Gamma'$, and the curve is no longer exactly symmetrical about the point $\\xi = \\xi_1$; however, these differences from the conditions for the infinitely long open tunnel are too small to be practically significant, so that the usual infinite-open-tunnel theory is still adequate for $\\xi_1 = 0.5$. For $\\xi_1$ less than $0.5$, the deviations from infinite-open-tunnel theory become larger rapidly, until, when the vortex is in the plane of the entrance ($\\xi_1 = 0$), the induced angle at the point $\\xi = \\xi_1$ is only $\\frac{\\Gamma'}{4}$.\n\nA similar discussion applies for the vortex in the closed portion of the tunnel ($\\xi_1 < 0$), although this case normally has no practical significance. For $\\xi_1 = -1$, the induced angles in the neighborhood of the vortex are practically identical with those for an infinitely long closed tunnel; however, in the open region ($\\xi > 0$), the curve is considerably different from that for the infinitely long closed tunnel (shown as the dashed curve in fig. 17).", "timestamp": "2026-07-22T05:07:54.380093+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 24, "total_pages": 36, "image_filename": "19930082614_p24.jpg", "text": "22\nNACA TN 1939\n\nTABLE III.— CALCULATION OF VARIATIONS WITH TIME OF AIRSPEED AND FLIGHT-PATH ANGLE FOR AN AIRPLANE ENTERING A 60° DIVE. WING LOADING, 50 POUNDS PER SQUARE FOOT. INITIAL ALTITUDE, 25,000 FEET\n\n| t (sec) | $\\Delta$t (sec) | $V_a$ (ft/sec) | $\\bar{V}_a$ (ft/sec) | n | $\\bar{\\gamma}$ (deg) | $\\Delta\\gamma$ (deg) | $\\gamma$ (deg) | $\\Delta$h (ft) | h (ft) | $t_a$ (lb/sq ft) | $C_L$ | $C_{D_0}$ | n/g | n (ft/sec²) | $\\bar{a}$ (ft/sec²) | $\\Delta V$ (ft/sec) | V (ft/sec) |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 0 | 0 | -- | -- | 1 | -- | 0 | 0 | -- | 25,000 | 261 | 0.19 | 0.019 | +0.08 | -2.6 | -- | -- | 700 |\n| 1 | 1 | 689 | 695 | -1.5 | -3.2 | -6.5 | -6.5 | 6.0 | -30 | 24,970 | 253 | -.30 | .118 | +.20 | +16.2 | -9.2 | -9.2 | 691 |\n| 1 | 1 | 691 | 686 | -1.5 | -7.3 | -6.6 | -6.6 | 6.0 | -30 | 24,970 | 255 | -.29 | .118 | +.48 | +15.5 | -8.9 | -8.9 | 691 |\n| 2 | 1 | 679 | 685 | -1.5 | -9.9 | -6.7 | -13.3 | 6.0 | -60 | 24,910 | 246 | -.30 | .118 | +.35 | -11.3 | -13.4 | -13.4 | 678 |\n| 3 | 1 | 686 | 676 | -1.5 | -16.6 | -6.6 | -19.9 | 5.9 | -100 | 24,730 | 241 | -.31 | .118 | +.20 | -7.1 | -9.2 | -9.2 | 669 |\n| 4 | 1 | 663 | 666 | -1.5 | -23.2 | -6.6 | -26.5 | 5.7 | -160 | 24,470 | 235 | -.32 | .118 | +.10 | -3.2 | -5.1 | -5.1 | 664 |\n| 6 | 2 | 664 | 664 | -1.5 | -23.0 | -13.0 | -39.5 | 5.3 | -270 | 23,900 | 245 | -.30 | .118 | .06 | 1.9 | -0.6 | -1.2 | 663 |\n| 8 | 2 | 670 | 666 | -1.5 | -45.7 | -12.0 | -51.5 | 6.1 | -770 | 23,130 | 256 | -.29 | .118 | .20 | 6.4 | 4.1 | 8.2 | 671 |\n| 9.5 | 1.5 | 680 | 676 | -1.5 | -55.7 | -4.5 | -60.0 | 6.7 | -2440 | 22,290 | 276 | -.08 | .118 | .23 | 7.4 | 6.9 | 10.3 | 681 |\n| 12 | 2.5 | 700 | 690 | 1.5 | -60.0 | 0 | -60.0 | = | 6-1490 | 20,900 | 298 | .08 | .113 | .18 | 5.8 | 6.6 | 16.5 | 698 |\n| 15 | 3 | 712 | 705 | .5 | -60.0 | 0 | -60.0 | = | -1850 | 18,970 | 332 | .08 | .113 | .09 | 2.9 | 4.3 | 12.9 | 711 |\n\n$^a \\Delta\\gamma = -60.0 + 51.5 = -8.5^\\circ$. $\\bar{\\gamma} = -51.5 - \\frac{8.5}{2} = -55.7^\\circ$. $\\Delta t = 1.5$ sec (figs. 5(c) and (d) for $V_a = 670$ ft/sec)\n$^b$Normal acceleration factor for a steady 60° dive, $n = \\cos(-60^\\circ) = 0.5$\n$^c C_D = \\bar{V}_a \\Delta t \\sin(-60^\\circ)$\n\n[Figure: NACA logo]\n\nExample calculation:\n\nAn airplane begins a dive from level flight at 700 feet per second, reaching an indicated normal acceleration of -1.5g during the first second, and maintaining this acceleration until the dive angle is 60°.\n\nNet drag coefficient of the airplane, $C_{D_0} = 0.013 + 0.060 \\ C_L^2 + \\Delta C_D$\nDrag increment due to air brakes, $\\Delta C_D = 0.100$\n\nAt t = 0\n$\\gamma = 0$. Proceeding as in the case of constant $\\gamma$ (table II), n/g = -0.08, n = -2.6 ft/sec²\n\nt = 1 sec\n$\\bar{a}$ is estimated, considering that $C_D = 0.019$ at t = 0 and $C_D = 0.118$ at t = 1.0,\n$\\bar{a}_a = -11$ ft/sec²\n$V_a = 700 - 11 (\\Delta t) = 689$ ft/sec ($\\Delta t = 1$ sec)\nn = -1.5\n$n = \\cos \\gamma = -0.5$, for $\\bar{\\gamma} = 0^\\circ$ (fig. 5(a))\n$\\gamma_a = 1/2 \\ \\Delta\\gamma = 1/2 \\ (6.5) = 3.25^\\circ$ (figs. 5(c) and (d). This may possibly require more than than a single trial in some instances.)\n$\\Delta\\gamma = -6.5^\\circ$ (figs. 5(a), (b), and (c))\nr/1000 = 6.0 (See guide lines, for example 1, figs. 5(a) and (b))\n$\\Delta h = -30$ ft (figs. 5(e) and (f) Minus sign chosen because a dive has been specified)\n$h = 25,000 - 30 = 24,970$ ft\n$t_a = 1/2 \\ \\sigma V_a^2 = 253$ lb/sq ft\n$C_L = \\frac{(-1.5)(50)}{253} = -0.30$\nn/g = -0.50 (fig. 1)\n$V = 700 - \\frac{32.2}{2} (-0.08 - 0.50) (1) = 691$ ft/sec\n\nTo show that the original estimate was close enough to give the correct velocity, the calculations have been repeated in the table with $V_a = 691$ ft/sec. This again gives $V = 691$ ft/sec.", "timestamp": "2026-07-22T05:07:57.274668+00:00"}
{"citation_id": "19930085879", "source_url": "https://ntrs.nasa.gov/api/citations/19930085879/downloads/19930085879.pdf", "page_number": 3, "total_pages": 29, "image_filename": "19930085879_p3.jpg", "text": "NACA RM L9D11\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nFLIGHT INVESTIGATION OF THE JETTISONABLE-NOSE METHOD  \nOF PILOT ESCAPE USING ROCKET-PROPELLED MODELS  \nBy Reginald R. Lundstrom and Burke R. O'Kelly\n\nSUMMARY\n\nTwo rocket-propelled models to test the jettisonable-nose method of pilot escape were launched by the Langley Laboratory. During the flight of the first model the nose came off during power-on flight due to the malfunction of a release latch and was damaged by collision with a wing of the main body. The nose section of the second model was jettisoned successfully at the end of its power-on flight at a Mach number of 0.87. Accelerations produced were well within human tolerance. The drag-weight ratios of nose and rear bodies were such that the deceleration of the nose was less than that of the rear body. The shielding effect of the nose on the rear body during separation was appreciable and forcible separation appears necessary.\n\nINTRODUCTION\n\nAs the speeds of piloted aircraft advance into the transonic and supersonic ranges, conventional means of pilot escape in cases of emergency appear to be inadequate. Ejection seats of the type mentioned in reference 1 should make escape much easier at subsonic speeds, but, in their present form, their use for escape at speeds in excess of 550 miles per hour at moderately low altitudes does not appear practical.\n\nThe human body is quite sensitive to accelerations and in any escape device the accelerations should be kept at a minimum for the safety and comfort of the pilot. Reference 2 lists the physiological effects of acceleration. Because high-speed airplanes are expected to travel at high altitudes, the escape method must provide the pilot with oxygen until a low altitude has been reached.\n\nOne method of escape which appears practical is to jettison the complete nose section of the airplane and, after it has been decelerated to a fairly low-subsonic speed, have the pilot leave the nose section with his own personal parachute. Reference 3 and recent unpublished low-speed", "timestamp": "2026-07-22T05:08:01.991271+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 32, "total_pages": 49, "image_filename": "19930082498_p32.jpg", "text": "NACA TN No. 1838\n\n[Figure: Two metallic, box-like configurations with multiple tubes protruding from the top and a single curved tube exiting from the bottom. The foreground configuration has a label reading \"NACA L-52706\" affixed to its side.]\n\n(b) Configuration 57.\n\nFigure 2.— Continued.\n\n31", "timestamp": "2026-07-22T05:08:02.767536+00:00"}
{"citation_id": "19930085544", "source_url": "https://ntrs.nasa.gov/api/citations/19930085544/downloads/19930085544.pdf", "page_number": 9, "total_pages": 33, "image_filename": "19930085544_p9.jpg", "text": "8\nNACA RM No. L8K26\n\nOscillating flow.- The expression for the total lift of an infinitely thin airfoil of infinite aspect ratio oscillating in an incompressible pulsating stream (see fig. 2) is given in reference 2 as\n\n$$\n\\begin{aligned}\nL = -\\pi\\rho \\frac{c^2}{4} \\Bigg[ & \\ddot{h} + \\dot{W}_{wt}\\dot{\\alpha}_P + \\dot{W}_{wt}(\\alpha + \\alpha_P) - a \\frac{c}{2} \\ddot{\\alpha}_P \\Bigg] \\\\\n& - 2\\pi\\rho W_{wt} \\frac{c}{2} \\Bigg\\{ \\dot{W}_o\\alpha + \\epsilon W_o\\alpha C\\left(k_{W_{wt}}\\right) e^{i\\omega t} \\\\\n& + \\left[ \\frac{c}{2}\\left(\\frac{1}{2} - a\\right)\\dot{\\alpha}_P + \\dot{W}_o\\alpha_P \\right] C\\left(k_{\\alpha_P}\\right) + \\dot{h}C\\left(k_h\\right) \\\\\n& + \\epsilon \\dot{W}_o\\alpha_P C\\left(k_{W_{wt}} + k_{\\alpha_P}\\right) e^{i\\omega t} \\Bigg\\}\n\\end{aligned}\n\\tag{7}\n$$\n\nwhere\n\n$$\nW_{wt} - W_o = W_o\\epsilon e^{i\\omega t}\n$$\n\n$$\nh = h_o e^{i\\omega t}\n$$\n\n$$\n\\alpha_P = \\alpha_{P_o} e^{i\\omega t}\n$$\n\n$$\nC(k) = F + iG \\quad \\text{(from reference 1)}\n$$\n\nIn the above expressions, $\\epsilon$ denotes the fractional amplitude of the perturbation part of the stream pulsations, $h_o$ the amplitude of the vertical displacement (flapping), and $\\alpha_{P_o}$ the amplitude of the incremental angle of attack due to the rotation of the airfoil about the axis $\\overline{x} = a$. The equations defining the quantities $(W_{wt} - W_o)$, $h$, and $\\alpha_P$ describe these quantities as pure sinusoidal variations; it thus becomes necessary to ascertain the applicability of these definitions to the inclined propeller case.", "timestamp": "2026-07-22T05:08:08.847707+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 1, "total_pages": 96, "image_filename": "19930085880_p1.jpg", "text": "NACA RM No. L9C03\n\nRM No. L9C03\n\n[Figure: NACA logo with wings]\n\nRESEARCH MEMORANDUM\n\nPLANING CHARACTERISTICS OF THREE SURFACES  \nREPRESENTATIVE OF HYDRO-SKI FORMS\n\nBy\n\nKenneth L. Wadlin and John R. McGehee\n\nLangley Aeronautical Laboratory  \nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE  \nFOR AERONAUTICS  \nWASHINGTON\n\nMarch 29, 1949  \nDeclassified April 24, 1956", "timestamp": "2026-07-22T05:08:09.377064+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 23, "total_pages": 36, "image_filename": "19930085487_p23.jpg", "text": "NACA RM No. E8J22\n\n[Figure: Sand pattern on turbine blades with scale marked \"INCHES\"]\n\nNACA\nC-14990\n5-20-46\n\nFigure 7. - Sand pattern for second bending-mode vibration on seventh-stage blade showing distorted node shape. Frequency of vibration, 4440 cycles per second.\n\n21", "timestamp": "2026-07-22T05:08:09.554333+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 12, "total_pages": 46, "image_filename": "19930085542_p12.jpg", "text": "10\nNACA RM No. L8L29\n\nEffect of Vertical Fins\n\nThe effects of the fins on $C_L$, $C_m$, and $C_X$ (fig. 14) are small as would be expected.\n\nAlthough the increment in $C_{n_\\psi}$ (fig. 15) caused by the small (A = 0.77) and large (A = 1.15) fins decreases from $C_L = 0$ to $C_L = 1.0$, $C_{n_\\psi}$ for the wing-fin combinations increases throughout the entire lift range. The addition of either fin causes a positive increment of $C_{l_\\psi}$ at $C_L = 0$ which decreases as the lift coefficient increases. Consequently, the rate of change of $C_{l_\\psi}$ with $C_L$ for the wing-fin combinations is smaller than for the wing alone. (See fig. 15.)\n\nAddition of the large fin caused negative displacements of $C_{Y_p}$ at $C_L = 0$; the addition of either fin caused an increase in the rate of change of $C_{Y_p}$ with $C_L$. (See fig. 16.) An almost constant positive increment in $C_{n_p}$ is the result of adding the large fin to model 2. (See fig. 16.) Both the large and small fins increased the damping in roll throughout the lift-coefficient range; the increase for the large fin amounts to about 30 percent of the damping in roll of the wing alone.\n\nEffect of Aspect Ratio of Modified Triangular Plan Forms\n\nThe models of the present group were formed by cutting various portions from the tips of a basic triangular wing (model 7) parallel to the plane of symmetry to obtain aspect ratio 3 (model 8), aspect ratio 2 (model 9), and aspect ratio 1 (model 10), including tips of revolution.\n\nFor the lowest test aspect ratio (A = 1.0) a definite nonlinear variation of $C_L$ with $\\alpha$ is noted in figure 17 which agrees with the results of similar models tested in reference 8. Except for a short range ($C_L = 0.15$ to $C_L = 0.3$) the variation of $C_m$ with $C_L$ is nonlinear for model 10 (A = 1.0). Reducing the aspect ratio of modified triangular wings results in a forward movement of the aerodynamic center which can be estimated with fair accuracy for aspect ratios of 2.3 or less by the theory of reference 3. (See fig. 23.) An increase in $C_{L_{max}}$ is noted as the aspect ratio is decreased. A similar trend was noted for the same range of aspect ratios reported in reference 8. The theories of references 3 and 5 predict a decrease in lift-curve slope with a decrease in aspect ratio but are only in fair agreement with the experimental results. (See fig. 23.) The theory of reference 2 shows poor agreement with experiment even for an aspect ratio of 1.0.", "timestamp": "2026-07-22T05:08:12.311530+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 46, "total_pages": 62, "image_filename": "19930082485_p46.jpg", "text": "1026\n\nNACA TN No. 1810\n\n| Ratio of specific heats, $\\gamma$ | | | | Ratio of specific heats, $\\gamma$ |\n| :--- | :--- | :--- | :--- | :--- |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | 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| | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n| | | | | |\n|", "timestamp": "2026-07-22T05:08:14.302250+00:00"}
{"citation_id": "19930085626", "source_url": "https://ntrs.nasa.gov/api/citations/19930085626/downloads/19930085626.pdf", "page_number": 6, "total_pages": 24, "image_filename": "19930085626_p6.jpg", "text": "NACA RM No. L8K23 CONFIDENTIAL 5\n\nreversal was obtained for the full-span ailerons, and except for the Mach number range for which reversal was obtained for the partial-span ailerons the sum of the measured values of effectiveness for the inboard and outboard partial-span ailerons was approximately equal to the effectiveness measured for the full-span ailerons.\n\nExtended-chord ailerons.- The results obtained for the extended-chord aileron deflected $5^0$ are shown in figure 4(d). The rolling effectiveness of this configuration was higher than that obtained for the original ailerons over the Mach number range investigated and no reversal was obtained. A small jog in the effectiveness curve, similar to that obtained for the true-contour ailerons, was encountered just below a Mach number of 1.\n\nBlunt trailing-edge ailerons.- The experimental results for the blunt trailing-edge ailerons, deflected $5^0$, are shown in figure 4(e). The effectiveness of the parallel-side ailerons varied smoothly and continuously over the entire Mach number range investigated. A small loss in effectiveness just below a Mach number of 1 was obtained for the flat-side aileron having a trailing-edge thickness equal to one-half of that of the parallel-side aileron. At the extremities of the investigated Mach number range, the effectiveness for both blunt trailing-edge ailerons was only slightly less than that of the original true-contour ailerons. At supersonic velocities the drag of the blunt trailing-edge-aileron configurations was only slightly larger than that of the original ailerons; at subsonic velocities a measurable increase in drag was obtained for the blunt trailing-edge configurations. The results obtained for the blunt trailing-edge ailerons indicate that separation of the boundary layer over the rearward part of the airfoil section was partly responsible for the reversal of effectiveness obtained for the true-contour ailerons.\n\nThe results obtained in the present investigation are in agreement with those of reference 2 as regards the relative effectiveness of the various configurations tested.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Field, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:08:14.772226+00:00"}
{"citation_id": "19930085869", "source_url": "https://ntrs.nasa.gov/api/citations/19930085869/downloads/19930085869.pdf", "page_number": 5, "total_pages": 36, "image_filename": "19930085869_p5.jpg", "text": "NACA RM L9D15 CONFIDENTIAL 3\n\nare presented in figures 2 and 3, respectively. The offsets of the hull are given in table I.\n\nThe basic plan-form section of the forebody (modified 16 series symmetrical airfoil section with a thickness ratio of 14.3 percent) was the same as that used for the unswepf forebodies of the models in reference 2. However, the water planes above the chine were progressively shifted aft, producing a swept profile as shown in figure 3. The length of the bow was decreased, but the volume was kept about the same by this manner of sweeping the hull. The afterbody was a simple conical boom.\n\nThe forebody chine was made straight in profile resulting in a continuous variation in deadrise angle and slightly convex buttock lines near the step. Vertical spray strips were installed at the chine to reduce propeller spray. Spray tests were first made with the spray strip shown in figure 4(a) which has the same depth as that used in reference 2 on the unswepf model. Since this spray strip allowed heavy spray to reach the propellers, the one shown in figure 4(b) was developed and used throughout the rest of the tests.\n\nThe configuration was a $\\frac{1}{16}$-scale model representing an assumed flying boat of 65,000 pounds gross load ($C_{\\Delta_0} = 3.87$). The wing loading and power loading of the Navy XPBB-1 flying boat (35.6 lb/sq ft and 14.8 lb/bhp) were simulated on the model. The size and locations of the aerodynamic surfaces corresponded to those of the XPBB-1. No flaps were used. The lateral spacing of the nacelles was the same as that of the twin-boom configuration described in reference 2. Leading-edge slats were installed to compensate for the low Reynolds number of the tests. The elevators had a range of deflection from -30° to 20°.\n\nThe test setup is shown in figure 5 with the model under way at a speed coefficient of 8.3. The model was free to trim about the pivot, which was located at the center of gravity and was free to move vertically, but was restrained in roll and yaw.\n\nPROCEDURE\n\nSpray\n\nThe range of speed over which spray was in the propellers was determined by making constant speed runs at full power and a series of gross loads. The model was free to trim about the 0.30$\\bar{c}$ location of the center of gravity with the elevators fixed at 0°. Observations were made of bow spray and spray that struck the wing and tail surfaces.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:08:15.251311+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 50, "total_pages": 66, "image_filename": "19930082245_p50.jpg", "text": "```markdown\nNACA TN No. 1596\n\nSection angle of attack, $\\alpha$, deg\n16\n14\n12\n10\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\n$\\delta_d$\n(deg)\n4\n6\n12\n18\n30\n\nSection pitching-moment coefficient, $c_m$\n.28\n.24\n.20\n.16\n.12\n.08\n.04\n0\n-.04\n-.08\n-.12\n-.16\n-.20\n.1 .2 .3 .4 .5 .6 .7 .8 .9\nMach number, M\n\n$\\delta_d$\n(deg)\n4\n6\n12\n18\n30\n\nNACA\n\n(h) $c_n = 0.8$.\nFigure 9.- Concluded.\n\n49\n```", "timestamp": "2026-07-22T05:08:21.571941+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 24, "total_pages": 28, "image_filename": "19930085471_p24.jpg", "text": "22\nNACA RM No. L9J11\n\nUNCLASSIFIED\nRECLASSIFIED\nCONFIDENTIAL\n\n<!-- Image (152, 174, 850, 747) -->\n\nFigure 8.- Comparison of experimental values of $V/b\\omega_\\alpha$ at M = 1.3\nwith theoretical incompressible values of $(V/b\\omega_\\alpha)_0$ at M = 0.", "timestamp": "2026-07-22T05:08:21.784968+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 25, "total_pages": 78, "image_filename": "19930082618_p25.jpg", "text": "NACA TN 1945\n\nSection lift coefficient, $c_l$\nMoment coefficient, $c_{m_{c/4}}$\nSection angle of attack, $\\alpha_s$, deg\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\square$ 1.0\n$\\diamond$ 2.0\n$\\triangle$ 3.0\n$\\nabla$ 6.0\nFlagged symbols denote standard roughness\n\n[Figure: Graph showing section lift coefficient and moment coefficient versus section angle of attack for various Reynolds numbers (R), with flagged symbols indicating standard roughness. The NACA logo is present on the graph.]\n\n(b) Section lift and pitching-moment characteristics of the NACA 64-409 airfoil section with a 0.20c simulated split flap deflected 60°.\n\nFigure 1.— Continued.\n\n23", "timestamp": "2026-07-22T05:08:22.611153+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 21, "total_pages": 46, "image_filename": "19930085519_p21.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:08:26.858587+00:00"}
{"citation_id": "19930082496", "source_url": "https://ntrs.nasa.gov/api/citations/19930082496/downloads/19930082496.pdf", "page_number": 36, "total_pages": 50, "image_filename": "19930082496_p36.jpg", "text": "NACA TN No. 1836\n35\n\n[Figure: Fractured surface of ceramal tensile specimen 3D5. Tensile strength, 34,600 pounds per square inch at 1800° F.]\n\n1 INCH\n\nNACA\nC-20939\n3-22-48\n\nFigure 8. - Fractured surface of ceramal tensile specimen 3D5. Tensile strength, 34,600 pounds per square inch at 1800° F.", "timestamp": "2026-07-22T05:08:27.340151+00:00"}