PDF/ocr_dataset_1014.jsonl

{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 24, "total_pages": 34, "image_filename": "19930086151_p24.jpg", "text": "```markdown\nCONFIDENTIAL\n\n22\n\nYawing-moment coefficient, $C_n$\nRolling-moment coefficient, $C_l$\n\n$\\delta_a$ (deg)\n$\\square$ 2.1\n$\\diamond$ 4.2\n$\\triangle$ 6.0\n$\\square$ 14.7\n$\\triangle$ 19.5\n$\\square$ 29.5\n\nAngle of attack, $\\alpha$, deg\n\nCONFIDENTIAL\n\nNACA\n\nNACA RM L9D28\n\nFigure 6.— The rolling-moment and yawing-moment characteristics of the 45° sweptback wing for various deflections of the parallelogram wing-tip aileron. Plain wing.\n```", "timestamp": "2026-07-22T04:36:53.958519+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 93, "total_pages": 118, "image_filename": "19930085838_p93.jpg", "text": "NACA RM No. L9B23\n91\n\n<!-- Image (165, 113, 893, 734) -->\n\n(h) $\\delta_F = 40^\\circ$.\nFigure 11.- Continued.", "timestamp": "2026-07-22T04:36:54.209915+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 13, "total_pages": 47, "image_filename": "19930093773_p13.jpg", "text": "```markdown\n12\nNACA RM E9G09\n\nMethods of Calculation\n\nFlight Mach number. - Complete ram-pressure recovery at the\nengine inlet was assumed. The flight Mach number was then\ndetermined from the following expression:\n\n$$\nM_0 = \\sqrt{\\frac{2}{\\gamma-1} \\left[ \\left(\\frac{P_1}{P_0}\\right)^{\\frac{\\gamma-1}{\\gamma}} - 1 \\right]} \\quad (1)\n$$\n\nTemperatures. - Total temperature was obtained from indicated\ntemperature by the use of an experimentally determined thermo-\ncouple impact-recovery factor of 0.85 in the following equation:\n\n$$\nT = \\frac{T_1 \\left(\\frac{P}{P_1}\\right)^{\\frac{\\gamma-1}{\\gamma}}}{1 + 0.85 \\left[ \\left(\\frac{P}{P_1}\\right)^{\\frac{\\gamma-1}{\\gamma}} - 1 \\right]} \\quad (2)\n$$\n\nEquivalent temperature. - Equivalent temperature was obtained\nfrom tunnel static pressure and engine-inlet total pressure and\ntemperature.\n\n$$\nt_e = \\frac{T_1}{\\left(\\frac{P_1}{P_0}\\right)^{\\frac{\\gamma-1}{\\gamma}}} \\quad (3)\n$$\n\nAir flow. - Engine air flow was calculated from pressure and\ntemperature measurements obtained at the engine inlet (station 1)\nby use of the equation\n\n$$\nW_a = A_1 P_1 \\sqrt{\\frac{2\\gamma g}{t_1 R(\\gamma-1)} \\left[ \\left(\\frac{P_1}{P_1}\\right)^{\\frac{\\gamma-1}{\\gamma}} - 1 \\right]} \\quad (4)\n$$\n\n1169\n```", "timestamp": "2026-07-22T04:36:55.494611+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 32, "total_pages": 44, "image_filename": "19930086081_p32.jpg", "text": "30\nNACA RM L9H05\n\nCONFIDENTIAL\n\n<!-- Image (93, 163, 852, 768) -->\n\nFigure 12.- Effect of tip thickness, airfoil nose radius, and airfoil contour on the lift-drag polar characteristics of a semispan delta wing. Small fuselage, fence off. R = 4.0 X $10^6$; M = 1.90.", "timestamp": "2026-07-22T04:37:01.504373+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 10, "total_pages": 21, "image_filename": "19930092013_p10.jpg", "text": "6\nREPORT 948—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (117, 78, 917, 522) -->\n\nFIGURE 6.—Lateral control characteristics in steady straight sideslip. $V_T=300$ knots.\n\nThe following formula was used in computing the values of stick fixed $C_{l\\beta}$ noted in figure 6\n\n$$\nC_{l\\beta}=C_{l\\beta}\\left[\\frac{d(pb/2V)}{d\\delta_a}\\right]\\left(\\frac{\\partial\\delta_a}{\\partial\\beta}\\right) \\quad (2)\n$$\n\nwhere a value of $C_{l\\beta}$ of $-0.45$ was obtained from reference 6. Abrupt rudder-fixed aileron-roll flight tests gave a value of $-0.00215$ for $d(pb/2V)/d\\delta_a$ (corrected approximately for effects of sideslip), and $\\partial\\delta_a/\\partial\\beta$ is the pilot-applied curve slope from figure 6. Also shown for comparison are values of the stick-fixed effective dihedral angle $\\Gamma_e$ computed from the equation\n\n$$\n\\Gamma_e=\\frac{C_{l\\beta}}{C_{l\\beta}/\\Gamma_e} \\quad (3)\n$$\n\nwhere a value of $C_{l\\beta}/\\Gamma_e$ of $-0.000225$ per degree squared was obtained from reference 6.\n\n**Abrupt rudder kicks.**—The pilot abruptly deflected and held the rudder pedals while the stick was held fixed. Several different rudder deflections, left and right, were employed for each servo-gearing ratio. Typical time histories showing the motions of the control surfaces and airplane are presented in figure 7. The computed aileron deflections due to servo action for an ideal servomechanism (no time lag) are shown for comparison with the measured values. The variation of the maximum value of the rolling parameter $pb/2V$ with change in rudder deflection $\\Delta\\delta_r$ for these maneuvers is given in figure 8 (a). The ratio of $pb/2V$ for unit $\\Delta\\delta_r$ for each servo-gearing ratio to the value for the normal airplane was computed from the slopes of these curves and is shown in figure 8 (b) as a function of the corresponding static stick-fixed effective dihedral angle. A similar predicted curve computed by the method of reference 7 is shown for comparison.\n\n**Controls-fixed lateral oscillations.**—Lateral oscillations were induced from an initial steady-sideslip attitude by abruptly returning and holding the rudder pedals and control stick in trim position. Typical time histories of the control deflections, including the computed aileron motion supplied by an ideal servomechanism, and resultant airplane motions are given in figure 9. The absolute values of servo-applied aileron deflection shown in figure 9 are probably in error because of initial misalignment at the trim sideslip angle. This factor is not important, however, because only the time variation of this quantity has any significance here. The oscillation period $P$ and number of cycles to damp to one-half amplitude $C_{1/2}$ were determined from the time histories of sideslip angle, and are plotted in figure 10 as a function of effective dihedral angle $\\Gamma_e$. Values", "timestamp": "2026-07-22T04:37:03.804416+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 84, "total_pages": 92, "image_filename": "19930085930_p84.jpg", "text": "CONFIDENTIAL\nNACA RM L9507\n\n[Figure: A black and white photograph showing a large, curved, metallic structure, possibly part of an aircraft or wind tunnel model. The structure has a smooth, aerodynamic shape with visible seams and joints. At the bottom, there is a smaller component with several vertical fins or blades.]\n\n(b) Passage corresponding to characteristics net shown in figure 10(b).\n\nFigure 29 - Concluded.\n\nCONFIDENTIAL\n\n83", "timestamp": "2026-07-22T04:37:06.405510+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 13, "total_pages": 66, "image_filename": "19930082245_p13.jpg", "text": "12\nNACA TN No. 1596\n\n(references 4, 7, and 8). A lower value of $\\left(\\frac{\\Delta c_n}{\\Delta \\alpha}\\right)_\\delta$ would be advantageous in roll since the damping-moment coefficient is a function of this parameter.\n\nAileron.— In addition to relieving hinge moments, thickening the aileron trailing edge reduced section aileron loads appreciably at constant airfoil section normal-force coefficient (figs. 7 and 10), as is to be expected from the action of the bevel on the air flow (fig. 16). The aileron load for the beveled-trailing-edge aileron was affected quite irregularly by changes in angle of attack, aileron deflection, and Mach number.\n\nSection Pitching Moment\n\nThe variation of section pitching-moment coefficient $c_m$ with Mach number and aileron deflection is seen to be more regular for the airfoil with the true-contour aileron than for the airfoil with the beveled-trailing-edge aileron (figs. 6 and 9). Thickening the trailing edge, however, reduced pitching-moment coefficients.\n\nThe pitching-moment coefficient of the airfoil with the beveled-trailing-edge aileron was approximately the same magnitude for deflections of $12^\\circ$ and $18^\\circ$ at several of the airfoil section normal-force coefficients (figs. 9(b), 9(c), 9(d), and 9(e)). For these conditions, the flow followed the contour of the airfoil at a deflection of $12^\\circ$; whereas at a deflection of $18^\\circ$, there was appreciable separation. The effect of separation was to change the section pitching-moment coefficient in a positive direction. For the conditions represented in figure 9(f), appreciable separation had occurred also at a deflection of $12^\\circ$, with a consequent spreading out of the section pitching-moment-coefficient curves for deflections of $12^\\circ$ and $18^\\circ$.\n\nThe section pitching-moment coefficient of the airfoil with the true-contour aileron at a deflection of $18^\\circ$ decreased notably for some of the combinations of Mach number and airfoil section normal-force coefficient. This decrease in section pitching-moment coefficient was mainly due to the development of appreciable separation of the air flow off the upper surface of the airfoil, the separation occurring at lower Mach numbers as the airfoil section normal-force coefficient was increased (fig. 6).\n\nThe section parameter of the rate of change of airfoil section pitching-moment coefficient with deflection per unit value of aileron section effectiveness $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ is a measure of the tendency of an aileron to twist a wing (as a result of the pitching moments produced by the aileron) in terms of the section effectiveness developed by the aileron (fig. 14). It is seen that increasing the Mach number increased", "timestamp": "2026-07-22T04:37:10.365454+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 7, "total_pages": 33, "image_filename": "19930082487_p7.jpg", "text": "NACA TN No. 1813\n\nnumber at several chordwise stations on the upper surface of the NACA 23015 airfoil section at $2^\\circ$ angle of attack. Here the airfoil crest is at the 22-percent-chord station on the upper surface. At points well forward of the crest, that is, at the 5- and 10-percent-chord stations, the local pressure coefficient ceases to fall with increasing Mach number somewhat before the drag-divergence Mach number. Closer to the crest, at the 15- and 20-percent-chord stations, the smooth decrease in pressure coefficient with increasing Mach number continues until the free-stream Mach number reaches the value for drag divergence. After the drag-divergence Mach number is exceeded, the pressure coefficients at points ahead of the airfoil crest rise in such a manner as to maintain virtually constant local Mach numbers. For points immediately behind the airfoil crest, the 25-, 30-, and 40-percent-chord stations on the upper surface, the local Mach number does not rise smoothly from subsonic to supersonic values since, at each chordwise position, as the shock wave moves past, there is an abrupt rise from a local Mach number slightly less than unity to a value greater than unity. However, at each of these chordwise positions, the free-stream Mach number for which the local velocity becomes sonic agrees quite well with the value obtained by assuming the pressure-coefficient variation with free-stream Mach number to follow the Prandtl-Glauert rule. For points far back on the upper surface, the 60- and 85-percent-chord stations, at a free-stream Mach number of about 0.73, there is a marked increase in the rate of fall of local pressure coefficient with increasing free-stream Mach number. The presence of a thick boundary layer over the rear of the airfoil is thus indicated by the pressure distribution and 0.73 is termed the shock-stall Mach number. A similar change in the variation of local pressure coefficient with free-stream Mach number occurs for the entire lower surface of the airfoil (fig. 3(b)) at a free-stream Mach number somewhat below 0.73. This marked decrease of local pressure coefficient with increasing free-stream Mach number resembles the pressure-distribution changes associated with increasing boundary-layer separation as the lift coefficient of an airfoil approaches maximum lift at low speeds. It is interesting to note (fig. 1) that, for the NACA 23015 airfoil section at $2^\\circ$ angle of attack, the marked decrease in lift coefficient with increasing Mach number which begins at a Mach number of about 0.7 is primarily a result of the rapid decrease of pressure coefficients on the lower surface of the airfoil (fig. 3(b)).\n\nMeasurements at Supercritical Speeds for NACA 23015 Airfoil at Various Angles of Attack\n\nIn figure 4 are presented the variations of pressure coefficient", "timestamp": "2026-07-22T04:37:14.023429+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 9, "total_pages": 62, "image_filename": "19930082485_p9.jpg", "text": "8\nNACA TN No. 1810\n\npressure ratio at any radius is the pressure ratio that would be obtained with radial equilibrium of pressures and free-vortex flow at design conditions. The experimental results show a pressure ratio at the root 2 percent lower than the design value and a pressure ratio at the tip 4 percent greater than the design value.\n\nDischarge angle. - The gas-discharge angle measured from the axial direction at a number of radial stations in a plane 0.1 chord downstream of the blades is shown in figure 9(a). In general, the air was turned approximately $1^\\circ$ more than the design value. The gas was considerably overturned in the boundary layer near the inner shroud. This region of severe overturn is followed by an adjacent region at a larger radius in which the gases were turned to the design leaving angle. Near the outer shroud the gases were slightly underturned.\n\nAlthough no data were taken at the outer shroud, the gases were probably overturned in this region. The overturning near the inner shroud is probably caused by secondary flow from the high-pressure region on the pressure surface of the blade to the low-pressure region on the suction surface as explained in reference 6.\n\nThe variation in discharge angle in the circumferential direction at the 10.35-inch radius is shown in figure 9(b). The gases near the suction surface of a blade were turned about $8^\\circ$ more than the gases near the pressure surface of the adjacent blade. Inasmuch as the included angle between suction and pressure surfaces at the trailing edge is $10^\\circ$, some difference between discharge angle near the two surfaces was to be expected.\n\nVane-surface-velocity distribution. - The experimental- and calculated-design-velocity distribution about the root, pitch, and tip sections of the blade are shown in figure 2. The experimental values of velocity on the blade surface were computed from experimental values of static pressure, assuming no loss in total pressure in the region outside the boundary layer on the blade surface. With this assumption, the surface velocity at a point on the blade surface is related to the static pressure at the same point as follows:\n\n$$\n\\frac{V}{V_{cr}} = \\sqrt{\\frac{\\gamma+1}{\\gamma-1} \\left[ 1 - \\left( \\frac{P_s}{P_t} \\right)^{\\frac{\\gamma-1}{\\gamma}} \\right]}\n$$\n\nwhere the total pressure $P_t$ is measured upstream of the blade row.", "timestamp": "2026-07-22T04:37:22.452356+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 24, "total_pages": 37, "image_filename": "19930090382_p24.jpg", "text": "26\nNACA RM L9I07\n\nCONFIDENTIAL\n\nThrust coefficient, $C_T$\nPower coefficient, $C_P$\nTip Mach number, $M_t$\nEfficiency, $\\eta$\n\nAdvance ratio, J\n(h) M=0.70.\nFigure 5 - Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:37:27.976626+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 10, "total_pages": 41, "image_filename": "19930082476_p10.jpg", "text": "8\nNACA TN No. 1801\n\nNeutral and down deflections of the elevator were favorable in preventing the spin; whereas up elevator deflections were conducive to the attainment of spinning equilibrium. From the foregoing results it appears that the fastest recoveries from any spin obtainable would have been effected by reversal of the wheel followed by a downward movement of the elevator.\n\nMass changes (loadings 2 and 3).— Test results obtained with the mass distribution increased along the fuselage are shown in chart 2, and results obtained with the mass distribution increased along the wings are shown in chart 3. These model conditions are represented, respectively, by loadings 2 and 3 in table II and points 2 and 3 in figure 4. More spins were obtained for loading 2, in which the elevator was set between neutral and full up for wheel settings with the spin, than were obtained for the normal-loading condition. Loading 3 gave results very similar to those for the normal loading.\n\nIncreased relative density (loading 4).— Chart 4 shows the results obtained with the weight of the model approximately doubled and with the radii of gyration about the center of gravity (and the mass-distribution parameters) kept approximately the same as for the normal loading (loading 4 in table II and point 4 in fig. 4). The test results obtained at this loading differed from results obtained at the normal loading in that definite spins were now obtained when the wheel was full with the spin and the elevator deflected up normally ($13^\\circ$). Test results obtained at other control configurations were generally the same as those obtained at the normal loading although, when the wheel was full with the spin and the elevator was either neutral or down, a spiral motion was obtained where definite \"no spin\" conditions had previously been obtained. At this loading, it was possible to obtain a spin with wheel-neutral control settings by deflecting the elevator to $30^\\circ$ up.\n\nUnlinked Controls\n\nIn order to establish the individual effects of the ailerons and the rudders in the spin, tests were made with the ailerons deflected when the rudders were neutral and with the rudders deflected when the ailerons were neutral. The results of these tests are presented in charts 5 to 7.\n\nEffect of ailerons.— With the rudders maintained at neutral, the aileron deflections were varied from full against to full with the spin for loadings 1 and 2. The elevator was kept at normal full up ($13^\\circ$) for these tests, and the results are presented in chart 5. Analysis of the results presented indicates that the greatest tendency to spin would occur for the model when the ailerons were placed at one-half or near one-half with the spin.\n\nChart 6 shows the results obtained at loadings 1 to 4 when the right and left ailerons were deflected individually and the rudders were kept at neutral.", "timestamp": "2026-07-22T04:37:28.771463+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 3, "total_pages": 99, "image_filename": "19930082511_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE No. 1826\n\nLINEAR THEORY OF BOUNDARY EFFECTS IN OPEN\n\nWIND TUNNELS WITH FINITE JET LENGTH\n\nBy S. Katzoff, Clifford S. Gardner,  \nLeo Diesendruck, and Bertram J. Eisenstadt\n\nSUMMARY\n\nIn the first part, the boundary conditions for an open wind tunnel (incompressible flow) are examined with special reference to the effects of the closed entrance and exit sections. Basic conditions are that the velocity must be continuous at the entrance lip and that the velocities in the upstream and downstream closed portions must be equal. For the two-dimensional open tunnel, interesting possibilities develop from the fact that the pressures on the two free surfaces need not be equal.\n\nElectrical analogies that might be used for solving the flow in open wind tunnels are outlined. Two types are described — one in which electrical potential corresponds to velocity potential, and another in which electrical potential corresponds to acceleration potential. The acceleration-potential analogies are probably experimentally simpler than the velocity-potential analogies.\n\nIn the second part, solutions are derived for four types of two-dimensional open tunnels, including one in which the pressures on the two free surfaces are not equal. Numerical results are given for every case. In general, if the lifting element is more than half the tunnel height from the inlet, the boundary effect at the lifting element is the same as for an infinitely long open tunnel.\n\nIn the third part is given a general method for calculating the boundary effect in an open circular wind tunnel of finite jet length. Numerical results are given for a lifting element concentrated at a point on the axis.\n\nINTRODUCTION\n\nThe basic theory of boundary corrections for an open wind tunnel was given by Prandtl many years ago (reference 1) and has since been used with reasonable success. The infinitely long open jet that was assumed in Prandtl’s analysis, however, has been frequently questioned as an adequate representation for an open wind tunnel, which normally has a relatively short jet between closed entrance and exit regions.", "timestamp": "2026-07-22T04:37:30.472182+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 85, "total_pages": 92, "image_filename": "19930085930_p85.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:37:32.825680+00:00"}
{"citation_id": "19930092013", "source_url": "https://ntrs.nasa.gov/api/citations/19930092013/downloads/19930092013.pdf", "page_number": 11, "total_pages": 21, "image_filename": "19930092013_p11.jpg", "text": "APPARATUS FOR VARYING EFFECTIVE DIHEDRAL IN FLIGHT\n7\n\n<!-- Image (47, 109, 902, 999) -->\n\n(a) $\\Gamma_e=6.3^\\circ$ (normal, servo inoperative).\n(b) $\\Gamma_e=14.9^\\circ$.\n(c) $\\Gamma_e=-2.7^\\circ$.\nFIGURE 7.—Time histories of abrupt, stick-fixed rudder kicks. $V_t=300$ knots.\n\n(a) $\\Gamma_e 6.3^\\circ$.\nFIGURE 9.—Time histories of lateral oscillations. $V_t=300$ knots.\n\n(a) Variation of maximum $p\\delta/2V$ with $\\Delta \\delta$.\n(b) Variation of rolling response parameter with $\\Gamma_e$.\nFIGURE 8.—Rolling response characteristics in abrupt rudder kicks $V_t=300$ knots.\n872865-59-2\n\n(b) $\\Gamma_e, 14.9^\\circ$.\nFIGURE 9.—Continued.", "timestamp": "2026-07-22T04:37:34.370492+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 49, "total_pages": 67, "image_filename": "19930085965_p49.jpg", "text": "48\nNACA RM E9E06\n\n[Figure: Vector diagram showing phase relations.\nVectors:\n- $I_{max\\ s}$ pointing up and left.\n- $\\phi_{max\\ s}$ pointing straight up.\n- $-j \\frac{P_{max\\ s}}{20}$ pointing up and right.\n- $\\frac{P_{max\\ s}}{20}$ pointing down and left.\n- Horizontal axis with $-E_{max\\ s}$ to the left and $E_{max\\ s}$ to the right.\n- Angle $\\theta = 45^\\circ$ marked between the vertical axis and the vector $\\frac{P_{max\\ s}}{20}$.\nLabels:\n- Left side: \"$-E_{max\\ s}$, voltage consumed by flux $\\phi_{max\\ s}$\"\n- Right side: \"$E_{max\\ s}$, voltage induced by flux $\\phi_{max\\ s}$\"]\n\n1125\n\nNACA\n\nFigure 5. - Vector diagram of phase relation of current, flux, and voltage at constant permeability.", "timestamp": "2026-07-22T04:37:34.821301+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 11, "total_pages": 37, "image_filename": "19930082450_p11.jpg", "text": "```markdown\n10\nNACA TN No. 1778\n\nTABLE 2.- N-FAIRL PROPERTIES\n$$ \\frac{t}{c} = 0.51; \\frac{t}{c_w} = 11.4; \\frac{t}{c_f} = 0.4; \\frac{t}{c_r} = 3; \\frac{t}{c_{tr}} = 4; \\frac{t}{c_{te}} = 1.50; \\frac{t}{c_{te}} = 10.0 $$\n\n| $t/c$ | 20 | 21 | 22 | 23 | 24 | 25 | 26 | 27 | 28 | 29 | 30 | 31 | 32 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 25 | 1.574 | 1.259 | 1.005 | 1.118 | 1.252 | 1.047 | 1.062 | 1.076 | 1.091 | 1.505 | 1.520 | 1.535 | 1.549 |\n| 26 | 1.568 | 1.255 | 1.004 | 1.117 | 1.250 | 1.045 | 1.060 | 1.074 | 1.089 | 1.503 | 1.518 | 1.533 | 1.548 |\n| 27 | 1.562 | 1.250 | 1.003 | 1.116 | 1.248 | 1.043 | 1.058 | 1.072 | 1.087 | 1.501 | 1.516 | 1.531 | 1.546 |\n| 28 | 1.556 | 1.246 | 1.002 | 1.115 | 1.246 | 1.041 | 1.056 | 1.070 | 1.085 | 1.499 | 1.514 | 1.529 | 1.544 |\n| 29 | 1.550 | 1.241 | 1.001 | 1.114 | 1.244 | 1.039 | 1.054 | 1.068 | 1.083 | 1.497 | 1.512 | 1.527 | 1.542 |\n| 30 | 1.544 | 1.236 | 1.000 | 1.113 | 1.242 | 1.037 | 1.052 | 1.066 | 1.081 | 1.495 | 1.510 | 1.525 | 1.540 |\n| 31 | 1.538 | 1.232 | 0.999 | 1.112 | 1.240 | 1.035 | 1.050 | 1.064 | 1.079 | 1.493 | 1.508 | 1.523 | 1.538 |\n| 32 | 1.532 | 1.227 | 0.998 | 1.111 | 1.238 | 1.033 | 1.048 | 1.062 | 1.077 | 1.491 | 1.506 | 1.521 | 1.536 |\n| 33 | 1.526 | 1.222 | 0.997 | 1.110 | 1.236 | 1.031 | 1.046 | 1.060 | 1.075 | 1.489 | 1.504 | 1.519 | 1.534 |\n| 34 | 1.520 | 1.218 | 0.996 | 1.109 | 1.234 | 1.029 | 1.044 | 1.058 | 1.073 | 1.487 | 1.502 | 1.517 | 1.532 |\n| 35 | 1.514 | 1.213 | 0.995 | 1.108 | 1.232 | 1.027 | 1.042 | 1.056 | 1.071 | 1.485 | 1.500 | 1.515 | 1.530 |\n| 36 | 1.508 | 1.208 | 0.994 | 1.107 | 1.230 | 1.025 | 1.040 | 1.054 | 1.069 | 1.483 | 1.498 | 1.513 | 1.528 |\n| 37 | 1.502 | 1.204 | 0.993 | 1.106 | 1.228 | 1.023 | 1.038 | 1.052 | 1.067 | 1.481 | 1.496 | 1.511 | 1.526 |\n| 38 | 1.496 | 1.199 | 0.992 | 1.105 | 1.226 | 1.021 | 1.036 | 1.050 | 1.065 | 1.479 | 1.494 | 1.509 | 1.524 |\n| 39 | 1.490 | 1.194 | 0.991 | 1.104 | 1.224 | 1.019 | 1.034 | 1.048 | 1.063 | 1.477 | 1.492 | 1.507 | 1.522 |\n| 40 | 1.484 | 1.190 | 0.990 | 1.103 | 1.222 | 1.017 | 1.032 | 1.046 | 1.061 | 1.475 | 1.490 | 1.505 | 1.520 |\n| 41 | 1.478 | 1.185 | 0.989 | 1.102 | 1.220 | 1.015 | 1.030 | 1.044 | 1.059 | 1.473 | 1.488 | 1.503 | 1.518 |\n| 42 | 1.472 | 1.180 | 0.988 | 1.101 | 1.218 | 1.013 | 1.028 | 1.042 | 1.057 | 1.471 | 1.486 | 1.501 | 1.516 |\n| 43 | 1.466 | 1.176 | 0.987 | 1.100 | 1.216 | 1.011 | 1.026 | 1.040 | 1.055 | 1.469 | 1.484 | 1.499 | 1.514 |\n| 44 | 1.460 | 1.171 | 0.986 | 1.099 | 1.214 | 1.009 | 1.024 | 1.038 | 1.053 | 1.467 | 1.482 | 1.497 | 1.512 |\n| 45 | 1.454 | 1.166 | 0.985 | 1.098 | 1.212 | 1.007 | 1.022 | 1.036 | 1.051 | 1.465 | 1.480 | 1.495 | 1.510 |\n| 46 | 1.448 | 1.162 | 0.984 | 1.097 | 1.210 | 1.005 | 1.020 | 1.034 | 1.049 | 1.463 | 1.478 | 1.493 | 1.508 |\n| 47 | 1.442 | 1.157 | 0.983 | 1.096 | 1.208 | 1.003 | 1.018 | 1.032 | 1.047 | 1.461 | 1.476 | 1.491 | 1.506 |\n| 48 | 1.436 | 1.152 | 0.982 | 1.095 | 1.206 | 1.001 | 1.016 | 1.030 | 1.045 | 1.459 | 1.474 | 1.489 | 1.504 |\n| 49 | 1.430 | 1.148 | 0.981 | 1.094 | 1.204 | 0.999 | 1.014 | 1.028 | 1.043 | 1.457 | 1.472 | 1.487 | 1.502 |\n| 50 | 1.424 | 1.143 | 0.980 | 1.093 | 1.202 | 0.997 | 1.012 | 1.026 | 1.041 | 1.455 | 1.470 | 1.485 | 1.500 |\n| 51 | 1.418 | 1.138 | 0.979 | 1.092 | 1.200 | 0.995 | 1.010 | 1.024 | 1.039 | 1.453 | 1.468 | 1.483 | 1.498 |\n| 52 | 1.412 | 1.134 | 0.978 | 1.091 | 1.198 | 0.993 | 1.008 | 1.022 | 1.037 | 1.451 | 1.466 | 1.481 | 1.496 |\n| 53 | 1.406 | 1.129 | 0.977 | 1.090 | 1.196 | 0.991 | 1.006 | 1.020 | 1.035 | 1.449 | 1.464 | 1.479 | 1.494 |\n| 54 | 1.400 | 1.124 | 0.976 | 1.089 | 1.194 | 0.989 | 1.004 | 1.018 | 1.033 | 1.447 | 1.462 | 1.477 | 1.492 |\n| 55 | 1.394 | 1.120 | 0.975 | 1.088 | 1.192 | 0.987 | 1.002 | 1.016 | 1.031 | 1.445 | 1.460 | 1.475 | 1.490 |\n| 56 | 1.388 | 1.115 | 0.974 | 1.087 | 1.190 | 0.985 | 1.000 | 1.014 | 1.029 | 1.443 | 1.458 | 1.473 | 1.488 |\n| 57 | 1.382 | 1.110 | 0.973 | 1.086 | 1.188 | 0.983 | 0.998 | 1.012 | 1.027 | 1.441 | 1.456 | 1.471 | 1.486 |\n| 58 | 1.376 | 1.106 | 0.972 | 1.085 | 1.186 | 0.981 | 0.996 | 1.010 | 1.025 | 1.439 | 1.454 | 1.469 | 1.484 |\n| 59 | 1.370 | 1.101 | 0.971 | 1.084 | 1.184 | 0.979 | 0.994 | 1.008 | 1.023 | 1.437 | 1.452 | 1.467 | 1.482 |\n| 60 | 1.364 | 1.096 | 0.970 | 1.083 | 1.182 | 0.977 | 0.992 | 1.006 | 1.021 | 1.435 | 1.450 | 1.465 | 1.480 |\n| 61 | 1.358 | 1.092 | 0.969 | 1.082 | 1.180 | 0.975 | 0.990 | 1.004 | 1.019 | 1.433 | 1.448 | 1.463 | 1.478 |\n| 62 | 1.352 | 1.087 | 0.968 | 1.081 | 1.178 | 0.973 | 0.988 | 1.002 | 1.017 | 1.431 | 1.446 | 1.461 | 1.476 |\n| 63 | 1.346 | 1.082 | 0.967 | 1.080 | 1.176 | 0.971 | 0.986 | 1.000 | 1.015 | 1.429 | 1.444 | 1.459 | 1.474 |\n| 64 | 1.340 | 1.078 | 0.966 | 1.079 | 1.174 | 0.969 | 0.984 | 0.998 | 1.013 | 1.427 | 1.4", "timestamp": "2026-07-22T04:37:35.896858+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 35, "total_pages": 42, "image_filename": "19930086078_p35.jpg", "text": "NACA RM L9H04\n33\n\nCONFIDENTIAL\n\n<!-- Image (108, 156, 932, 773) -->\n\nFigure 12.- Lateral control characteristics of 45° sweptback wing with large-chord wing-tip aileron at various deflections, fully extended.", "timestamp": "2026-07-22T04:37:37.110548+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 14, "total_pages": 47, "image_filename": "19930093773_p14.jpg", "text": "NACA RM E9G09\n\nAir-flow values obtained from measurements in the venturi of the inlet-air duct and at the exhaust nozzle agreed within 3 percent with those obtained from measurements at the engine inlet.\n\nThrust. - The thrust of the installation was independently determined from balance-scale measurements and also from pressures and temperatures measured near the exhaust-nozzle outlet by means of a survey rake. Because of the inefficiency of the exhaust nozzle, the scale thrust is less than the rake thrust.\n\nJet thrust was determined from balance-scale measurements by the use of the following equation:\n\n$$\nF_{j,s} = D + B + D_r + \\frac{W_a V_x}{g} + A_x (p_x - p_0)\n\\tag{5}\n$$\n\nNet thrust is then given by the equation\n\n$$\nF_{n,s} = F_{j,s} - \\frac{W_a}{g} V_e\n\\tag{6}\n$$\n\nThe last two terms of equation (5) represent the momentum and the pressure forces on the installation at the slip joint in the inlet-air duct. The drag of the installation was determined by runs with the engine inoperative and with a blocking plate installed in the inlet to prevent air flow through the engine.\n\nThe rake thrust, which is the ideal thrust available, is given by the following equation and values obtained at station 7, 1 inch upstream of the nozzle outlet:\n\n$$\nF_{j,r} = \\frac{2C_t A_7 p_7 \\gamma_7}{\\gamma_7 - 1} \\left[ \\left( \\frac{p_7}{p_7} \\right)^{\\frac{\\gamma_8 - 1}{\\gamma_8}} - 1 \\right] + C_t A_7 (p_7 - p_0)\n\\tag{7}\n$$\n\nAlternate thrust equation. - When the assumption is made that $p_7 = p_8$, an alternate equation for jet thrust is as follows:", "timestamp": "2026-07-22T04:37:40.004236+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 81, "total_pages": 104, "image_filename": "19930085842_p81.jpg", "text": "NACA RM L9C29\n77\n\n[Figure: Graph plotting Propeller advance diameter ratio, $V/nD$ and Lift coefficient, $C_L$ against Torque coefficient, $Q_c$; and Resultant-drag coefficient, $C_{DR}$ against Torque coefficient, $Q_c$.]\n\nPropeller advance\ndiameter ratio, $V/nD$\n10\n8\n6\n4\n\n$V/nD$\n\n$C_{DR}$\n\nResultant-drag coefficient, $C_{DR}$\n4\n0\n-4\n-8\n\nLift coefficient, $C_L$\n12\n10\n8\n6\n4\n2\n0\n\n$C_L$\n\n$B, deg$\n$\\circ$ 10\n$\\diamond$ 14\n$\\triangle$ 20\n\n0\n0.1\n0.2\n0.3\n0.4\nTorque coefficient, $Q_c$\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(d) $\\alpha_u = 15^\\circ$.\nFigure 43.— Continued.", "timestamp": "2026-07-22T04:37:42.565675+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 12, "total_pages": 24, "image_filename": "19930082447_p12.jpg", "text": "10\nNACA TN No. 1775\n\nTABLE I. - OFFSETS OF LANGLEY IMPACT-BASIN\nFLOAT MODEL M-3 (SEE FIG. 1)\n[All dimensions are in inches]\n\n| Station | Half breadth | | Height above datum line | | |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| | Chine | Deck | Keel | Chine | Deck |\n| 0 | 0 | 0.33 | 27.06 | 27.06 | 37.60 |\n| 2 | 2.15 | 1.45 | 21.34 | 26.08 | 38.17 |\n| 5 | 4.25 | 3.05 | 17.12 | 25.97 | 38.81 |\n| 9 | 7.80 | 4.58 | 12.85 | 27.06 | 39.51 |\n| 14 | 10.31 | 5.93 | 9.05 | 24.90 | 40.09 |\n| 21 | 12.81 | 7.23 | 5.62 | 21.90 | 40.52 |\n| 29 | 15.09 | 8.15 | 3.01 | 19.08 | 40.59 |\n| 38 | 16.86 | 8.71 | 1.13 | 16.55 | 40.59 |\n| 47 | 18.04 | 8.94 | .27 | 15.53 | 40.59 |\n| 58 | 18.87 | 9.00 | 0 | 15.56 | 40.59 |\n| 72 | 19.33 | 9.00 | 0 | 15.94 | 40.59 |\n| 87.25 | 19.40 | 9.00 | 0 | 16.00 | 40.59 |\n| 106.625 | 19.40 | 9.00 | 0 | 16.00 | 40.59 |\n| 120.75 | 19.40 | 9.00 | 0 | 16.00 | 40.59 |\n\n[NACA logo]", "timestamp": "2026-07-22T04:37:47.939285+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 25, "total_pages": 34, "image_filename": "19930086151_p25.jpg", "text": "NACA RM L9D28\n\nCONFIDENTIAL\n\nYawing-moment coefficient, $C_n$\nRolling-moment coefficient, $C_l$\n\n$\\delta_a$ (deg)\n$\\square$ 2.1\n$\\diamond$ 4.2\n$\\triangle$ 9.6\n$\\nabla$ 14.7\n$\\circ$ 19.5\n$\\triangledown$ 29.5\n\nAngle of attack, $\\alpha$, deg\n\nCONFIDENTIAL\n\nNACA\n\nFigure 7.— The rolling-moment and yawing-moment characteristics of the 45° sweptback wing for various deflections of the triangular wing-tip aileron. Wing with end plate.\n\n23", "timestamp": "2026-07-22T04:37:49.515896+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 42, "total_pages": 54, "image_filename": "19930086015_p42.jpg", "text": "NACA RM A9E24 CONFIDENTIAL 41\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:37:52.343578+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 14, "total_pages": 66, "image_filename": "19930082245_p14.jpg", "text": "NACA TN No. 1596\n13\n\nthe value of $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ for both configurations (fig. 14) and thereby made the wing-twist problem at high speeds worse. The increase of $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ with Mach number was mainly due to the decrease in section effectiveness $-\\left(\\frac{\\Delta c_l}{\\Delta \\delta_a}\\right)_{c_n}$ with Mach number (fig. 12). The section pitching-moment parameter $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ was generally less, and the Mach number effects were not so pronounced, for the configuration with the beveled-trailing-edge aileron as for the configuration with the true-contour aileron (fig. 14). The reduction in the parameter $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ for the airfoil with the beveled-trailing-edge aileron resulted from smaller values of the section parameter $\\left(\\frac{\\Delta c_m}{\\Delta \\delta_a}\\right)_{c_n}$, which also decreased with Mach number instead of increasing as was the case for the airfoil with the true-contour aileron.\n\nSection Critical Speed\n\nThe section critical Mach number of the airfoil with either aileron was 0.70 at an airfoil section normal-force coefficient of zero and with the aileron neutral (fig. 13). At positive aileron deflections the section critical Mach number was essentially the same for the airfoil with either aileron except at an airfoil section normal-force coefficient of 0.6, at which value the airfoil with the beveled-trailing-edge aileron had lower section critical-speed values. At negative deflections, the section critical Mach number of the airfoil with the beveled-trailing-edge aileron generally was lower than that of the airfoil with the true-contour aileron.\n\nIn the data of figure 13, the upper surface of the main portion of the airfoil was the critical surface, except at negative aileron deflections greater than -7.5° at an airfoil section normal-force coefficient of zero, at which conditions the lower surface of the main portion of the airfoil was the critical one. The pressures over the ailerons were more positive than the minimum pressure occurring on the main portion of the airfoil for all test conditions, so that the critical Mach number of the ailerons was greater than that of the main portion of the airfoil.\n\nFigures 5 and 8 illustrate the characteristically flat chordwise pressure distributions of the 66-series airfoil section and the effect of Mach number on the pressures. The pressure diagrams over the main portion of the airfoil are seen to be very similar for both aileron configurations. The typical large changes in pressure distribution at supercritical Mach numbers for relatively small changes in Mach number are to be noted.", "timestamp": "2026-07-22T04:37:52.537527+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 33, "total_pages": 44, "image_filename": "19930086081_p33.jpg", "text": "NACA RM L9H05\nCONFIDENTIAL\n31\n\n<!-- Image (204, 78, 886, 824) -->\n\nFigure 13.- Aerodynamic characteristics of a 3-percent-thick half-delta tip control surface mounted on a semispan delta wing. Fence off. Coefficients based on dimensions of the complete semispan wing. M = 1.90. Flagged symbols denote repeat tests. Initial series of tests.", "timestamp": "2026-07-22T04:38:17.975020+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 50, "total_pages": 67, "image_filename": "19930085965_p50.jpg", "text": "NACA RM E9E06\n49\n\n1125\n\n[Figure: Vector diagram of phase relation of current, flux, and voltage under conditions of saturation.]\n\nFigure 6. - Vector diagram of phase relation of current, flux, and voltage under conditions of saturation.", "timestamp": "2026-07-22T04:38:26.237470+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 25, "total_pages": 37, "image_filename": "19930090382_p25.jpg", "text": "NACA RM L9I07\n27\n\nCONFIDENTIAL\n\nTip Mach number, $M_t$\n1.5\n1.0\n0.5\n0\n\nEfficiency, $\\eta$\n1.00\n.75\n.50\n.25\n0\n\nPower coefficient, $C_P$\n1.20\n1.10\n1.00\n.90\n.80\n.70\n.60\n.50\n.40\n.30\n.20\n.10\n0\n\nThrust coefficient, $C_T$\n.300\n.275\n.250\n.225\n.200\n.175\n.150\n.125\n.100\n.075\n.050\n.025\n0\n\nAdvance ratio, J\n0 .5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.5 7.0 7.5 8.0 8.5 9.0\n\nCONFIDENTIAL\n(h) M=0.70 Concluded\nFigure 5 - Continued.\n\n[Graph showing curves for $C_P$, $C_T$, $\\eta$, and $M_t$ plotted against Advance ratio, J. The curves are labeled with $\\beta_{0.75R} = 65^\\circ$.]\n\nNACA", "timestamp": "2026-07-22T04:38:27.511108+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 36, "total_pages": 42, "image_filename": "19930086078_p36.jpg", "text": "34\nNACA RM L9H04\n\nCONFIDENTIAL\n\n<!-- Image (58, 183, 888, 759) -->\n\nFigure 13.- Lateral control characteristics of 45° sweptback wing with large-chord wing-tip aileron at various extensions. $\\delta_a = 4^\\circ$.", "timestamp": "2026-07-22T04:38:29.965327+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 26, "total_pages": 34, "image_filename": "19930086151_p26.jpg", "text": "```markdown\n24\n\nCONFIDENTIAL\n\nYawing-moment coefficient, $C_n$\nRolling-moment coefficient, $C_l$\n\n$\\delta_a$ (deg)\n$\\square$ 2.1\n$\\diamond$ 4.2\n$\\nabla$ 6.0\n$\\triangle$ 9.6\n$\\square$ 14.7\n$\\square$ 19.5\n$\\square$ 29.5\n\nCONFIDENTIAL\n\nNACA\n\nAngle of attack, $\\alpha$, deg\n\nFigure 8.—The rolling-moment and yawing-moment characteristics of the 45° sweptback wing for various deflections of the parallelogram wing-tip aileron. Wing with end plate.\n\nNACA RM L9D28\n```", "timestamp": "2026-07-22T04:38:30.291143+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 15, "total_pages": 47, "image_filename": "19930093773_p15.jpg", "text": "14\nNACA RM E9G09\n\n$$F_j = \\frac{2\\gamma_8}{\\gamma_8-1} C_t A_8 p_8 \\left[ \\left( \\frac{p_8}{p_8} \\right)^{\\frac{\\gamma_8-1}{\\gamma_8}} - 1 \\right] + A_8 C_t (p_8 - p_0) \\quad (8)$$\n\nwhere\n\n$$p_8 = \\frac{P_8}{\\left( \\frac{\\gamma_8+1}{2} \\right)^{\\frac{\\gamma_8}{\\gamma_8-1}}}$$\n\nand for supersonic jet velocities where\n\n$$\\frac{P_8}{P_0} > 1.9$$\n\nFor subsonic jet velocities where\n\n$$\\frac{P_8}{P_0} < 1.9$$\n\nequation (8) reduces to\n\n$$F_j = \\frac{2\\gamma_8 A_8 p_0 C_t}{\\gamma_8-1} \\left[ \\left( \\frac{p_8}{p_0} \\right)^{\\frac{\\gamma_8-1}{\\gamma_8}} - 1 \\right] \\quad (9)$$", "timestamp": "2026-07-22T04:38:42.909437+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 11, "total_pages": 41, "image_filename": "19930082476_p11.jpg", "text": "NACA TN No. 1801\n\nThe results indicated that: When the inboard aileron was maintained at neutral, no spin was obtained regardless of the outboard aileron deflection; whereas, when the inboard aileron was deflected from approximately three-tenths to six-tenths of its maximum full-up deflection, a spin was obtained regardless of the position of the outboard aileron.\n\nIt thus appears from the results that in order to spinproof an airplane proportioned similarly to the model tested, limiting the up aileron to about $5^\\circ$ would be desirable. The normal differential aileron movements employed for the linked-control tests appear ineffective in preventing the spin.\n\nEffect of rudders.—With the ailerons maintained at neutral, the rudder deflections for loadings 2, 3, and 4 were varied from neutral to as much as $20^\\circ$ with the spin for the outboard rudder and to as much as $45^\\circ$ with the spin for the inboard rudder. The elevator was kept at its normal full-up deflection ($13^\\circ$) for these tests, and the results are presented in chart 7. The results show that if the outboard rudder was at or near neutral, no spin could be obtained regardless of the position of the inboard rudder. If the outboard rudder was set with the spin, however, the results indicate that spins could be obtained even if the inboard rudder was at neutral. The amount the outboard rudder had to be set with the spin in order to obtain a spinning condition varied somewhat with loading. The results show that the outboard rudder was the more effective rudder during the spin and that differential rudder deflection in which the outboard rudder is maintained at or near neutral is effective in preventing the attainment of spinning equilibrium when the ailerons are neutral.\n\nTests in which the model was launched with the rudders set against the spin are presented in chart 8 for loadings 3 and 4. The results indicate that for loading 4 (increased relative density) the model would not spin when both rudders were $20^\\circ$ against the spin even though the aileron deflection was such as to be very conducive in causing the model to spin. The model ceased spinning quickly after being launched into the tunnel, thereby indicating that recovery by movement of the rudders from with the spin to against the spin would have been rapid. When, however, the mass was distributed heavily along the wings (loading 3), the results indicate that rudder reversal alone would not effect recovery. Inasmuch as references 1 and 4 indicate that rudder effectiveness decreases and elevator effectiveness increases as the mass distribution of airplanes is increased along the wings, this result appears reasonable; thus, in order to obtain satisfactory recovery at loading 3, rudder reversal would have to be followed by a downward movement of the elevator. For loading 4, on the other hand, the results indicate that even though the relative density was comparatively high ($\\mu = 10$ approx.), the rudders were effective in terminating the spin for this mass distribution $\\left(\\frac{I_x - I_y}{mb^2} = -18 \\times 10^{-4}\\right)$. On the basis of the results obtained at loading 4 and on the basis of reference 5, which indicates that decreased relative density improves recovery, it can be concluded that rudder action alone would have been effective in terminating spins obtained for loadings 1 and 2.", "timestamp": "2026-07-22T04:38:45.130579+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 15, "total_pages": 66, "image_filename": "19930082245_p15.jpg", "text": "14\nNACA TN No. 1596\n\nWing Drag\n\nThe wing-drag data from force-test measurements (fig. 17) showed that, at constant airfoil section normal-force coefficient and with the aileron neutral, beveling the trailing edge of the aileron increased the drag of the airfoil. The increment in drag became greater at higher values of airfoil section normal-force coefficient. The variation of drag coefficient with Mach number was essentially the same for both configurations. The low-speed drag coefficient shown in figure 17 is appreciably higher than the low-speed profile-drag-coefficient value of 0.004 obtained in other tests (reference 11) for the NACA 66,1-115 airfoil section. The main reason for the discrepancy is the air leakage through the gap between the model and the tunnel walls, the effect of the leakage being to increase the force drag of the wing. The span-wise gaps on the upper and lower surfaces of the wing between the aileron and the aileron cover plates also probably increased the drag of the basic section somewhat.\n\nCONCLUSIONS\n\nAn investigation was made in the Langley 8-foot high-speed tunnel of the section characteristics of a 24-inch-chord NACA 66,1-115 airfoil section equipped with unsealed 20-percent-chord plain ailerons of true-airfoil-contour profile and 30° beveled-trailing-edge profile. The airfoil was tested with aerodynamically smooth surfaces for Mach numbers up to 0.75, and for various airfoil angles of attack and aileron deflections. The test Reynolds number at the highest speed was $7.5 \\times 10^6$. The following conclusions are indicated:\n\n1. The aileron section effectiveness parameter $-\\left(\\frac{\\Delta c_l}{\\Delta \\delta_a}\\right)_{c_n}$ for both of the aileron profiles investigated decreased between the Mach numbers 0.25 and 0.75 by about one-half the low-speed value, at the lower values of airfoil section normal-force coefficient.\n\n2. The section pitching-moment parameter $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$ for both aileron profiles increased with Mach number and thereby aggravated the wing-twist problem at high speeds.\n\n3. Changing the aileron profile from the true-contour profile to the 30° beveled-trailing-edge profile caused\n(a) a decrease in the aileron section effectiveness parameter $-\\left(\\frac{\\Delta c_l}{\\Delta \\delta_a}\\right)_{c_n}$", "timestamp": "2026-07-22T04:38:45.286493+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 37, "total_pages": 42, "image_filename": "19930086078_p37.jpg", "text": "NACA RM L9H04\n35\n\nCONFIDENTIAL\n\n<!-- Image (103, 139, 928, 800) -->\n\nFigure 14.- Lateral control characteristics of 45° sweptback wing with triangular wing-tip aileron at various deflections, fully extended.", "timestamp": "2026-07-22T04:38:51.931542+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 27, "total_pages": 34, "image_filename": "19930086151_p27.jpg", "text": "NACA RM L9J28\n25\n\nCONFIDENTIAL\n\nRolling-moment coefficient, $C_l$\n\n(a) Triangular wing-tip aileron.\n\n$\\alpha$ (deg) | $C_L$\n---|---\n0 | 0.01\n5 | .21\n10 | .43\n16 | .72\n22 | 1.02\n\n(b) Parallelogram wing-tip aileron.\n\nCONFIDENTIAL\n\nAileron deflection, $\\delta_a$, deg\n\nFigure 9.— The rolling-moment characteristics of the 45° sweptback wing at various angles of attack and aileron deflections. Plain wing.", "timestamp": "2026-07-22T04:38:55.337246+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 86, "total_pages": 92, "image_filename": "19930085930_p86.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n\nNACA RM L9G07\n\nTotal-pressure tube\n\n[Figure: A photograph of an aerodynamic setup showing a wing section and a probe.]\n\nNACA\n\nFigure 40.- A photograph of the angle of attack setup without flow.\nUNCLASSIFIED\nCONFIDENTIAL\n\n85", "timestamp": "2026-07-22T04:39:00.924268+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 43, "total_pages": 54, "image_filename": "19930086015_p43.jpg", "text": "42\n\nCONFIDENTIAL\n\nNACA RM A9E24\n\n| | | |\n| :--- | :--- | :--- |\n| $D=165.12$<br>$M=1.23$<br>$\\Delta p/q = .03$ | $D=147.20$<br>$M=1.32$<br>$\\Delta p/q = -.06$ | $D=129.32$<br>$M=1.43$<br>$\\Delta p/q = -.06$ |\n| $D=113.43$<br>$M=1.53$<br>$\\Delta p/q = -.01$ | $D=100.47$<br>$M=1.63$<br>$\\Delta p/q = .02$ | $D=88.52$<br>$M=1.73$<br>$\\Delta p/q = .02$<br>[NACA logo] |\n\nDistance from wall, $\\delta$, in.\nRatio of total pressures, $\\frac{H}{H_0}$\n\n(a) Upper rake.\n\nFigure 12.— The boundary-layer profiles on the nozzle walls of the Ames 6- by 6-foot supersonic wind tunnel. Stagnation pressure = 9 lb/sq in. abs.", "timestamp": "2026-07-22T04:39:01.961444+00:00"}
{"citation_id": "19930090382", "source_url": "https://ntrs.nasa.gov/api/citations/19930090382/downloads/19930090382.pdf", "page_number": 26, "total_pages": 37, "image_filename": "19930090382_p26.jpg", "text": "28\nNACA RM L9I07\n\nCONFIDENTIAL\n\nCONFIDENTIAL\n\nThrust coefficient, $C_T$\nPower coefficient, $C_P$\n\nTip Mach number, $M_t$\n\nEfficiency, $\\eta$\n\nAdvance ratio, J\n$\\beta_{0.75}=50^\\circ$\n(1) M=0.75\nFigure 5 - Continued.\n\nNACA", "timestamp": "2026-07-22T04:39:08.210285+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 16, "total_pages": 47, "image_filename": "19930093773_p16.jpg", "text": "NACA RM E9G09\n\n15\n\nREFERENCE\n\n1. Sanders, Newell D.: Performance Parameters for Jet-Propulsion Engines. NACA TN 1106, 1946.", "timestamp": "2026-07-22T04:39:12.530426+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 82, "total_pages": 104, "image_filename": "19930085842_p82.jpg", "text": "78\nNACA RM L9C29\n\nPropeller advance-\ndiameter ratio, $V/nD$\n1.4\n1.2\n1.0\n.8\n\nLift coefficient, $C_L$\n12\n8\n4\n0\n\nTorque coefficient, $Q_c$\n0 .01 .02\n\n$V/nD$\n\n$C_{DR}$\n\n$Q_c$\n\n$\\beta, deg$\n30\n\nResultant drag\ncoefficient, $C_{DR}$\n4\n0\n\nNATIONAL ADVISORY\nCOMMITTEE FOR AERONAUTICS\n\n(e) $\\alpha_u = 21^\\circ$.\nFigure 43.- Continued.", "timestamp": "2026-07-22T04:39:18.060646+00:00"}
{"citation_id": "19930086081", "source_url": "https://ntrs.nasa.gov/api/citations/19930086081/downloads/19930086081.pdf", "page_number": 34, "total_pages": 44, "image_filename": "19930086081_p34.jpg", "text": "32\nCONFIDENTIAL\nNACA RM L9H05\n\n$$\n\\begin{array}{c}\n.04 \\\\\nC_L \\quad 0 \\\\\n-.04\n\\end{array}\n$$\n\n$$\n\\begin{array}{c}\n.01 \\\\\nC_D \\quad 0\n\\end{array}\n$$\n\n$$\n\\begin{array}{c}\n.02 \\\\\nC_m \\quad 0 \\\\\n-.02\n\\end{array}\n$$\n\n$$\n\\begin{array}{c}\n\\delta \\\\\n(deg) \\\\\n\\circ \\quad 0.3 \\\\\n\\circ \\quad 2.3 \\\\\n\\diamond \\quad 3.4 \\\\\n\\triangleleft \\quad 7.0 \\\\\n\\triangleright \\quad 9.6\n\\end{array}\n$$\n\n$$\n\\begin{array}{c}\n.008 \\\\\n.004 \\\\\nC_l \\quad 0 \\\\\n-.004\n\\end{array}\n$$\n\n$$\n\\begin{array}{c}\n.004 \\\\\nC_n \\quad 0 \\\\\n-.004\n\\end{array}\n$$\n\n$$\n\\begin{array}{ccccccccccc}\n-7 & -6 & -5 & -4 & -3 & -2 & -1 & 0 & 1 & 2 & 3 \\\\\n& & & & & & & & & & \\text{NACA}\n\\end{array}\n$$\n\n$$\n\\alpha, \\text{deg}\n$$\n\nCONFIDENTIAL\n\nFigure 14.- Aerodynamic characteristics of a 3-percent-thick half-delta tip control surface mounted on a semispan delta wing. Large fence on. Coefficients based on dimensions of the complete semispan wing. M = 1.90. Flagged symbols denote repeat tests. Initial series of tests.", "timestamp": "2026-07-22T04:39:21.112614+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 10, "total_pages": 62, "image_filename": "19930082485_p10.jpg", "text": "NACA TN No. 1810\n\nIn general, the experimental velocities are higher than design values on the pressure surface and lower on the suction surface. The differences are more pronounced at the root and tip sections than at the pitch section; however, these sections are at the edge of the boundary layers on the inner and outer shrouds. The shear stresses in the boundary layer cause an increase in entropy and deviation from the theoretical velocities.\n\nTangential- and axial-velocity distribution. - The gas-discharge-velocity components were computed from the pressure ratios of figure 8 and the discharge angles of figure 9(a), assuming that the radial component of velocity is small enough to be neglected.\n\nThe experimental and design values of critical velocity ratio $V/V_{cr}$ are plotted against radius in figure 10(a). The measured values are 3 percent higher than design near the blade tip and 3 percent lower near design near the blade root.\n\nThe measured values of tangential component of critical velocity ratio $V_u/V_{cr}$ are 3 percent higher than design near the blade tip and 3 percent lower near the blade root, as shown in figure 10(b). Vortex distribution of $V_u/V_{cr}$ was not obtained; but inasmuch as the values are, in general, greater than design values, the turbine power will probably exceed the design power by about 2 percent if the rotor is assumed to perform as designed.\n\nConstant axial velocity was not achieved (fig. 11(a)); the experimental values were about 6 percent lower than the design value at the 10-inch radius and 6 percent higher near the blade tip.\n\nThe weight-flow parameter $\\frac{\\rho_s V_x}{\\rho_t V_{cr}}$ obtained from the data is compared with the design values in figure 11(b). This parameter is quite close to the design value at the blade sections at which the blade-surface-pressure distributions were obtained; at the 10-inch radius, however, this parameter is 6 percent lower than the design value. Although no accurate computation of the experimental boundary-layer thicknesses at the blade root and tip can be made because of insufficient data describing conditions near the tunnel shrouds, the check between design weight flow and experimental weight flow is considered close enough to say that the design allowance for boundary-layer growth at the shroud was adequate. The principal reason that vortex flow was not achieved is because radial equilibrium of pressure was not attained. The heavy boundary layer at the inner shroud also tends to deflect the air toward the outer shroud and distorts the flow near the inner shroud.", "timestamp": "2026-07-22T04:39:21.440524+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 1, "total_pages": 49, "image_filename": "19930082498_p1.jpg", "text": "NATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE\nNo. 1838\n\nDYNAMOMETER-STAND INVESTIGATION OF A GROUP OF MUFFLERS\n\nBy Don D. Davis, Jr., and K. R. Czarnecki\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\n[Figure: NACA logo]\n\nWashington\nMarch 1949", "timestamp": "2026-07-22T04:39:23.381954+00:00"}
{"citation_id": "19930086078", "source_url": "https://ntrs.nasa.gov/api/citations/19930086078/downloads/19930086078.pdf", "page_number": 38, "total_pages": 42, "image_filename": "19930086078_p38.jpg", "text": "36\nNACA RM L9H04\n\nCONFIDENTIAL\n\n<!-- Image (70, 198, 892, 763) -->\n\nFigure 15.- Lateral control characteristics of 45° sweptback wing with triangular wing-tip aileron at various extensions. $\\delta_a = 4^\\circ$.", "timestamp": "2026-07-22T04:39:24.887713+00:00"}
{"citation_id": "19930086151", "source_url": "https://ntrs.nasa.gov/api/citations/19930086151/downloads/19930086151.pdf", "page_number": 28, "total_pages": 34, "image_filename": "19930086151_p28.jpg", "text": "26\nNACA RM L9J28\n\nCONFIDENTIAL\n\nRolling-moment coefficient, $C_l$\n\n(a) Triangular wing-tip aileron.\n\n$\\alpha$ $C_L$\n(deg)\n0 0\n5 .23\n10 .46\n16 .69\n22 .84\n\n(b) Parallelogram wing-tip aileron.\n\nCONFIDENTIAL\n\nAileron deflection, $\\delta_a$, deg\n\nFigure 10.—The rolling-moment characteristics of the 45° sweptback wing at various angles of attack and aileron deflections. Wing with end plate.", "timestamp": "2026-07-22T04:39:25.715518+00:00"}
{"citation_id": "19930082245", "source_url": "https://ntrs.nasa.gov/api/citations/19930082245/downloads/19930082245.pdf", "page_number": 16, "total_pages": 66, "image_filename": "19930082245_p16.jpg", "text": "NACA TN No. 1596\n15\n\n(b) a reduction in hinge moments, but made the hinge-moment\ncharacteristics irregular with an overbalance of the section\nhinge-moment parameter $\\left(\\frac{\\Delta c_h}{\\Delta \\delta_a}\\right)_{\\alpha=0^\\circ}$ at moderate aileron\ndeflections, the overbalance worsening with increase in\nMach number\n\n(c) a decrease in section normal-force-curve slopes $\\left(\\frac{\\Delta c_n}{\\Delta \\alpha}\\right)_{\\delta_a=0^\\circ}$\nand $\\left(\\frac{\\Delta c_n}{\\Delta \\delta_a}\\right)_{\\alpha=0^\\circ}$\n\n(d) a decrease in aileron section loads, at constant airfoil\nsection normal-force coefficient\n\n(e) a general decrease in the section pitching-moment\nparameter $\\left(\\frac{\\Delta c_m}{\\Delta \\alpha}\\right)_{c_n}$\n\n(f) generally only small change in the section critical Mach number\nof the airfoil at positive aileron deflections and a\ndecrease in section critical Mach number at the larger\nnegative aileron deflections, at constant airfoil section\nnormal-force coefficient\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Field, Va., July 1, 1948", "timestamp": "2026-07-22T04:39:26.279103+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 87, "total_pages": 92, "image_filename": "19930085930_p87.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:39:26.480596+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 95, "total_pages": 118, "image_filename": "19930085838_p95.jpg", "text": "```markdown\nNACA RM No. L9B23\n93\n\n<!-- Image (164, 125, 894, 808) -->\n\n(a) $\\delta_f = 25^\\circ$.\n\nFigure 12.- Hinge-moment characteristics of a flap on the approximately 15.4-percent-chord thick NACA 7-series-type airfoil with double slotted flap and straight-sided Frise aileron. Aileron balance, 0.351ca; R = 6 x $10^6$ (approx.)\n```", "timestamp": "2026-07-22T04:39:31.228723+00:00"}
{"citation_id": "19930082447", "source_url": "https://ntrs.nasa.gov/api/citations/19930082447/downloads/19930082447.pdf", "page_number": 13, "total_pages": 24, "image_filename": "19930082447_p13.jpg", "text": "NACA TN No. 1775\n\n11\n\nTABLE II\n\nIMPACT-LOADS DATA FROM TESTS OF A PRISMATIC FLOAT\n\nWITH 40° ANGLE OF DEAD RISE\n\n| Run | At contact | | | Approach parameter | At n₁ₘₐₓ | | | | Time, t, at chine immersion (sec) | Time, t, at bow immersion (sec) | At jₘₐₓ | | | At rebound | |\n|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|\n| | j₀ (fps) | j̇₀ (fps) | γ₀ (deg) | κ | t (sec) | n₁w (g) | j (ft) | j̇ (fps) | | | t (sec) | j (ft) | n₁w (g) | t (sec) | j̇ (fps) |\n| | | | | | | | | | τ = 3° | | | | | | |\n| 1 | 9.24 | 23.31 | 21.62 | 0.129 | 0.082 | 1.60 | 0.68 | 7.32 | 0.085 | 0.035 | 0.377 | 1.49 | 0.32 | ----- | ----- |\n| 2 | 7.75 | 23.26 | 18.43 | .154 | .103 | 1.17 | .74 | 5.97 | None | .040 | .396 | 1.39 | .25 | ----- | ----- |\n| 3 | 9.39 | 29.82 | 17.16 | .165 | .098 | 1.76 | .78 | 7.04 | .315 | .033 | .344 | 1.46 | .32 | No exit | No exit |\n| 4 | 8.82 | 29.07 | 16.88 | .170 | .077 | 1.70 | .69 | 7.39 | .337 | .035 | .356 | 1.45 | .25 | -do- | Do. |\n| 5 | 7.61 | 28.74 | 14.83 | .195 | .100 | 1.26 | .71 | 5.69 | None | .038 | .370 | 1.33 | .15 | -do- | Do. |\n| 6 | 7.82 | 29.94 | 14.64 | .197 | .093 | 1.30 | .72 | 5.90 | -do- | .038 | .360 | 1.30 | .32 | -do- | Do. |\n| 7 | 9.39 | 40.32 | 13.11 | .222 | .098 | 1.99 | .83 | 6.26 | -do- | .038 | .309 | 1.29 | .28 | -do- | Do. |\n| 8 | 8.75 | 39.53 | 12.48 | .233 | .097 | 1.82 | .73 | 6.47 | -do- | .035 | .312 | 1.30 | .25 | -do- | Do. |\n| 9 | 7.89 | 40.32 | 11.07 | .264 | .094 | 1.48 | .68 | 6.19 | -do- | .041 | .322 | 1.20 | .32 | -do- | Do. |\n| 10 | 4.05 | 23.15 | 9.92 | .296 | .180 | .41 | .66 | 2.70 | -do- | .142 | .455 | .96 | .20 | -do- | ----- |\n| 11 | 9.46 | 56.50 | 9.51 | .309 | .082 | 2.17 | .72 | 6.90 | -do- | .032 | .247 | 1.16 | .32 | 1.067 | -0.57 |\n| 12 | 5.83 | 40.32 | 8.23 | .359 | .128 | .87 | .63 | 4.05 | -do- | .058 | .353 | 1.00 | .25 | No exit | No exit |\n| 13 | 8.11 | 56.50 | 8.17 | .361 | .102 | 1.70 | .75 | 5.55 | -do- | .042 | .266 | 1.08 | .36 | -do- | Do. |\n| 14 | 7.47 | 54.35 | 7.83 | .377 | .098 | 1.52 | .67 | 5.26 | -do- | .040 | .268 | 1.04 | .28 | -do- | Do. |\n| 15 | 9.39 | 69.93 | 7.65 | .386 | .080 | 2.37 | .69 | 6.97 | -do- | .030 | .232 | 1.06 | .38 | .871 | -.85 |\n| 16 | 3.13 | 23.36 | 7.63 | .387 | .213 | .25 | .57 | 1.99 | -do- | .104 | .488 | .85 | .08 | No exit | No exit |\n| 17 | 4.05 | 30.30 | 7.61 | .389 | .162 | .45 | .58 | 2.59 | -do- | .084 | .420 | .94 | .19 | -do- | Do. |\n| 18 | 8.03 | 68.49 | 6.69 | .443 | .097 | 1.76 | .69 | 5.33 | -do- | .036 | .242 | .97 | .38 | -do- | Do. |\n| 19 | 5.97 | 56.50 | 6.03 | .492 | .117 | 1.05 | .64 | 4.12 | -do- | .052 | .286 | .89 | .32 | -do- | Do. |\n| 20 | 2.84 | 27.32 | 5.93 | .501 | .252 | .27 | .62 | 1.49 | -do- | .110 | .473 | .77 | .08 | -do- | Do. |\n| 21 | 3.13 | 30.30 | 5.90 | .503 | .213 | .27 | .58 | 1.99 | -do- | .102 | .428 | .77 | .15 | -do- | Do. |\n| 22 | 4.05 | 39.37 | 5.87 | .506 | .165 | .50 | .59 | 2.70 | -do- | .076 | .360 | .81 | .28 | -do- | Do. |\n| 23 | 3.98 | 40.16 | 5.66 | .525 | .161 | .50 | .58 | 2.70 | -do- | .076 | .366 | .83 | .15 | -do- | Do. |\n| 24 | 3.84 | 39.21 | 5.59 | .531 | .170 | .45 | .56 | 2.49 | -do- | .086 | .401 | .79 | .10 | -do- | Do. |\n| 25 | 7.82 | 91.50 | 5.11 | .582 | .084 | 2.03 | .59 | 5.69 | -do- | .031 | .197 | .86 | .63 | .612 | -1.71 |\n| 26 | 5.69 | 68.49 | 4.75 | .626 | .120 | 1.14 | .59 | 3.77 | -do- | .053 | .260 | .79 | .32 | No exit | No exit |\n| 27 | 2.84 | 39.06 | 4.16 | .716 | .248 | .30 | .58 | 1.28 | -do- | .116 | .408 | .68 | .08 | -do- | Do. |\n| 28 | 3.77 | 54.64 | 3.29 | .794 | .163 | .57 | .59 | 2.20 | -do- | .073 | .321 | .73 | .15 | -do- | Do. |\n| 29 | 5.69 | 90.10 | 3.61 | .826 | .111 | 1.26 | .58 | 3.56 | -do- | .054 | .228 | .73 | .50 | -do- | Do. |\n| 30 | 4.27 | 68.03 | 3.59 | .830 | .148 | .72 | .55 | 2.70 | -do- | .074 | .278 | .69 | .31 | -do- | Do. |\n| 31 | 3.20 | 56.18 | 3.26 | .915 | .183 | .41 | .52 | 1.85 | -do- | .094 | .350 | .66 | .20 | -do- | Do. |\n| 32 | 3.56 | 68.03 | 3.00 | .955 | .184 | .57 | .56 | 1.78 | -do- | .087 | .284 | .63 | .40 | -do- | Do. |\n| 33 | 3.20 | 68.49 | 2.68 | 1.114 | .185 | .50 | .53 | 1.78 | -do- | .107 | .314 | .61 | .15 | -do- | Do. |\n| 34 | 4.12 | 90.10 | 2.64 | 1.140 | .141 | 1.01 | .51 | 1.42 | -do- | .069 | .241 | .63 | .45 | .710 | -0.57 |\n| 35 | 3.63 | 90.09 | 2.31 | 1.293 | .154 | .75 | .50 | 1.85 | -do- | .088 | .247 | .58 | .35 | No exit | No exit |\n| 36 | 2.77 | 90.91 | 1.75 | 1.708 | .193 | .49 | .44 | 1.14 | -do- | .078 | .277 | .48 | .15 | -do- | Do. |\n| (a) | 9.02 | 89.88 | 5.73 | .520 | .083 | 2.59 | .67 | 5.90 | -do- | .037 | .192 | .95 | .72 | .756 | -1.70 |\n| | | | | | | | | | τ = 6° | | | | | | |\n| 37 | 8.89 | 34.13 | 14.60 | 0.388 | 0.107 | 1.72 | 0.89 | 6.26 | -do- | 0.072 | 0.296 | 1.35 | 0.32 | 1.085 | -1.07 |\n| 38 | 8.75 | 34.60 | 14.19 | .400 | .108 | 1.85 | .91 | 5.90 | -do- | .068 | .293 | 1.33 | .32 | 1.044 | -1.28 |\n| 39 | 8.89 | 35.21 | 14.17 | .404 | .094 | 1.76 | .81 | 6.33 | -do- | .068 | .284 | 1.30 | .36 | .975 | -1.35 |\n| 40 | 9.03 | 43.47 | 11.74 | .489 | .100 | 1.98 | .85 | 6.33 | -do- | .067 | .253 | 1.22 | .50 | .836 | -2.28 |\n| 41 | 8.89 | 43.29 | 11.61 | .495 | .102 | 1.99 | .85 | 6.04 | -do- | .071 | .253 | 1.22 | .45 | .849 | -1.71 |\n| 42 | 8.75 | 44.85 | 10.19 | .515 | .107 | 1.98 | .84 | 6.11 | -do- | .062 | .252 | 1.18 | .50 | .839 | -1.71 |\n| 43 | 8.89 | 46.73 | 10.77 | .536 | .094 | 2.03 | .81 | 6.47 | -do- | .069 | .234 | 1.18 | .58 | .772 | -1.99 |\n| 44 | 8.89 | 58.14 | 8.69 | .669 | .100 | 2.29 | .81 | 5.90 | -do- | .071 | .209 | 1.08 | .75 | .602 | -2.56 |\n| 45 | 8.96 | 58.47 | 8.46 | .688 | .100 | 2.33 | .81 | 5.76 | -do- | .063 | .209 | 1.07 | .80 | .612 | -2.42 |\n| 46 | 8.96 | 60.97 | 8.36 | .697 | .101 | 2.29 | .82 | 5.8", "timestamp": "2026-07-22T04:39:32.477193+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 44, "total_pages": 54, "image_filename": "19930086015_p44.jpg", "text": "NACA RM A9E24\n\nCONFIDENTIAL\n\n| | | |\n| :--- | :--- | :--- |\n| $D=165.12$<br>$M=1.23$<br>$\\Delta p_t/q = .17$ | $D=147.20$<br>$M=1.32$<br>$\\Delta p_t/q = .26$ | $D=129.32$<br>$M=1.43$<br>$\\Delta p_t/q = .31$ |\n| $D=113.43$<br>$M=1.53$<br>$\\Delta p_t/q = .31$ | $D=100.47$<br>$M=1.63$<br>$\\Delta p_t/q = .31$ | $D=88.52$<br>$M=1.73$<br>$\\Delta p_t/q = .32$ |\n\nDistance from wall, $\\delta$, in.\nRatio of total pressures, $\\frac{H}{H_0}$\n\n(b) Lower front rake.\n\nFigure 12—Continued.\n\nCONFIDENTIAL\n\n43", "timestamp": "2026-07-22T04:39:35.305993+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 4, "total_pages": 99, "image_filename": "19930082511_p4.jpg", "text": "2\nNACA TN No. 1826\n\nThe present examination of the problem was occasioned by the need for\nboundary corrections for tests in the Langley full-scale tunnel of a\nlarge helicopter, of which the forward edge of the rotor disk reached\nalmost to the mouth of the entrance bell while the rear edge approached\nthe exit bell. Previous studies (reference 2) had shown that the\nPrandtl theory was satisfactory for a wing in the usual position in the\ntunnel (about 20 feet downstream of the entrance); but it was felt that\nthis simple theory was inadequate for such far forward and rearward\nlocations of the lifting surface, and that some further development was\ndesirable. The only previous analysis bearing directly on the problem\nseemed to be that of reference 3, which considered a lifting element\nconcentrated at a point on the axis of a circular open tunnel of finite\njet length; however, the treatment therein was not rigorous, and was\njustified only by a somewhat heuristic discussion, so that its general\napplicability was not obvious. Other studies treated either two-\ndimensional or axially symmetrical conditions (references 4 and 5) and\nalso did not consider the closed exit region, so that the extent of\ntheir applicability to the present problem was not at first apparent. A\nsimilar German wartime report (reference 6), which did not become available\nuntil after the present paper was written, would have been more useful\nin this respect because of the generality of its physical discussion.\n\nBecause of the particular shape of the tunnel cross section, a\nreasonably simple solution in terms of available functions seemed unlikely;\naccordingly, the initial effort was directed toward defining the problem\nin such a way that it could be solved by analogy methods in an electrical\ntank. Identification of the necessary boundary conditions appeared at\nfirst to be somewhat perplexing; however, after recognition of some of\nthe basic physical phenomena, the boundary conditions were readily\nclarified. The problem is thus considered now to be fairly well under-\nstood, at least insofar as it can be considered linear and uninfluenced\nby turbulent mixing at the free surfaces or by the irregular nature of\nthe flow at the exit. As will appear later, however, grave technical\ndifficulties exist in the exact solution by electrical-analogy methods,\nso that, for example, actual evaluation of the tunnel interference for\nthe large helicopter in the Langley full-scale tunnel, which problem\ninstigated the present research, has not yet been accomplished.\n\nAfter the boundary conditions were clarified, analytical methods of\nsolution were developed for two-dimensional and circular open tunnels.\nThese studies have been combined with the discussion of the boundary\nconditions and the electrical analogies to form the present paper, which,\nit is hoped, will serve to clarify basic concepts and establish a sound\nbasis for any further work.\n\nThe report is divided into three parts. In part I, the boundary\nconditions are defined and discussed for the open wind tunnel with closed\nentrance and exit sections, and an outline is given of suggested electri-\ncal analogies applicable to the problem. In part II, analytical\nsolutions are given for various two-dimensional open-tunnel types,", "timestamp": "2026-07-22T04:39:36.979535+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 8, "total_pages": 33, "image_filename": "19930082487_p8.jpg", "text": "6\nNACA TN No. 1813\n\nwith Mach number at various chordwise stations for the NACA 23015\nairfoil section at several angles of attack. The local pressure-\ncoefficient variation with Mach number at the crest location, for\neach angle of attack shown in figure 4, is denoted by a solid\ncurve. It is seen that for each angle of attack the drag-divergence\nMach number differs only slightly from the free-stream Mach number\nat which sonic velocity occurs at the airfoil crest. As the Mach\nnumber is increased beyond the value for drag divergence there is a\nrise in pressure coefficient with Mach number at points ahead of the\ncrest; and, at least up to a Mach number of 0.8, this pressure-\ncoefficient variation with free-stream Mach number is such as to\nmaintain practically constant local Mach numbers.\n\nOrigin of Drag Rise\n\nThe relation between the pressure-distribution changes noted for\nthe NACA 23015 airfoil section and the supercritical drag rise will\nnow be considered. At approximately the drag-divergence Mach number,\nthere is established over the forward portion of the upper surface of\nthe NACA 23015 airfoil a Mach number distribution which remains rela-\ntively unchanged with some further increase of free-stream Mach\nnumber. In figure 3, it is seen that, while pressure coefficients\nahead of the crest are becoming less negative, pressure coefficients\nat points well aft of the airfoil crest become more negative as the\nMach number increases above that for drag divergence. This combina-\ntion of pressure changes produces a rising pressure drag which appears\nto be the primary factor in the supercritical drag rise. It should be\npointed out that the negatively increasing pressure coefficients aft\nof the airfoil crest may result from two different causes: the\nrearward growth of the supersonic region, and marked boundary-layer\nthickening. Both of these causes may be present, as for example, on\nthe NACA 23015 airfoil section at $2^\\circ$ angle of attack.\n\nPrediction of the Drag-Divergence Mach Number\n\nIn the preceding sections the flow changes over the NACA 23015\nairfoil section at supercritical speeds were considered for various\nangles of attack. It was noted that increasing the free-stream Mach\nnumber produces no abrupt changes in the airfoil section character-\nistics until the supersonic region extends behind the airfoil crest.\nWhen this occurs, the drag begins to rise rapidly with increasing\nMach number. Figure 5 shows that for the NACA 23015 airfoil section\nthe calculated and experimental values of $M_p$, the Mach number at\nwhich sonic velocity occurs at the airfoil crest, are in reasonable", "timestamp": "2026-07-22T04:39:46.794489+00:00"}