PDF/ocr_dataset_1081.jsonl

{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 18, "total_pages": 30, "image_filename": "19930085970_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:55:07.020396+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 68, "total_pages": 72, "image_filename": "19930085491_p68.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 67\n\n(a) WF-57; $C_L$, 0.\n\n(b) WF-57; $C_L$, 0.27.\n\n(c) WF-60; $C_L$, 0.\n\n(d) WF-60; $C_L$, 0.26.\n\nFigure 16.— Schlieren photographs of swept-back wing and fuselage configurations at a Reynolds number of 0.62 million.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:07.272354+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 14, "total_pages": 25, "image_filename": "19930085979_p14.jpg", "text": "NACA RM E9E12\n13\n\n[Figure: Photograph of a man crouching next to a large, metallic, cylindrical model mounted on a stand inside a tunnel test section.]\n\nNACA\nC-22171\n8-31-48\n\nFigure 1. - Photograph of model installation in tunnel test section.", "timestamp": "2026-07-22T06:55:14.692348+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 2, "total_pages": 20, "image_filename": "19930085999_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:55:15.975291+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 14, "total_pages": 46, "image_filename": "19930085983_p14.jpg", "text": "```markdown\nTABLE I.- AIRFOIL-SECTION COORDINATES\n[All values given in percent chord]\n\n| $c_0$ | | | | $c_1$ | | | | $c_2$ | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Lower surface** | | **Upper surface** | | **Lower surface** | | **Upper surface** | | **Lower surface** | | **Upper surface** | |\n| Station | Ordinate | Station | Ordinate | Station | Ordinate | Station | Ordinate | Station | Ordinate | Station | Ordinate |\n| 0 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 0 | 0 |\n| .5 | -.404 | .5 | .404 | .526 | -.366 | .474 | .438 | .532 | -.357 | .469 | .457 |\n| .75 | -.488 | .75 | .488 | .778 | -.433 | .726 | .536 | .788 | -.419 | .713 | .557 |\n| 1.25 | -.616 | 1.25 | .616 | 1.283 | -.536 | 1.221 | .690 | 1.295 | -.507 | 1.208 | .720 |\n| 2.50 | -.847 | 2.50 | .847 | 2.535 | -.706 | 2.463 | .984 | 2.547 | -.663 | 2.447 | 1.026 |\n| 5.00 | -1.166 | 5.00 | 1.166 | 5.039 | -.927 | 4.956 | 1.401 | 5.056 | -.851 | 4.944 | 1.477 |\n| 10.0 | -1.599 | 10.0 | 1.599 | 10.041 | -1.211 | 9.959 | 1.984 | 10.056 | -1.089 | 9.944 | 2.103 |\n| 20.0 | -2.131 | 20.0 | 2.131 | 20.036 | -1.530 | 19.964 | 2.725 | 20.044 | -1.345 | 19.956 | 2.912 |\n| 30.0 | -2.413 | 30.0 | 2.413 | 30.026 | -1.682 | 29.974 | 3.138 | 30.031 | -1.458 | 29.969 | 3.373 |\n| 40.0 | -2.499 | 40.0 | 2.499 | 40.010 | -1.700 | 39.990 | 3.297 | 40.013 | -1.446 | 39.987 | 3.548 |\n| 50.0 | -2.354 | 50.0 | 2.354 | 50.00 | -1.530 | 50.000 | 3.179 | 50.000 | -1.270 | 50.000 | 3.436 |\n| 60.0 | -2.032 | 60.0 | 2.032 | 59.990 | -1.231 | 60.010 | 2.828 | 59.987 | -.982 | 60.013 | 3.085 |\n| 70.0 | -1.589 | 70.0 | 1.589 | 69.985 | -.860 | 70.015 | 2.313 | 69.981 | -.632 | 70.019 | 2.547 |\n| 80.0 | -1.071 | 80.0 | 1.071 | 79.985 | -.479 | 80.015 | 1.669 | 79.975 | -.288 | 80.025 | 1.852 |\n| 90.0 | -.541 | 90.0 | .541 | 89.985 | -.155 | 90.015 | .927 | 89.981 | -.031 | 90.019 | 1.045 |\n| 100.0 | -.011 | 100.0 | .011 | 100.000 | --- | 100.000 | --- | 100.000 | --- | 100.000 | --- |\n\n| $c_3$ | | | | $c_4$ | | | | $c_5$ | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Lower surface** | | **Upper surface** | | **Lower surface** | | **Upper surface** | | **Lower surface** | | **Upper surface** | |\n| Station | Ordinate | Station | Ordinate | Station | Ordinate | Station | Ordinate | Station | Ordinate | Station | Ordinate |\n| 0.533 | -0.350 | 0.470 | 0.462 | 0.536 | -0.350 | 0.470 | 0.460 | 0.543 | -0.350 | 0.473 | 0.455 |\n| .788 | -.414 | .709 | .577 | .789 | -.405 | .711 | .558 | .788 | -.420 | .718 | .560 |\n| 1.298 | -.502 | 1.202 | .725 | 1.291 | -.503 | 1.204 | .722 | 1.278 | -.508 | 1.208 | .718 |\n| 2.555 | -.645 | 2.444 | 1.043 | 2.549 | -.646 | 2.440 | 1.039 | 2.557 | -.648 | 2.452 | 1.033 |\n| 5.056 | -.828 | 4.944 | 1.497 | 5.057 | -.832 | 4.945 | 1.489 | 5.061 | -.823 | 4.956 | 1.494 |\n| 10.056 | -1.051 | 9.944 | 2.150 | 10.055 | -1.050 | 9.945 | 2.144 | 10.070 | -1.048 | 9.965 | 2.119 |\n| 20.048 | -1.290 | 19.952 | 2.978 | 20.044 | -1.291 | 19.956 | 2.976 | 20.053 | -1.313 | 19.947 | 2.960 |\n| 30.032 | -1.377 | 29.968 | 3.447 | 30.033 | -1.389 | 29.967 | 3.425 | 30.035 | -1.419 | 29.965 | 3.415 |\n| 40.016 | -1.361 | 39.984 | 3.639 | 40.022 | -1.368 | 39.978 | 3.621 | 40.018 | -1.401 | 39.982 | 2.608 |\n| 50.000 | -1.186 | 50.000 | 3.527 | 50.000 | -1.193 | 50.000 | 3.512 | 50.000 | -1.208 | 50.000 | 3.485 |\n| 59.984 | -.892 | 60.016 | 3.169 | 59.989 | -.908 | 60.011 | 3.162 | 59.982 | -.928 | 60.018 | 3.135 |\n| 69.976 | -.557 | 70.024 | 2.627 | 69.978 | -.569 | 70.022 | 2.604 | 69.982 | -.595 | 70.018 | 2.592 |\n| 79.976 | -.223 | 80.024 | 1.911 | 79.978 | -.230 | 80.022 | 1.915 | 79.982 | -.245 | 80.018 | 1.891 |\n| 89.984 | -.008 | 90.016 | 1.091 | 89.978 | -.011 | 90.022 | 1.083 | 89.982 | -.018 | 90.018 | 1.068 |\n| 100.000 | --- | 100.000 | --- | 100.000 | --- | 100.000 | --- | 100.000 | --- | 100.000 | --- |\n\nNote: Spanwise positions of airfoil sections $c_0$ to $c_5$ are shown in Figure 3.\nFor all sections: Leading-edge radius = 0.175. Trailing-edge radius = 0.014.\n\n[Figure: NACA logo]\n\nCONFIDENTIAL\n12\nCONFIDENTIAL\nNACA RM A9I27\n```", "timestamp": "2026-07-22T06:55:18.008598+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 12, "total_pages": 132, "image_filename": "19930085990_p12.jpg", "text": "```markdown\n10\nCONFIDENTIAL\nNACA RM A9I01\n\nMounting the tail above the fuselage necessitated a supporting\nbracket with a streamlined body to serve as a fairing for the fittings\nby which the stabilizer was attached. The force and moment characteristics\nof the wing and fuselage with the bracket and the fairing body are presented\nin figures 16, 17, 18, and 19. These data indicate no noticeable effects\nof the bracket on the characteristics of the wing-fuselage combination\nexcept a slight increase in the minimum drag. (See fig. 17.)\n\nLift, drag, and pitching-moment characteristics of the complete\nsemispan model with the horizontal tail mounted above the extended wing-\nchord plane are presented in figures 20, 21, and 22 for Mach numbers up\nto 0.95 and for stabilizer settings from $4^\\circ$ to $-6^\\circ$. Comparison of the drag\ndata of figure 21 with those of figure 11 indicates a slight increase in\nthe minimum drag which may be attributed to the addition of the tail\nbracket and the fairing body and not to the raising of the horizontal tail.\nThe model with the high tail was longitudinally stable at all lift\ncoefficients below the stall and at all Mach numbers, as can be seen from\nfigure 22. At a Mach number of 0.20, addition of the horizontal tail\nshifted the aerodynamic center from 14 percent to 53 percent of the wing\nmean aerodynamic chord. The contribution of the horizontal tail to the\nlongitudinal stability decreased with increasing Mach number. As will be\ndiscussed later, this reduction in the contribution of the tail to the\nstability was due primarily to an increase in $\\partial\\epsilon/\\partial\\alpha$ with increasing Mach\nnumber. The all-movable stabilizer retained effectiveness in longitudinal\ncontrol at all Mach numbers and all lift coefficients.\n\nThere was a marked change in the pitching-moment coefficient at zero\nlift as a result of raising the tail above the fuselage. Whereas with the\ntail in the extended wing-chord plane, zero pitching moment occurred at\nzero lift with a stabilizer angle of $0^\\circ$, with the tail raised above the\nextended wing-chord plane a stabilizer setting of approximately $2^\\circ$ was\nrequired to produce zero pitching moment at zero lift. To investigate the\ncause of this shift in the zero-lift pitching-moment coefficient the\nReynolds number was increased from 2,000,000 to 12,000,000 while the Mach\nnumber remained 0.20. This increase had no effect on the pitching-moment\ncoefficient at zero lift. Visual observation, by means of tufts, of the\nflow at the afterend of the fuselage and on the tail-supporting bracket\nrevealed a sizable stream angle in the region of the tail due to the rapid\nconvergence of the rear end of the fuselage. This convergence was reduced\nby modifying the afterpart of the fuselage as shown in figure 1(b). The\nresults of tests with the modified fuselage are shown in figure 23. These\ndata show that, for the model with the tail mounted above the extended\nwing-chord plane, modification of the fuselage caused a decrease in the\nzero-lift pitching-moment coefficient greater than the increase accompanying\nthe raising of the tail on the original fuselage.\n\nThe lift, drag, and pitching-moment characteristics of the complete\nsemispan model with the high tail and the original fuselage and with the\nwing flaps deflected are presented in figures 24, 25, and 26. Raising\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:55:18.971347+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 17, "total_pages": 30, "image_filename": "19930085975_p17.jpg", "text": "$$5 \\times 1 0 ^ { 6 }$$\n\nCONFIDENTIAL\n\nReynolds number, R\n\nMach number,M\n\nCONFIDENTIAL\n\nFigure 4.- The variation of test Reynolds number with Mach number based on the mean aerodynamic chord of 0.765 foot.\n\nNACA RM L9E10", "timestamp": "2026-07-22T06:55:21.856536+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 9, "total_pages": 17, "image_filename": "19930085988_p9.jpg", "text": "CONFIDENTIAL\n\nWings:\nUsed on wing-body\nconfiguration only\n\nMAC = 23.9\n\n6.46 Diam.\n\nFins:\n(2) shown for\nwinged model\n(4) used on wing-\nless model\n\nNACA\n\nA-A\nTypical fin section\n\nFigure 1.— General arrangement of test model. Wing-body configuration shown. Body-alone configuration\nidentical except as noted.\n\nNACA RM L9H30", "timestamp": "2026-07-22T06:55:22.110045+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 59, "total_pages": 149, "image_filename": "19930083192_p59.jpg", "text": "NACA TN 1976\n55\n\ndifferent routes in the U.S. and abroad and had an average value of 0.1.\nSince, on the average, 500/6 gusts with an effective gust velocity greater\nthan 0.3 foot per second are encountered per mile of rough air (reference 9), the number of gusts that will be encountered on the average\nduring the lifetime of an airplane can be estimated. No path-ratio\ndata are available for postwar operations. Attempts to obtain such data\nfrom the indirect evidence of V-G records lead to erroneous results.\n\nDISCUSSION\n\nApplied acceleration increments.- For the prewar period, table XXI\nshows a wide scatter in the flight miles to exceed the limit acceleration increment. The values vary by a factor of about 300. A variation\nby a factor of 42 occurs for the same type of airplane operated on\ndifferent routes by different airlines. In four out of the six samples,\nthe flight miles to exceed the limit acceleration increment would be\ngreater than 13 million, or, if a cruising speed of 200 miles per hour\nis assumed, it might be expected that the limit load factor would be\nexceeded about once on the average in every 60,000 hours of flight.\n\nThe data for the wartime period do not show as much variation of\nflight miles and the variations do not appear significant within that\ngroup. In comparison with the prewar operations, however, the flight\nmiles to exceed limit acceleration have consistently decreased; this\nfact indicates that the pressure of the emergency on wartime commercial\noperations resulted in higher imposed loads on the airplane.\n\nAtmospheric gustiness.- The flight miles to exceed the acceleration\nincrement corresponding to an effective gust velocity of 37.5K feet per\nsecond at the probable speed (table XXI) indicate that the operational\nexperience on the basis of a \"gust intensity\" have much less scatter.\nThe spread in flight miles was 15:1 for the prewar operations and\nabout 7:1 for the wartime period. For prewar conditions, when the\ndata for airplane D on route V are neglected, the scatter is within the\narbitrary 5:1 criterion and, in comparison with the spread in flight\nmiles for limit acceleration increment, indicates that the difference\nin the level of roughness encountered on various routes is not of\nengineering concern. A similar observation is indicated for the wartime\nperiod although the smaller average flight miles indicates that flights\nduring that period were through more severe weather conditions than for\nthe prewar period.\n\nIn connection with prewar data for airplane D on route V of\ntable XXI, the low value of flight miles shown may be significant and\nindicates that early transpacific flights encountered more severe\nweather than was experienced along other routes. When the distances", "timestamp": "2026-07-22T06:55:22.561792+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 9, "total_pages": 40, "image_filename": "19930085997_p9.jpg", "text": "NACA RM A9I29 CONFIDENTIAL 7\n\nThe total-pressure surveys at position 2 were made in a section of the duct where the computed average Mach number was approximately 0.5, and these measurements represent more closely the pressure recovery that would exist at the compressor intake. The variation of $(\\overline{H}_2/H_0)_{\\text{max}}$ with $M_0$ for configuration D is shown in figure 5. The losses in total-pressure ratio between positions 2 and 3 amounted to from 0.015 to 0.030.\n\nSince the nose-type duct inlet is generally accepted, at present, as a design in which the highest pressure recovery can be realized, the pressure recovery of typical nose inlet models (reference 4) is also shown in figure 5. A comparison of these results with those of the present model indicates that, at Mach numbers less than about 1.8, total-pressure recovery within 0.05 of that of nose inlets was attained with the present design, without considering energy expended in removing boundary-layer air.\n\nEffect of mass-flow ratio $m_1/m_0$ on pressure recovery.— The variation of $H_3/H_0$ and $\\overline{H}_2/H_0$ with $m_1/m_0$ is presented in figure 7 for configuration D at an angle of attack of $0^\\circ$. In the range of mass-flow ratios indicated by the dashed curves, schlieren photographs demonstrated that the boundary layer ahead of the scoops was separated. The fact that the total-pressure ratio $H_3/H_0$ remained high at 1.36 Mach number possibly was caused by a condition in which the losses ahead of the inlet were compensated for by reduced losses within the subsonic diffusers at these low mass-flow ratios. Apparently, as the Mach number increased the energy dissipated in turbulence ahead of the inlet increased and caused reduced values of $H_3/H_0$ noted at the higher Mach numbers.\n\nMaximum values of $\\overline{H}_2/H_0$ occurred at larger mass-flow ratios than those at which $(H_3/H_0)_{\\text{max}}$ was recorded. This difference can be attributed to the subsonic diffuser efficiency between the two positions. The attainment of a constant rate of mass flow through the air-induction system indicates that supersonic flow into the inlet has been established.\n\nEffect of mass-flow ratio $m_4/m_0$ on pressure recovery.— The previously discussed results were obtained with the maximum rate of flow through the boundary-layer scoops. In tests of the model with inlet configuration D, reductions in the mass of air flowing through the boundary-layer scoops influenced the recovery of total pressure in the settling chamber as shown in figure 8. A nearly linear variation of $(H_5/H_0)_{\\text{max}}$ with $m_4/m_0$ occurred at free-stream Mach numbers of 1.36 and 1.70. However, for $M_0$ equal to 2.01, a reduction in $m_4/m_0$ from the value at which the flow in the boundary-layer duct was choked caused the main-scoop flow to be unsteady and $H_5/H_0$ to decrease markedly.\n\nFor configuration D with choked flow in the boundary-layer ducts, the mass-flow ratio $m_4/m_0$ of the air entering the boundary-layer scoops and the total-pressure recovery $H_5/H_0$ in the sting are presented in figure 9 as functions of $m_1/m_0$. At mass-flow ratios at which separated flow occurred ahead of the main scoops (see fig. 7), the reduced pressure\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:24.068492+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 25, "total_pages": 34, "image_filename": "19930085913_p25.jpg", "text": "24\nNACA RM L9F24\n\n| | 1 | 2 | 3 | 4 |\n| :--- | :--- | :--- | :--- | :--- |\n| **82** | 50 | 50 | 20 | 20 |\n| **83** | 30 | 30 | 15 | 15 |\n| **84** | 30 | 30 | 15 | 15 |\n| **85** | 30 | 30 | 20 | 20 |\n| **86** | 30 | 30 | 15 | 15 |\n| **87** | 30 | 30 | 15 | 15 |\n| **88** | 30 | 30 | 15 | 15 |\n| **89** | 30 | 30 | 15 | 15 |\n| **90** | 30 | 30 | 15 | 15 |\n| **91** | 30 | 30 | 15 | 15 |\n| **92** | 30 | 30 | 15 | 15 |\n| **93** | 50 | 30 | 15 | 15 |\n| **94** | 30 | 30 | 20 | 20 |\n| **95** | 30 | 30 | 15 | 30 |\n\n[Figure: NACA logo]\n\n(r) Model C; $\\Lambda = 60^\\circ$; $e_w = 0$.\nFigure 1.— Concluded.", "timestamp": "2026-07-22T06:55:25.416963+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 9, "total_pages": 24, "image_filename": "19930085991_p9.jpg", "text": "NACA RM L9I28\n\nwhere $C_{m_\\alpha}$ was determined graphically. This equation is similar to one developed and applied to a jettisonable nose section in reference 2 based on the relationship from reference 3\n\n$$\n\\Delta F = q \\frac{dA}{dx} \\cos^2 \\epsilon \\sin 2\\alpha\n$$\n\nwhere $\\Delta F$ can be taken as a measure of the transverse or normal force per unit of length for a symmetrical airship hull.\n\nThe stabilizing effectiveness of the fin system was calculated by the following general equation:\n\n$$\nC_{m_\\alpha \\text{fins}} = -2 C_{L_\\alpha \\text{fins}} \\cos^2 \\theta \\frac{S_{F_t} L_T}{S_p L} - 2 C_{L_\\alpha \\text{fins}} \\cos^2 \\phi \\frac{S_{F_t} L_T}{S_p L}\n$$\n\nwhere $\\theta$ is the angle between the plane of any two of the fins and the axis about which $C_{m_\\alpha}$ is calculated and $\\phi$ is the angle between the plane of the other two fins and this axis. The angle $\\theta$ plus $\\phi$ equals $90^\\circ$. Therefore,\n\n$$\nC_{m_\\alpha \\text{fins}} = -2 C_{L_\\alpha \\text{fins}} \\frac{S_{F_t} L_T}{S_p L} (\\cos^2 \\theta + \\cos^2 \\phi)\n$$\n\nand since\n\n$$\n\\theta = 90 - \\phi\n$$\n\nand\n\n$$\n\\cos \\phi = \\sin \\theta\n$$", "timestamp": "2026-07-22T06:55:25.776426+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 31, "total_pages": 47, "image_filename": "19930085919_p31.jpg", "text": "30\nCONFIDENTIAL\nNACA RM No. A9C21\n\n[Figure: Graph showing Lift coefficient, $C_L$ vs Drag coefficient, $C_D$ for three flap deflection angles: $\\delta_f, 45^\\circ$; $\\delta_f, 60^\\circ$; and $\\delta_f, 75^\\circ$. The graph includes schematic diagrams of wing configurations labeled \"Plain wing + short fuselage\" with dimensions such as 0.45c, 0.60c, 0.75c, 1.05c, and 0.25c. The NACA logo is present in the lower right corner of the graph area.]\n\n(b) $C_L$ vs $C_D$.\nFigure 9- Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:26.938178+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 29, "total_pages": 92, "image_filename": "19930085951_p29.jpg", "text": ".48\n.44\n.40\n.36\n.32\n.28\n.24\n.20\n.16\n.12\n.08\n.04\n0\nPower coefficient, $C_P$\n\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0 3.2 3.4 3.6 3.8\nAdvance ratio, J\n\n$\\beta_{0.75R}$ 20° 25° 30° 35° 40° 45° 50° NACA 55°\n\n(b) Power coefficient.\nFigure 8.— Continued. Rotational speed, 1140 rpm.\n\nNACA RM L9D29\n27", "timestamp": "2026-07-22T06:55:27.145718+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 69, "total_pages": 72, "image_filename": "19930085491_p69.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:55:28.193317+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 50, "total_pages": 92, "image_filename": "19930085870_p50.jpg", "text": "```markdown\nNACA RM No. L9D07\n51\n\nCONFIDENTIAL\n\n<!-- Image (66, 110, 903, 896) -->\n\n(j) Wing 10.w=1.371; R=660,000.\nFigure 6. - Continued.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:55:29.279376+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 15, "total_pages": 25, "image_filename": "19930085979_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:55:30.638690+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 19, "total_pages": 30, "image_filename": "19930085970_p19.jpg", "text": "NACA RM A9E09 CONFIDENTIAL 17\n\n[Figure: (a) M = 0.]\n\n[Figure: (b) M = 1.09, side view.]\n\n[Figure: (c) M = 1.14, plan view.]\n\n[Figure: (d) M = 1.14, side view.]\n\nFigure 4.— Typical schlieren photographs of the flow about the model at supersonic Mach numbers.\n\nCONFIDENTIAL\n\nNACA\nA-13971", "timestamp": "2026-07-22T06:55:30.884178+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 98, "total_pages": 114, "image_filename": "19930086061_p98.jpg", "text": "94\nNACA RM L9J07\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n$$ \\frac{y}{b/2} \\text{, percent} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\circ$ 4.1 | 0.12 |\n| $\\square$ 8.1 | 0.27 |\n\n(a) Angles of attack: 4.1°, 8.1°.\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n$$ \\frac{y}{b/2} \\text{, percent} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\diamond$ 14.1 | 0.46 |\n| $\\triangle$ 24.1 | 0.69 |\n| $\\triangleright$ 32.1 | 0.83 |\n\n(b) Angles of attack: 14.1°, 24.1°, 32.1°.\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n$$ \\frac{y}{b/2} \\text{, percent} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\square$ 36.1 | 0.86 |\n| $\\circ$ 44.1 | 0.72 |\n\n(c) Angles of attack: 36.1°, 44.1°.\n\nFigure 43.- Span load distribution of wing 2 at various angles of attack; $\\psi = 20^\\circ$. Flagged symbols represent data taken with left semispan at $\\psi = -20^\\circ$.", "timestamp": "2026-07-22T06:55:31.429824+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 38, "total_pages": 52, "image_filename": "19930085911_p38.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 37\n\n1152\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:55:31.662014+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 19, "total_pages": 55, "image_filename": "19930085966_p19.jpg", "text": "18\nCONFIDENTIAL\nNACA RM L9B17\n\n$$\n\\frac{p_4}{p_3} = \\frac{\\gamma_3}{\\gamma_4} \\left[ \\frac{1 + \\gamma_3 \\left(1 - \\frac{K_3}{2}\\right) M_3^2}{\\frac{\\gamma_3}{\\gamma_4} + \\gamma_3 \\left(1 + \\frac{K_4}{2}\\right) M_4^2} \\right] \\tag{A1}\n$$\n\n$$\n\\frac{pt_4}{pt_3} = \\frac{p_4}{p_3} \\frac{\\left(1 + \\frac{\\gamma_4 - 1}{2} M_4^2\\right)^{\\frac{\\gamma_4}{\\gamma_4 - 1}}}{\\left(1 + \\frac{\\gamma_3 - 1}{2} M_3^2\\right)^{\\frac{\\gamma_3}{\\gamma_3 - 1}}} \\tag{A2}\n$$\n\n$$\n\\frac{Tt_4}{Tt_3} = \\frac{\\gamma_4}{\\gamma_3} \\left(\\frac{p_4}{p_3}\\right)^2 \\frac{R_3}{R_4} \\frac{M_4^2}{M_3^2} \\frac{\\left(1 + \\frac{\\gamma_4 - 1}{2} M_4^2\\right)}{\\left(1 + \\frac{\\gamma_3 - 1}{2} M_3^2\\right)} \\tag{A3}\n$$\n\nAn inspection of the equations reveals that a simultaneous solution is possible which expresses the total-pressure ratio as a function of Mach number before combustion, total-temperature ratio, friction coefficients, specific heats, and gas constants. Also, through use of equations (A1) and (A2) it is possible to express total-pressure ratio as a function of Mach numbers before and after combustion, friction coefficients, and specific heats. Figure 22 illustrates these functions for standard air values of the ratio of specific heats and the gas constant and approximately the value of friction loss and distribution corresponding to the test ram-jet combustion chamber. The plot assumes that the friction-loss coefficients remain constant for all conditions of combustion-chamber operation. A test-data plot, similar to that of figure 22, is presented in figure 23, which was used to determine the total-pressure ratio across the combustion chamber in the supersonic thrust-coefficient calculations. A comparison of the hypothetical combustion-chamber characteristics and the actual characteristics is made in figure 24. An inspection of the figure reveals that a closer comparison probably could have been attained by choosing a hypothetical combustion chamber with a slightly lower friction-loss coefficient concentrated more heavily at the combustion-chamber outlet.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:33.890022+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 10, "total_pages": 17, "image_filename": "19930085988_p10.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9H30\n\n(a) Body alone.\n(b) Wing-body configuration.\n\nFigure 2.— Photographs of models in launching position.\n\nCONFIDENTIAL\n\n9", "timestamp": "2026-07-22T06:55:35.847488+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 3, "total_pages": 20, "image_filename": "19930085999_p3.jpg", "text": "NACA RM E9I07\nRESTRICTED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nVIBRATION OF LOOSELY MOUNTED TURBINE BLADES DURING SERVICE\nOPERATION OF A TURBOJET ENGINE WITH CENTRIFUGAL COMPRESSOR\nAND STRAIGHT-FLOW COMBUSTION CHAMBERS\n\nBy W. C. Morgan, R. H. Kemp\nand S. S. Manson\n\nSUMMARY\n\nAn experimental investigation was conducted to determine the\nvibration characteristics of loosely mounted turbine blades during\nservice operation of a turbojet engine. High-temperature strain\ngages were used to measure the turbine-blade vibrations.\n\nThe turbine blades studied were 4 inches long; the experimental\ninvestigation was confined to blades having 0.03-inch and 0.06-inch\namplitude of tip movement at room temperature in the plane of the\nturbine wheel. Vibration during service operation occurred in the\nfirst bending and first torsional modes of the turbine blades. In\naddition, a small number of complex modes of vibration was observed.\nSources of vibration excitation were present at frequencies related\nto the number of nozzle blades and combustion chambers and to multi-\nples of the first-order turbine speed.\n\nA comparison was made between the vibrations of the loosely\nmounted blades and those observed during a previous investigation\nof similar blades tightly mounted in a turbine wheel. For the par-\nticular engine used, in which the vibratory-stress levels were gen-\nerally low, the comparison did not show the existence of any appre-\nciable vibration damping attributable to looseness in turbine-blade\nmounts. The loosely mounted blades appeared to be more susceptible\nto vibration in the first bending mode. The possibility exists,\nhowever, that in the presence of higher excitation forces, a rela-\ntive lowering of vibratory-stress levels might result from loose\nblade mounting.\n\nINTRODUCTION\n\nA recent modification in turbine-blade fastening has been that\nof increasing the clearance between blade-base serrations and the\n\nRESTRICTED", "timestamp": "2026-07-22T06:55:42.828626+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 32, "total_pages": 47, "image_filename": "19930085919_p32.jpg", "text": "NACA RM No. A9G21\nCONFIDENTIAL\n\nLift coefficient, $C_L$\nAngle of attack, $\\alpha$, deg\nPitching-moment coefficient, $C_m$\n\n$\\delta_f$, deg\n$\\triangle$ 0\n$\\nabla$ 25\n$\\square$ 45\nSplit flap of\ntrapezoidal plan form\n\n$\\delta_f$, deg\n$\\triangle$ 0\n$\\nabla$ 25\n$\\diamond$ 45\n0.28c\nsplit flap\n\nPlain wing +\nshort fuselage\n\nNACA\n\n(a) $C_L$ vs $\\alpha$ and $C_m$.\nFigure 10.- Effect of the split flaps of different plan forms on the lift, drag, and pitching-moment characteristics of the model. R, $4.2 \\times 10^6$.\n\nCONFIDENTIAL\n31", "timestamp": "2026-07-22T06:55:47.458489+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 60, "total_pages": 149, "image_filename": "19930083192_p60.jpg", "text": "56 NACA TN 1976\n\ninvolved in transpacific operations and the lack of weather ships in this period are considered, more frequent accidental encounters with severe weather might be expected than for the transcontinental routes or operations in more populated regions.\n\nOn the basis of the information concerning the flight miles to equal or exceed the acceleration increment corresponding to 37.5K feet per second at the probable speed, the level of route roughness is concluded to be largely independent of the route, airplane, or operator. It is not possible at this time to state whether the maintenance of a constant level of roughness is due to dispatching practices, meteorological forecasting abilities, or both.\n\nFrequency of encountering gusts.- The available data on path ratio given in reference 9 indicate a wide spread in values obtained, and at this time an average value of 0.1 seems to be the best estimate available. Since the path ratio is the proportion of the total miles flown that are spent in rough air and the number of gusts per mile of rough air has been found to be essentially constant at 500/δ, the path ratio can be defined as the actual number of gusts divided by the total number expected if the total flight path were rough. It should be noted that the path ratio defined by gust counts depends on the threshold (in this case, the threshold is 0.3 fps). When used with a gust count of 500/δ gusts per mile of rough air and the gust frequency distributions of figure 8, the recommended average value of path ratio can be used to estimate the probability of an airplane encountering a gust of any specified intensity.\n\nAlthough no information is available as to the variation of the number of gusts or path ratio with altitude, terrain, or weather, the path ratio would probably decrease with altitude.\n\nProbable speed $V_p$.- The determination of the miles to exceed 37.5K feet per second at the probable speed showed the flight miles to be relatively constant, and, therefore, the wide variations in the flight miles for limit load factor might be ascribed to variations in the probable speed. On this basis, the results indicate that the probable speed is important in determining flight miles to limit acceleration increment.\n\nThe data on the probable-speed ratio (table XXI) indicate a scatter in the speed ratios for any given period. The speed ratio for maximum acceleration increased from the prewar period to the postwar period for the same airplanes and routes from an average value of 0.75 to about 0.86. It appears that the speed ratio might be a function of the route, airline policy, and the airplane characteristics, but no conclusion can be drawn as to the significant parameters that determine the speed ratio.", "timestamp": "2026-07-22T06:55:48.008332+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 51, "total_pages": 92, "image_filename": "19930085870_p51.jpg", "text": "52\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O CL\n{□ Cm\n.16\nWedge L.E. {△ CL\n{◇ Cm\n\n.08\nCL\n0\n-.08\n-.16\n-.24\n\n.01\nCm\n0\n-.01\n\n.06\n.04\nCD\n.02\n0\n\n6\n.4\nL/D\n2\n\nElliptical L.E. {O CD\n{□ L/D\nWedge L.E. {△ CD\n{◇ L/D\n\n-8 -6 -4 -2 0 2 4 6 8\nα, deg\n\n(k) Wing II. w = 1.647, R = 570,000.\nFigure 6. - Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:48.467344+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 15, "total_pages": 46, "image_filename": "19930085983_p15.jpg", "text": "NACA RM A9I27 CONFIDENTIAL 13\n\n[Figure: A man in a shirt and tie stands beside a model aircraft wing mounted on a stand. The wing is white, with a swept-back leading edge and a dark cylindrical fuselage section. The background appears to be a wind tunnel or test facility. In the bottom right corner of the image is a NACA logo with the number A-13253.]\n\n(a) Rear view.\n\n[Figure: A top-down (plan) view of the same model aircraft wing. The man is visible behind the model, holding it steady. The wing’s planform shows its swept-back leading edge and tapered shape. The background includes structural elements of the test facility. In the bottom right corner is a NACA logo with the number A-13254.]\n\n(b) Plan view.\n\nFigure 1.— Model of the cambered and twisted wing with the leading edge swept back $63^\\circ$ in combination with a fuselage.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:49.311367+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 16, "total_pages": 25, "image_filename": "19930085979_p16.jpg", "text": "1130\n\nNACA RM E9E12\n\n-Flush-type skin\nthermocouples\n\n-Model center line\n\nAccessory housing\n-Streamline support\nstruts\n\n-Screen\n\n-Orifices\n\n-Plenum\nchamber\n\n-Static-\npressure\ntaps\n\n-Thermocouple\nprobes\n\n-Thermocouple\nrakes\n\n-Front pressure\nrakes\n\n-Static-\npressure\ntubes\n\n-Rear pressure rakes\n\n-Adjustable\ntail cone\n\nMounting strut-\n\nNACA\n\nFigure 2. - Sketch of model showing instrumentation used in investigation.\n\n15", "timestamp": "2026-07-22T06:55:54.641419+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 13, "total_pages": 132, "image_filename": "19930085990_p13.jpg", "text": "NACA RM A9I01 CONFIDENTIAL 11\n\nthe tail above the fuselage had little effect on the lift and drag of the model with the flaps deflected. However, the model with the high tail had more nearly linear pitching-moment characteristics than the model with the tail in the extended wing-chord plane.\n\nWing Wake and Effective Downwash at the Horizontal Tail\n\nThe dynamic pressure at the horizontal tail, the velocity distribution in the wake of the wing-fuselage combination, the effective angles of downwash at the horizontal tail, and the tail efficiency factors are presented in figures 27 through 36.\n\nLocation of the wing wake.- The location of the point of maximum total-pressure loss and the wake boundaries have been determined from measurements of the stagnation pressure behind the wing-fuselage combination at a position corresponding longitudinally to the midchord of the horizontal tail (3.508 wing mean aerodynamic chords behind the quarter point of the wing mean aerodynamic chord) and laterally to the mean aerodynamic chord of the horizontal tail semispan (0.428 wing mean aerodynamic chord from the plane of symmetry). The results of these measurements are presented in figures 27 and 28 where the location of the wake is presented as a function of angle of attack for various Mach numbers and Reynolds numbers. The location of the wake is given with respect to the wing-chord plane at $0^\\circ$ angle of attack. The two alternate positions of the horizontal tail are also identified in these figures so that the location of the tail with respect to the wing-fuselage wake can be readily determined.\n\nThe tail mounted in the extended wing-chord plane was in the wake of the wing at all test angles of attack and at all test Mach numbers. The high tail did not enter the wake until the angle of attack exceeded about $7^\\circ$ at Mach numbers below 0.70. As the Mach number was increased above 0.70, the high tail entered the wake at progressively lower angles of attack. With the wing flaps deflected the high tail was above the wake at all angles of attack. (See fig. 28.)\n\nAt moderate to large angles of attack and at Mach numbers above 0.85, the wing-fuselage wake was characterized by two distinct regions of large total-pressure loss. These are shown in figure 29 which presents the variation of total-pressure loss across the wake at an angle of attack of $6^\\circ$ and a Mach number of 0.85. The secondary peak of total-pressure loss is believed to be associated with separation at the wing leading edge and usually occurred near the angle of attack at which the aerodynamic center of the wing moved forward. Figure 29 also indicates that the presence of the fuselage influenced the magnitude and the location of the total-pressure losses and the location of the wake boundaries.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:56.178466+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 39, "total_pages": 52, "image_filename": "19930085911_p39.jpg", "text": "```markdown\n38\nCONFIDENTIAL\nNACA RM E9F22\n\n1152\n\n[Figure: A graph with two plots. The top plot shows Net acceleration, $a_n$, g's on the y-axis (ranging from -2.0 to 1.0) versus Time after release, $\\tau$, sec on the x-axis (ranging from 0 to 50). The bottom plot shows Free-stream Mach number, $M_0$ on the y-axis (ranging from .2 to 1.4) versus Time after release, $\\tau$, sec on the x-axis. An arrow labeled \"Impact\" points to the x-axis at 50 sec. The NACA logo is in the top right corner of the graph area.]\n\n(a) Resultant flight conditions.\nFigure 9. - Time history of flight data and performance of ram-jet unit 16-A-4.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:55:56.179030+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 18, "total_pages": 30, "image_filename": "19930085975_p18.jpg", "text": "CONFIDENTIAL\n\n$\\Lambda_{c/4}$ Test $\\alpha$\n(deg) (deg)\n—— 36 30,345,8&650\n----326 30,345,8&650\n-- 467 30 & 345\n---467 650\n\nDamping-in-roll correction\nfactor due to wing distortion, K\n\n$C_{l_{\\delta_{corr.}}} = K \\cdot C_{l_{\\delta_{test}}}$\n\n1.2\n1.1\n1.0\n\n4 5 6 7 8 9 1.0\nMach number, M\n\nCONFIDENTIAL\n\nFigure 5.— Correction factor for elastic distortion of test wings under load.\n\nNACA RM L9E10\n16", "timestamp": "2026-07-22T06:55:57.560147+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 10, "total_pages": 40, "image_filename": "19930085997_p10.jpg", "text": "8 CONFIDENTIAL NACA RM A9I29\n\nrecovery and mass flow in the boundary-layer scoops indicate that the boundary layer separated ahead of the boundary-layer scoops as well. In the tests with the maximum rate of flow through the boundary-layer ducts, the pressure in the sting $H_{\\mathrm{S}}$ was maintained at the highest value at which no adverse influence on the main duct system was noted. The relatively low values of $H_{\\mathrm{S}} / H_{0}$ can be attributed to pressure losses in the boundary-layer ducts in addition to the energy dissipated in the boundary layer along the model forebody.\n\nTotal-pressure distribution in duct.- The survey of the total pressure within the duct at position 2 affords a means of estimating the asymmetry in the total pressure of the air flow that would be supplied to a compressor. This factor is important when considering compressor performance or the repeated stresses likely to be imposed upon the compressor blades. In figure 10, the pressure distribution across the height and width of one duct is shown for three values of $m_{1} / m_{0}$ and Mach number. The occurrence of greater asymmetry in the pressure distribution as the free-stream Mach number was increased possibly was caused by the greater intensity of the effects of boundary-layer shock-wave interaction which would result in thickening or separation of the boundary layer. Total pressures near the floor of the duct were consistently low at all Mach numbers and mass-flow ratios indicating that, in the presence of the adverse pressure gradient at the entrance to the scoops, the boundary layer thickens rapidly from the leading edge of the boundary-layer scoop. The maximum variation in total-pressure ratio occurred at a Mach number of 2.01, in which case the difference between the maximum and minimum pressure recovery was approximately 40 percent of the average total-pressure recovery at position 2. This variation is large, but at full scale a smaller variation could be expected because of the much greater Reynolds number and reduced viscous effects.\n\nInteraction between duct systems.- An interaction between the flows in the two main diffusers was manifest in measurements of the total pressure at position 2. With decreasing values of $m_{1} / m_{0}$ from that at which separated flow occurred ahead of the inlets, the pressure recovery at this position in the two ducts diverged about an average value approximately equal to $H_{\\mathrm{S}} / H_{0}$. It was impossible to predict the particular duct passage in which the pressure recovery would diverge above or below the average; however, once the recovery in one side of the induction system had been established above the average, it continued to diverge in this direction with decreasing values of $m_{1} / m_{0}$. This result indicated a possible reversal of flow in one duct at this condition.\n\nTo observe the effects upon the pressure recovery and flow stability of single-duct operation, tests were performed with one duct sealed. The boundary layer was removed ahead of the closed duct in order to reduce the possibility that this flow would influence the flow in the open duct. Results of these tests indicated that separation occurred at about 10-percent-lower values of $m_{1} / m_{0}$ and 2-percent-higher total-pressure ratios $H_{\\mathrm{S}} / H_{0}$ than are shown in figure 7.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:59.750218+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 11, "total_pages": 17, "image_filename": "19930085988_p11.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:56:00.014065+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 20, "total_pages": 55, "image_filename": "19930085966_p20.jpg", "text": "NACA RM L9B17 CONFIDENTIAL 19\n\nUsing equation (1), figure 23 and standard air values of specific heats and the gas constant, thrust coefficients were calculated for assumed values of flight Mach number, combustion-chamber inlet Mach number and total-temperature-rise ratio, and diffuser total-pressure-recovery ratio. Figures 25, 26, 27, and 28 present calculated curves of constant thrust coefficient plotted on ordinates of total-temperature-rise ratio and inlet Mach number for flight Mach numbers of 1.0 and 2.0 with diffuser total-pressure-recovery ratios of 80 and 100 percent. The curves indicate that the highest supersonic flight thrust coefficients are obtained at the test-data boundary which extends from high temperature-rise ratios and moderate inlet Mach numbers to high inlet Mach numbers and moderate temperature-rise ratios. Therefore, in order to indicate the variation with flight Mach number of approximately the maximum thrust coefficient, two conditions along this boundary of inlet Mach number and temperature-rise ratio were chosen. This variation is illustrated in figures 13 and 14 and is discussed under the section \"Results and Discussion.\"\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:56:01.340623+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 26, "total_pages": 34, "image_filename": "19930085913_p26.jpg", "text": "```markdown\nNACA RM L9F24\n25\n\nNo weight, 0% l\n5.55%\n11.11%\n16.66%\n22.20%\n27.77%\n33.33%\n38.90%\n44.40%\n50.00%\n\n55.50%\n61.11%\n66.66%\n72.20%\n77.80%\n83.30%\n88.90%\n94.40%\n100.00%\n\n[Figure: NACA logo]\n\n--- 2nd natural frequency nodal lines\n--- 3rd natural frequency nodal lines\n\n(a) Unswept, untapered wing; $e_w = -1$.\n\nFigure 2.- Progressive change in nodal lines with spanwise weight position.\n```", "timestamp": "2026-07-22T06:56:01.726272+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 99, "total_pages": 114, "image_filename": "19930086061_p99.jpg", "text": "NACA RM L9J07\n95\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n$$ \\frac{y}{b/2} \\text{ , percent} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\circ$ 4.1 | 0.10 |\n| $\\square$ 8.1 | 0.20 |\n\n(a) Angles of attack: 4.1°, 8.1°.\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n$$ \\frac{y}{b/2} \\text{ , percent} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\diamond$ 14.1 | 0.34 |\n| $\\triangle$ 24.1 | 0.55 |\n| $\\triangledown$ 32.1 | 0.64 |\n\n(b) Angles of attack: 14.1°, 24.1°, 32.1°.\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n$$ \\frac{y}{b/2} \\text{ , percent} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\triangle$ 36.1 | 0.67 |\n| $\\diamond$ 44.1 | 0.73 |\n\n(c) Angles of attack: 36.1°, 44.1°.\n\nFigure 44.- Span load distribution of wing 2 at various angles of attack; $\\psi = 35^\\circ$. Flagged symbols represent data taken with left semispan at $\\psi = -35^\\circ$.", "timestamp": "2026-07-22T06:56:02.180900+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 70, "total_pages": 72, "image_filename": "19930085491_p70.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 69\n\n[Figure: (e) WF-63; C_L, 0.]\n\n[Figure: (f) WF-63; C_L, 0.21.]\n\n[Figure: (g) WF-67; C_L, 0.]\n\n[Figure: (h) WF-67; C_L, 0.22.]\n\nFigure 16.— Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:56:11.330480+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 33, "total_pages": 47, "image_filename": "19930085919_p33.jpg", "text": "CONFIDENTIAL\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\n$\\delta_f$, deg\n$\\triangleright$ 0\n$\\triangle$ 25\n$\\square$ 45\n\nSplit flap of\ntriangular plan form\n\n$\\delta_f$, deg\n$\\triangleleft$ 0\n$\\nabla$ 25\n$\\diamondsuit$ 45\n\n0.25c split\nflap\n\nNACA\n\n(b) $C_L$ vs $C_D$.\n\nFigure 10- Concluded.\n\nCONFIDENTIAL\n\nNACA RM NO. A9G21\n\n32", "timestamp": "2026-07-22T06:56:12.914416+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 17, "total_pages": 25, "image_filename": "19930085979_p17.jpg", "text": "16\nNACA RM E9E12\n\nAngle of\nattack\n(deg)\n8\n4\n0\n\n<!-- Image (263, 124, 565, 466) -->\n\n<!-- Image (263, 520, 631, 919) -->\n\nFigure 3. - Effect of angle of attack on inlet-lip pressure distribution.\n\nNACA\n\n1150", "timestamp": "2026-07-22T06:56:14.213802+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 10, "total_pages": 24, "image_filename": "19930085991_p10.jpg", "text": "8\nNACA RM L9I28\n\nthen\n\n$$C_{m_{\\alpha_{fins}}} = -2C_{L_{\\alpha_{fins}}} \\frac{S_F L_T}{S_p L} (\\cos^2\\theta + \\sin^2\\theta)$$\n\nAlso, since\n\n$$\\cos^2\\theta + \\sin^2\\theta = 1$$\n\nthen\n\n$$C_{m_{\\alpha_{fins}}} = -2C_{L_{\\alpha_{fins}}} \\frac{S_F L_T}{S_p L}$$\n\nIn terms of the parameter $S_F$ which was used in the empirical criterion of the present investigation, since $S_F$ is equal to $2S_{F_t} \\cos 45^\\circ$, the stabilizing effectiveness of the fins could have been calculated by\n\n$$C_{m_{\\alpha_{fins}}} = -2C_{L_{\\alpha_{fins}}} \\frac{S_F L_T}{2 \\cos 45^\\circ S_p L}$$\n\n$$= -C_{L_{\\alpha_{fins}}} \\frac{S_F L_T}{\\cos 45^\\circ S_p L}$$\n\nThe values used for $C_{L_{\\alpha_{fins}}}$ were obtained from reference 4 which presents the variation of $C_{L_\\alpha}$ with aspect ratio for low-aspect-ratio wings. Based on information in reference 5, the aspect ratio of each fin was assumed to be effectively 1.5 times its geometric aspect ratio in order to allow for increased fin lift effectiveness caused by end-plate effects of the nose. The calculated stabilizing contribution of the fins was added to the calculated instability of the nose to obtain the resultant $C_{m_\\alpha}$ for the finned nose.", "timestamp": "2026-07-22T06:56:16.778059+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 30, "total_pages": 92, "image_filename": "19930085951_p30.jpg", "text": "```markdown\n28\n\n[CONFIDENTIAL]\n\nEfficiency, $\\eta$\n1.0\n.9\n.8\n.7\n.6\n.5\n.4\n.3\n.2\n.1\n0\n\n$\\beta_{0.75R}$ 20° 25° 30° 35° 40° 45° 50° 55°\n\nHelical-tip Mach number\nAir-stream Mach number\n\nMach number, M\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\nAdvance ratio, J\n0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 2.8 3.0 3.2 3.4 3.6 3.8\n\n(c) Efficiency.\n\nFigure 8.- Concluded. Rotational speed, 1140 rpm.\n\n[CONFIDENTIAL]\n\nNACA RM L9D29\n\n[NACA logo]\n```", "timestamp": "2026-07-22T06:56:16.778848+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 4, "total_pages": 20, "image_filename": "19930085999_p4.jpg", "text": "2\nNACA RM E9I07\n\ncorresponding serrations in the rim of the turbine wheel. Several\nturbojet engines have been fabricated with loosely mounted turbine\nblades. The use of loose blades may possibly increase the vibration\ndamping inherent in inserted-blade design.\n\nThe amplitude of blade-tip movement in the plane of the turbine\nwheel varies among the several existing turbine-blade designs and\namong the blades of any individual turbine wheel because of manufac-\nturing tolerance limits. For example, one manufacturer specifies\nthat the tip amplitude shall be within the limits 0.002 to 0.060 inch\nfor a blade 4 inches long. The factor of blade length must be con-\nsidered when the degree of looseness in blade-base fit is specified\nby amplitude of blade-tip movement.\n\nThe problem of turbine-blade vibration has hindered the devel-\nopment of the turbojet engine (reference 1). A study of some recent\nservice failures in turbine blades, however, has indicated that the\nfailures might be attributed to vibration fatigue (reference 2).\n\nA British investigation has been made of certain blade-vibration\nproblems encountered during development of gas turbines and compres-\nsors (reference 3). In this study of the damping characteristics of\nturbine blades, approximately 70 percent of the damping was attributed\nto the blade-root attachment in the wheel; the remainder was shared\nby aerodynamic and material damping. These values apply only to\nturbine blades retained in serrated dovetails. Another conclusion\ndrawn from the results of the British investigation was that a\nturbine-blade root could be considered as tightly mounted at service\noperation turbine speeds, irrespective of the root tightness when\nstationary. This conclusion is at variance with the concept that\nlooseness of a turbine blade in the wheel rim will effect a con-\nsiderable increase in damping of the turbine-blade vibration.\n\nA program is in progress at the NACA Lewis laboratory to inves-\ntigate the problems associated with vibration in turbine blades. As\na part of this general program, an investigation was made to deter-\nmine the vibration characteristics of loosely mounted turbine blades\nduring actual service operation of a turbojet engine. The measure-\nments of vibration frequency and strain were obtained from high-\ntemperature strain gages mounted on the turbine blades. The tip\namplitudes approximated the average and maximum movement specified\nby the manufacturer for the type of turbine blade employed in the\nexperimental investigation.\n\nA comparison was made between the results from a study of vibra-\ntion characteristics of tightly mounted blades (reference 4) and those", "timestamp": "2026-07-22T06:56:17.218374+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 19, "total_pages": 30, "image_filename": "19930085975_p19.jpg", "text": "NACA RM L9E10\n17\n\nCONFIDENTIAL\n\n<!-- Image (319, 101, 718, 898) -->\n\nFigure 6.- The variation with Mach number of the rolling-moment characteristics of a wing for various aileron deflections at several angles of attack. $\\Lambda_{c/4} = 3.6^\\circ$.", "timestamp": "2026-07-22T06:56:24.258943+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 21, "total_pages": 55, "image_filename": "19930085966_p21.jpg", "text": "20\nCONFIDENTIAL\nNACA RM L9B17\n\nREFERENCES\n\n1. Perchonok, Eugene, Wilcox, Fred A., and Sterbentz, William H.:\n   Preliminary Development and Performance Investigation of a\n   20-Inch Steady-Flow Ram Jet. NACA ACR No. E6D05, 1946.\n\n2. Rubert, Kennedy F.: An Analysis of Jet-Propulsion Systems Making\n   Direct Use of the Working Substance of a Thermodynamic Cycle.\n   NACA ACR No. L5A30a, 1945.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:56:29.148296+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 12, "total_pages": 17, "image_filename": "19930085988_p12.jpg", "text": "NACA RM L9H30\n\nCONFIDENTIAL\n\nPressure tube\n$\\frac{3}{16}$ I.D.\n\nOuter wall\nof nozzle\n\n7.5\n\n0.064\n\n0.042\n\n6.462\n\nFin\n\n30°\n\n3.25 R\n\nSta.\n0\n\n130.0\n\nCONFIDENTIAL\n\nNACA\n\nFigure 3.- Detail of base-pressure-tube installation.\n\n11", "timestamp": "2026-07-22T06:56:29.363419+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 40, "total_pages": 52, "image_filename": "19930085911_p40.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 39\n\n[Figure: Graph with three panels]\n\nTop panel:\nY-axis: Free-stream static temperature, $T_0$, °R\nRange: 400 to 550\n\nMiddle panel:\nY-axis: Free-stream static pressure, $P_0$, lb/sq ft\nRange: 0 to 2500\n\nBottom panel:\nY-axis: Fuel flow, $W_f \\times 3600$, lb/hr; Free-stream total pressure, $P_0$, lb/sq ft\nRange: 0 to 7000\nCurves labeled: $W_f$ and $P_0$\nArrow at τ = 50 sec labeled “Impact”\nNACA logo in bottom right corner of graph area\n\nX-axis (common to all panels): Time after release, $\\tau$, sec\nRange: 0 to 50\n\n(b) Independent test variables.\n\nFigure 9. – Continued. Time history of flight data and performance of ram-jet unit 16-A-4.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:56:30.631541+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 27, "total_pages": 34, "image_filename": "19930085913_p27.jpg", "text": "26\nNACA RM L9F24\n\nNo weight, -11.11%\n11.11%\n16.66%\n22.20%\n27.77%\n33.33%\n38.90%\n44.40%\n50.00%\n55.50%\n\n61.11%\n66.66%\n72.20%\n77.80%\n83.30%\n88.90%\n94.40%\n100.00%\n\n[NACA logo]\n\n—— 2nd natural frequency nodal lines\n- - - - 3rd natural frequency nodal lines\n\n(b) Swept, untapered wing; $\\Lambda = 45^\\circ$, $e_w = -1$.\n\nFigure 2.— Continued.", "timestamp": "2026-07-22T06:56:30.861597+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 11, "total_pages": 40, "image_filename": "19930085997_p11.jpg", "text": "NACA RM A9I29 CONFIDENTIAL 9\n\nEstimate of the energy expended in boundary-layer removal.- The equivalent pressure recovery obtained by subtracting the energy required for boundary-layer removal from the measured recovery does not provide a completely adequate criterion for the worth of this system. In considering a specific application, the advantages of arrangement that may result from use of side scoops with boundary-layer control, possible uses of the boundary-layer air such as in engine cooling, and the results of a detailed analysis of the effects of the inlets upon the external drag would also have to be considered. In order to obtain an indication of the effective recovery, however, the energy expended in the boundary-layer scoops was subtracted from the energy recovered in the main scoops in order to arrive at an equivalent value of the pressure recovery at position 2, $(\\overline{H}_2/H_0)_e$.\n\nThe tables of reference 5 were used in these calculations and the experimental results were applied to full-scale flight at an altitude of 35,000 feet. It was assumed that the energy required to remove the boundary layer was equal to that necessary to compress isentropically the mass flow in the boundary-layer scoop $m_4$ from the total pressure after diffusion $H_5$ to a total pressure $H_4$ corresponding to a 50-percent decrease in the free-stream kinetic energy. The latter assumes a turbulent boundary layer at the entrance to the boundary-layer scoops and that the energy contained in the boundary layer is equal to that of the free stream depleted of 50 percent of its kinetic energy. The results of calculations making use of the preceding assumptions are shown in figure 11. Values of $(\\overline{H}_2/H_0)_e$ were calculated within the range of mass-flow ratio at which the flow into the scoops of configuration D was steady. A comparison of these results with the measured pressure recovery at position 2 indicates that the energy required to remove the boundary layer was equivalent to a loss in total pressure $\\overline{H}_2/H_0$ of approximately 0.08.\n\nThe extension of the data to full scale, however, requires some consideration of the effect of the model scale, which was taken as 1.4 percent. In estimating the influence of the model scale, the following assumptions were made:\n\n1. The boundary layer at the entrance to the boundary-layer scoops was turbulent.\n\n2. The boundary-layer thickness on the model and at full scale varies as\n\n$$\n\\frac{(\\delta/x)_{\\text{model}}}{(\\delta/x)_{\\text{full scale}}} = \\left[ \\frac{(Rx)_{\\text{full scale}}}{(Rx)_{\\text{model}}} \\right]^{1/5}\n$$\n\nwhere $x$ is a characteristic length. (See reference 6.)\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:56:34.036960+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 14, "total_pages": 132, "image_filename": "19930085990_p14.jpg", "text": "12 CONFIDENTIAL NACA RM A9I01\n\nDynamic-pressure ratio and Mach number at the tail.- To determine the ratio of the dynamic pressure at the tail to the free-stream dynamic pressure, measurements were made of the stagnation and static pressures in the region of the horizontal tail. The results of these measurements are presented in figure 30 for various free-stream Mach numbers as a function of angle of attack. The dynamic-pressure ratio at the centroid position of the horizontal tail in the extended wing-chord plane for $0^\\circ$ angle of attack varied from 0.945 at a free-stream Mach number of 0.20 to 0.865 at a free-stream Mach number of 0.95. Due to the symmetry of the model about the wing-chord plane, the dynamic-pressure ratio at the tail mounted in the extended wing-chord plane increased with increasing or decreasing angle of attack, attaining a value of approximately 0.98 at all Mach numbers at angles of attack of $\\pm 6^\\circ$.\n\nAt a Mach number of 0.95, the dynamic-pressure ratio at the centroid position of the high horizontal tail was unity at angles of attack less than $2.5^\\circ$ and less than unity at larger angles of attack. (See fig. 30(b).) As free-stream Mach number decreased, the minimum angle of attack for which the dynamic pressure remained at the free-stream value increased to $7^\\circ$ for Mach numbers less than 0.70.\n\nWith the wing flaps deflected, the dynamic-pressure ratio at the high tail position was unity, and at the position of the tail in the extended wing-chord plane it varied from approximately 0.99 at $0^\\circ$ angle of attack to approximately 0.84 at $10^\\circ$ angle of attack. The effect of increasing the Reynolds number from 2,000,000 to 10,000,000 was to increase the dynamic-pressure ratio approximately 5.5 percent at an angle of attack of $10^\\circ$ with less effect as the angle of attack was reduced.\n\nThe Mach numbers at the tail have been computed from the wake-survey data and are presented as functions of angle of attack for various free-stream Mach numbers in figure 32.\n\nEffective angles of downwash at the tail.- The effective angles of downwash at the horizontal tail have been computed from the moment data and are presented as average values over the stabilizer angle range in figures 33 and 34. The expression used for calculation of the effective angle of downwash is as follows:\n\n$$\n\\epsilon = \\alpha + i_t - \\frac{(\\Delta C_{m_t})_\\alpha}{(\\partial C_m / \\partial i_t)_\\alpha}\n$$\n\nwhere $(\\Delta C_{m_t})_\\alpha$ is the increment in pitching-moment coefficient due to the addition of the tail for a constant angle of attack and $(\\partial C_m / \\partial i_t)_\\alpha$ is the stabilizer effectiveness at a constant angle of attack. This expression does not permit the downwash due to the wing to be separated from the downwash due to other components of the model, and thus the stream angle at the horizontal tail due to convergence of the rear end\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:56:34.255508+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 7, "total_pages": 32, "image_filename": "19930085992_p7.jpg", "text": "NACA RM L9E17\n\nrange. A cross-sectional view of the wing is given in figure 2 and the wing properties were as follows:\n\nChord, inches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T06:56:35.546311+00:00"}