PDF/ocr_dataset_1073.jsonl

{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 10, "total_pages": 22, "image_filename": "19930085922_p10.jpg", "text": "2C\nNACA RM No. 19C23\n\nTunnel ceiling\nBalance strut\nStrut fairing\nSting fairing\nTunnel center line\nSting support\nShaft locking screws\nTest wing\nRoller bearing\nAngle of attack\nchanging block\nBall thrust\nbearing\nTunnel floor\nNACA\n0 5 10\nScale, inches\n\nFigure 2.- Schematic drawing of the free-rolling sting mounted in the Langley high-speed 7- by 10-foot\ntunnel test section.\n9", "timestamp": "2026-07-22T06:40:37.611903+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 66, "total_pages": 78, "image_filename": "19930082483_p66.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:40:37.804457+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 76, "total_pages": 98, "image_filename": "19930086073_p76.jpg", "text": "74\nNACA RM A9H04\n\nLift coefficient, $C_L$\n\nPitching - moment coefficient, $C_m$\n\n$\\circ$ $\\square$\n$\\delta_{a_L} = +10.8, \\& \\delta_{a_R} = -10.8$\nAileron deflection, $\\delta_a$, deg\n\n(c) $C_L$ vs $C_m$.\n\nFigure 16. — Continued.", "timestamp": "2026-07-22T06:40:39.358575+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 64, "total_pages": 65, "image_filename": "19930082546_p64.jpg", "text": "NACA TN No. 1870\n63\n\n<!-- Image (164, 156, 870, 732) -->\n\nFundamental frequency of panel excitation, cps\n(b) Circular steel wall with two-blade propeller.\nFigure 22.- Continued.", "timestamp": "2026-07-22T06:40:40.134241+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 27, "total_pages": 92, "image_filename": "19930085870_p27.jpg", "text": "NACA RM No. L9D07\n27\n\nCONFIDENTIAL\n\n[Figure: Diagram showing dimensions of 8-percent-thick triangular-wing models. Includes planform view with labels A-A, E, W, MAC, b, 2.25, .32, .165, 1.75, 0.125; side profile with $c_r$, $xc$, $yc$, .45, 0.30, .15; and cross-sections labeled E (elliptical) and W (wedge). NACA logo present.]\n\nFigure 2.— Dimensions of 8-percent-thick triangular-wing models. (Sting dimensions identical for all wings. E, elliptical leading edge; W, wedge leading edge.) CONFIDENTIAL", "timestamp": "2026-07-22T06:40:42.695690+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 22, "total_pages": 29, "image_filename": "19930085899_p22.jpg", "text": "NACA RM No. L9J21\n\n| Angle of attack, $\\alpha$, deg | Lift coefficient, $C_L$ | Pitching-moment coefficient, $C_m$ | Drag coefficient, $C_D$ |\n| :--- | :--- | :--- | :--- |\n| 0 | -0.2 | 0 | 0 |\n| 0 | 0 | 0 | 0 |\n| 0 | 0.2 | 0 | 0 |\n| 12 | 0.4 | 0 | 0 |\n| 8 | 0.6 | 0.2 | 0.08 |\n| 4 | 0.8 | 0.1 | 0.04 |\n| 0 | | 0 | 0 |\n| -4 | | -0.1 | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | M | |\n| | | 1.10 | |\n| | | 1.00 | |\n| | | 0.90 | |\n| | | 0.80 | |\n| | | 0.60 | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | M | |\n| | | 1.10 | |\n| | | 1.00 | |\n| | | 0.90 | |\n| | | 0.80 | |\n| | | 0.60 | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| 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| |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n\n| | | | |\n| :--- | :--- | :--- | :--- |\n| | | | |\n| | | | |\n| | | | |\n| | | | |", "timestamp": "2026-07-22T06:40:45.084491+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 39, "total_pages": 60, "image_filename": "19930085862_p39.jpg", "text": "```markdown\nNACA RM No. L9A07\n37\n\n<!-- Image (159, 110, 832, 874) -->\n\n(b) $C_{h_a}$ and $P_R$ against $\\alpha$.\nFigure 10.— Continued.\n```", "timestamp": "2026-07-22T06:40:46.894230+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 9, "total_pages": 47, "image_filename": "19930085919_p9.jpg", "text": "8 CONFIDENTIAL NACA RM No. A9C21\n\nattack and flap deflections investigated (an increment of at least 0.4 up to an angle of attack of 24°) and increased the lift coefficient attained before the occurrence of longitudinal instability from about 0.5 to 0.8. As the hinge line of the split flap was moved forward from 100 to 40 percent of the wing chord, the flap effectiveness decreased rapidly. Although the split flap with its hinge line at the wing trailing edge produced the largest lift increases, this flap also produced the largest changes in longitudinal balance.\n\nThe split flaps with their hinge lines normal to the air stream provided less negative pitching moments and smaller lift increments than did the split flap with its hinge line at the trailing edge of the wing (fig. 9(a)). With the split flap deflected 60° in the forward position with its hinge normal to the air stream the lift coefficient attained before the occurrence of longitudinal instability was increased from about 0.5 to 0.6. Surface tufts indicated that the split flaps with their hinge lines normal to the air stream caused flow separation to occur initially near the midsemispan of the wing at an angle of attack of 0°. At angles of attack greater than 12°, these split flaps caused a larger portion of the wing to stall, which is probably responsible for the decreased lift-curve slope and the decreased maximum lift coefficient (fig. 9 (a)).\n\nIncrease of the deflection of the split flaps from 45° to 75° caused relatively small changes in the lift and pitching-moment characteristics (fig. 9(a)). Deflecting some of the split flaps more than 45°, for example, the flap hinged at 100 percent of the wing chord, decreased the maximum lift coefficient. Only the split flap at the trailing edge of the wing greatly reduced the drag of the model at high lift coefficients (fig. 9(b)).\n\nThe data for the model with the split flap of triangular plan form and with the split flap of constant-percent chord (both hinged along the wing trailing edge) are presented in figure 10. At high angles of attack the split flap of triangular plan form produced slightly larger increments of lift coefficient than the split flap of constant-percent chord of the same area. Longitudinal instability occurred at approximately the same lift coefficient with the same deflection of either flap, but the split flap of triangular plan form deflected 45° produced slightly less negative pitching moments at small angles of attack. With either flap at 0° angle in the extended position, the lift-curve slope was increased from 0.046 to 0.052 per degree and the aerodynamic center was shifted rearward 1.5 percent of the mean aerodynamic chord at small lift coefficients.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:40:49.799774+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 45, "total_pages": 47, "image_filename": "19930083221_p45.jpg", "text": "NACA TN No. 1824\n43\n\n| | |\n| :--- | :--- |\n| $v_r$ | radial component of perturbation velocity |\n| $x,y,z$ | Cartesian coordinates |\n| $\\alpha$ | angle of attack in radians |\n| $\\beta$ | $\\sqrt{|1-M_0^2|}$ |\n| $\\gamma$ | ratio of specific heats, for air $\\gamma = 1.4$ |\n| $\\Delta\\Phi_0, \\Delta u_0, \\Delta w_0$ | discontinuity in component in $z = 0$ plane |\n| $\\theta$ | semivertex angle of swept back wing |\n| $\\nu$ | $\\frac{\\omega c_0}{2}$ |\n| $\\rho_0$ | free-stream density |\n| $\\Phi$ | total velocity potential |\n| $\\varphi$ | perturbation velocity potential |\n| $\\omega'$ | impressed frequency (reference to true time) |\n| $\\omega$ | $\\frac{\\omega'}{a_0}$ |\n\nREFERENCES\n\n1. von Kármán, Th.: The Similarity Law of Transonic Flow. Journal of Math. and Physics. vol. XXVI, no. 3, Oct. 1947, pp. 182-190.\n2. Stewart, H. J.: The Lift of a Delta Wing at Supersonic Speeds. Quart. App. Math., vol. IV, no. 3, Oct. 1946, pp. 246-254.\n3. Oswatitsch, Klaus, and Wieghardt, K.: Theoretical Analysis of Stationary Potential Flows and Boundary Layers at High Speed. NACA TM No. 1189, 1948.\n4. Sauer, R.: General Characteristics of the Flow Through Nozzles at Near Critical Speeds. NACA TM No. 1147, 1947.", "timestamp": "2026-07-22T06:40:54.786682+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 7, "total_pages": 22, "image_filename": "19930085928_p7.jpg", "text": "```markdown\n6\nCONFIDENTIAL\nNACA RM No. A9A31\n\nDucts Without Slots\n\nA comparison of the curves of figure 3 shows that the changes\nin the internal shape of the ducts without slots produced a large\nincrease in the maximum total-pressure ratios $(H_3/H_0)_{max}$ beyond\nthose attained with the unslotted inlet of reference 1. The increase\nwas 3 percent at a Mach number of 1.36 and 14 percent at a Mach\nnumber of 2.01. The degree of constriction had only a small effect\non this improvement. Schlieren photographs, such as those of figure 4,\nof the flow about models A and B indicate that a normal shock wave\nexisted upstream of the duct entrances through the Mach number range\nof the tests when the total-pressure ratio was at the maximum value.\nTherefore, the flow into the scoops was subsonic for the conditions\nof figure 3 and underwent no supersonic compression through the\ncontracting inlet passages. Apparently, the increase in the pressure\nrecovery above that of the model tested in reference 1 was the result\nof an improvement of the flow in the subsonic diffusor.\n\nThe variation of total-pressure ratio $H_3/H_0$ with mass-flow\nratio$^1$ $m_1/m_0$ for models A and B is shown in figure 5. Although the\nmaximum total-pressure ratios are nearly the same, the range of flow\nratios over which a high recovery can be maintained and also the\nmaximum flow rate are greater with model A. For nearly all of the\ntest conditions represented on these curves, a normal shock wave\nexisted in the stream ahead of the inlets. Only for the greatest\nvalues of mass-flow ratio at a test Mach number of 2.01 did this\nshock wave retreat into the ducts, and then only with model A. The\nfact that this anticipated event did not occur until a Mach number\ngreater than that calculated was reached is probably caused by the\npresence of the forebody boundary layer. The relatively large\ndisplacement thickness of this boundary layer increased the effective\ncontraction of the inlet passages and thereby delayed the entrance of\nthe normal shock wave beyond the Mach number for which it would be\nswallowed in unidimensional, inviscid flow.\n\nWhen the flow through the inlet was subsonic, the stream was\naccelerated in the constricted passage, and at large mass-flow ratios\nsonic velocity probably existed in the throat. Downstream of the\nthroat the flow expanded and became supersonic again until a second\nnormal shock wave or a complex pattern of shock waves reduced it to\nsubsonic velocity and it was finally diffused. Since the throat\n\n$^1$Mass-flow ratio is defined as the mass of fluid entering the inlet\ndivided by that which would flow through a tube of the same area\nin the free stream.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:40:58.391821+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 42, "total_pages": 149, "image_filename": "19930083192_p42.jpg", "text": "38\nNACA TN 1976\n\nare used. The calculations were performed by a step-by-step procedure that took into account the possible lack of coincidence in the gust peaks. Other calculations indicated that for gusts with a distance of 10 chords, the maximum lag in the acceleration peak would be about 2 chords when the wing had an infinite mass parameter.\n\nExperimental studies.- Most of the experimental data available are from general tests in the Langley gust tunnel with upward-acting gusts perpendicular to the flight path. Most of the research can be classed as general and includes the material presented in reference 19. The results of tests of specific models that are applicable to general studies are included.\n\nA general study was made of the effect of the stability of airplanes on the gust load factor; the stability characteristics were varied over a wide range within which conventional airplanes would be expected to fall. The tests were performed on an arbitrary \"stability\" airplane model for correlation with the extended analysis previously described. The characteristics of this model are given in table XI and the total acceleration increments obtained from the tests are shown in figure 35. The tests were made for three gust-gradient distances and with the center-of-gravity positions at 17.5, 25, and 30 percent of the mean geometric chord. The combinations of tail area and tail length used on the test model are listed in figure 35.\n\nThe unconventional airplane models listed in table XI include a flying wing and a canard airplane. Tests of the flying wing were made with the center of gravity at 20 percent of the mean geometric chord. The acceleration increment, forward speed, gust velocity, and pitch of the model were measured during each test flight. The tests consisted of a series of five or more flights of the airplane at a forward speed of 40 miles per hour through vertical gusts with gradient distances of 0, 8, and 16 chords. The test results have been included in table XIII and are shown in figure 38(a). The test conditions for the canard airplane have previously been reported in reference 15. The pertinent characteristics of the model have been included in table XI and the results of the experiments are given in table XIV and in figure 38(b).\n\nA few flight tests have been made to obtain data on the behavior of an airplane subjected to unsymmetrical gusts and on the gust loads on the vertical and horizontal tail surfaces. The results for the unsymmetrical gusts have been discussed previously in the section entitled \"The Structure of Atmospheric Gusts\" and indicate reasonable agreement between calculation and experiment. The results of investigations of gust tail loads on the C-2H and XB-15 airplanes are given in table XV as ratios of the effective gust velocity on a tail surface to that determined for the wing. The XB-15 airplane is shown in figure 6 and described in the section on gust structure and the general characteristics", "timestamp": "2026-07-22T06:41:04.850927+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 11, "total_pages": 34, "image_filename": "19930085913_p11.jpg", "text": "10\nNACA RM L9F24\n\nin reference 1. The ratios of the weights and mass moment of\ninertias $\\left(\\frac{W_w}{W} \\text{ and } \\frac{I_w}{I}\\right)$ used in the reference papers were approximately\nthe same as those used in this paper. In reference 3 and this paper a\ncomplete spanwise survey of the unswept wing could not be experimentally\nobtained because the flutter velocity exceeded the divergence velocity\nover a large range of weight positions. However, a complete spanwise\nsurvey of the wing in reference 3 was analytically obtained in refer-\nence 1. It is of interest to mention that the trend with spanwise\nposition of the analytical results obtained in reference 1 for an\nunswept wing was similar to that reported in this paper for the swept\nwings, each being weighted at the leading edge.\n\nTwo flutter speeds were obtained from the unswept, unweighted\nwing. In the neighborhood of the lower flutter speed the flutter was\nnot of a destructive nature and the lower speed range could be exceeded\nand the higher flutter speed obtained. The flutter occurring at the\nlower speed appeared to involve a significant amount of wing second\nbending, and at the higher speed the model appeared to vibrate very\nlittle in bending and its motion was predominantly torsional. The\nunweighted, unswept wing was the only model for which two flutter\nvelocities were recorded. All other flutter velocities reported herein\nwere the lowest values obtained regardless of the violence of the\nflutter.\n\nIn table I several of the recorded flutter frequencies are marked.\nThe flutter in these cases was of an unusual nature. (See fig. 1.)\nConsider run 71, for example, in which case two distinct frequencies\nwere obtained simultaneously at flutter. This case was unexpected\nsince flutter usually involves only one frequency or, in some cases,\noccurs with a burst of one frequency, then a burst of a different\nfrequency. Run 71 (weight at 41.67 percent l) was not an isolated\ncase unrelated to those in which the weight position was nearby on\neither side. As the weight was moved spanwise (runs 67 to 74) the\nmodel experienced a change in flutter mode. The amplitude of the tip\ngage traces diminished as the weight was moved toward the tip of the\nwing (runs 67 to 70). In run 71 the tip traces came in at a higher\nfrequency, and a high frequency persisted on the tip traces in runs 71\nto 74. The root traces, on the other hand, had a relatively constant\namplitude and frequency in runs 67 to 71 and then were inactive in\nruns 72 to 74. On the basis of the records presented in figure 1,\nrun 71 can be considered to be part of a flutter-mode change.\n\nA possible explanation of such an unusual type of flutter is that\nthe flexibility of the thin plate-like structure and the associated\nnodal-line patterns may be involved. Furthermore, these models had low\nstructural damping.", "timestamp": "2026-07-22T06:41:05.175615+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 11, "total_pages": 22, "image_filename": "19930085922_p11.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:41:12.725263+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 17, "total_pages": 52, "image_filename": "19930085911_p17.jpg", "text": "16 CONFIDENTIAL NACA RM E9F22\n\nand the total pressure is assumed equal to the free-stream total pressure minus external shock losses.\n\n$$\nM = \\sqrt{\\frac{2}{\\gamma - 1} \\left[ \\left( \\frac{p}{P} \\right)^{\\frac{\\gamma - 1}{\\gamma}} - 1 \\right]}\n$$\n\nThe ram-jet air flow is determined from the data available at station 3 by assuming that the total pressure at station 3 equals the measured total pressure at station 4 and that the total temperature equals the free-stream total temperature. The method involves determining\n\n$$\np_3 = P_3 - (P_3 - p_3)\n$$\n\nthe Mach number at station 3 from equation (5), and\n\n$$\nW_a = p_3 A_3 M_3 \\left( \\frac{P_3}{p_3} \\right)^{\\frac{\\gamma - 1}{2\\gamma}} \\sqrt{\\frac{\\gamma g}{R T_3}}\n$$\n\nFor the conditions at the diffuser outlet, the total pressure $P_4$ is known and Mach number $M_4$ is determined from $M_3$ by calculating $A_{cr,3}/A_3$ from the general equation\n\n$$\n\\frac{A_{cr}}{A} = M \\left( \\frac{\\frac{\\gamma + 1}{2}}{1 + \\frac{\\gamma - 1}{2} M^2} \\right)^{\\frac{\\gamma + 1}{2(\\gamma - 1)}}\n$$\n\nand then, because $P_4 = P_3$ and $T_4 = T_3$\n\n$$\n\\frac{A_{cr,4}}{A_4} = \\left( \\frac{A_{cr,3}}{A_3} \\right) \\left( \\frac{A_3}{A_4} \\right)\n$$\n\nFrom equation (8), $M_4$ can be determined for the value of $A_{cr,4}/A_4$. It is then possible to calculate $p_4$ from equation (3), $t_4$ from equation (2), and $V_4$ from equation (1).\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:41:14.735930+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 67, "total_pages": 78, "image_filename": "19930082483_p67.jpg", "text": "NACA TN No. 1807\n65\n\n1032\n\nLocation of turbine rotor\nLow-pressure exhaust duct\nExternal hot-gas producer\nExplosion shield\nExpansion bellows\nGas-flow\nEddy-current-type\nabsorption dynamometer\n\nNACA\nC.21411\n5.11.48\n\nFigure 6. - Over-all test setup used for partial-admission investigation.", "timestamp": "2026-07-22T06:41:15.188737+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 50, "total_pages": 62, "image_filename": "19930082918_p50.jpg", "text": "NACA TN 1940\n61\n\n<!-- Image (174, 110, 812, 934) -->\n\nFigure 10.- Effect of aging on (111) line intensity of low-carbon N-155 alloy solution-treated 10 hours at 2200° F and water-quenched. I, line peak intensity, indicated time; $I_0$, line peak intensity, unaged.", "timestamp": "2026-07-22T06:41:20.826208+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 65, "total_pages": 65, "image_filename": "19930082546_p65.jpg", "text": "64\nNACA TN No. 1870\n\n<!-- Image (90, 168, 847, 852) -->\n\nFundamental frequency of panel excitation, cps\n(c) Flat vertical wooden wall with four-blade propeller.\nFigure 22.- Concluded.", "timestamp": "2026-07-22T06:41:22.210429+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 75, "total_pages": 114, "image_filename": "19930086061_p75.jpg", "text": "NACA RM L9J07\n71\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n\n-3\n-2\nP -1\n0\n1\nRight semispan\n\nUpper\nLower\n\n20°\n\n(c) $\\psi = 20^\\circ$\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n\n-3\n-2\nP -1\n0\n1\nRight semispan\n\nUpper\nLower\n\n35°\n\n(d) $\\psi = 35^\\circ$\n\nNACA\n\nFigure 24.- Concluded.", "timestamp": "2026-07-22T06:41:24.469441+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 10, "total_pages": 47, "image_filename": "19930085919_p10.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 9\n\nThe drag characteristics of the model with the triangular flap were similar to those of the model with the 25-percent-chord split flap for the same flap deflections (fig. 10(b)).\n\nElevons\n\nThe characteristics of the model with various deflections of the constant-percent-chord elevon and the constant-chord elevon are presented in figure 11. The pitching moments with the constant-chord elevon undeflected were slightly different from the pitching moments with the constant-percent-chord elevon undeflected. Similar discrepancies may be found in other figures of this report. These discrepancies are believed to have been caused by small differences in the contour or in the $0^\\circ$ settings of the various controls. At small angles of attack, the rates of change of pitching- and rolling-moment coefficients with elevon deflection for the two elevons were approximately in proportion to their area moments about the pitch or roll axes.³ The rate of change of lift, pitching-moment, and rolling-moment coefficient with elevon deflection remained nearly constant up to an angle of attack of $9^\\circ$, decreased between angles of attack of $9^\\circ$ and $17^\\circ$, but increased at higher angles of attack for negative deflection of the elevons. In the low lift range, the rate of change of pitching-moment and rolling-moment coefficients with elevon deflection decreased as the negative deflection of the elevon exceeded $30^\\circ$.\n\nThe characteristics of the model with the constant-chord elevon and the 0.25-chord split flap deflected $45^\\circ$ at the wing trailing edge are presented in figure 12. This was the elevon split-flap combination tested with the leading-edge flaps which will be discussed in the succeeding sections of this report. The rate of change of lift and rolling-moment coefficients with elevon deflection remained nearly constant to an angle of attack of $5^\\circ$, but decreased at higher angles of attack. Therefore, with the split flap the effectiveness of the constant-chord elevon began to decrease at a smaller angle of attack than without the split flap (figs. 11 and 12).\n\nAs mentioned hereinbefore the split flap hinged at the wing trailing edge produced large changes in balance; therefore, the\n\n---\n\n³The moment of the area of the constant-chord elevon about either the pitch or the roll axis was approximately 1.5 times that of the constant-percent-chord elevon.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:41:25.201671+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 41, "total_pages": 46, "image_filename": "19930085548_p41.jpg", "text": "40\nNACA RM No. E8L30\n\n<!-- Image (156, 119, 825, 780) -->\n\nFigure 18. - Comparison of experimental and calculated indicated thermal efficiencies on basis of compression ratio. Inlet-manifold temperature, 400° F; inlet-manifold pressure, 100 pounds per square inch absolute; calculated thermal efficiency $\\eta_t = K \\left(1 - \\frac{1}{r^{0.4}}\\right)$.", "timestamp": "2026-07-22T06:41:25.201828+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 77, "total_pages": 98, "image_filename": "19930086073_p77.jpg", "text": "```markdown\nNACA RM A59E04\n\nLift coefficient, $C_L$\n\nRolling-moment coefficient, $C_l$\n\nYawing-moment coefficient, $C_n$\n\nSide-force coefficient, $C_Y$\n\n$\\delta_a$, deg\nO 0\n$\\square$ $\\delta_{a_L}=+10.8$, $\\delta_{a_R}=-10.8$\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 16.—Concluded.\n\n75\n```", "timestamp": "2026-07-22T06:41:25.610978+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 28, "total_pages": 92, "image_filename": "19930085870_p28.jpg", "text": "28\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n$\\frac{3}{8}\"$ bevel on all edges\n\n$\\epsilon$\n\nMAC.\n\nC\n\nNACA\n\n$t$\n\nFigure 3.- Dimensions of flat-plate triangular-wing models. (Sting supports identical with thick-wing installation.)\nCONFIDENTIAL", "timestamp": "2026-07-22T06:41:27.280530+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 40, "total_pages": 60, "image_filename": "19930085862_p40.jpg", "text": "38\nNACA RM No. L9A07\n\n<!-- Image (144, 109, 736, 874) -->\n\n(c) $C_L$ and $C_m$ against $\\alpha$.\nFigure 10.- Concluded.", "timestamp": "2026-07-22T06:41:28.767839+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 45, "total_pages": 72, "image_filename": "19930085491_p45.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:41:33.480563+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 8, "total_pages": 22, "image_filename": "19930085928_p8.jpg", "text": "NACA RM No. A9A31 CONFIDENTIAL 7\n\narea of model B was less than that of model A and since the throats were choked at high flow ratios, less air flowed through the inlet of model B. As the outlet area of the passages through the models were reduced from the maximum values, the mass-flow ratio could not change because the velocity of the flow in the inlet throats was sonic. However, the total-pressure ratio increased because the shock losses in the subsonic diffuser moved toward the throat and occurred at a lower Mach number. The portion of the curves of figure 5 that indicates that the mass-flow ratio decreased while there was little change in the pressure recovery suggests that the flow through the ducts for this condition was entirely subsonic and there was an increase in the spillage around the lips. This fact is indicated by the schlieren photographs of figure 4 which show that the normal shock wave moved farther ahead of the inlet as the flow ratio was reduced. When the mass-flow ratio was reduced sufficiently, the boundary layer separated and the flow through the ducts became unsteady.\n\nDucts With Slots\n\nThe purpose of cutting slots in the duct walls of the inlet of models A and B was to permit the boundary layer of the flow over the forebody to escape from the ducts and thereby not only to remove low-energy air from the internal stream, but also to permit a sufficient mass of air to escape so that the normal shock wave which forms upstream of a constricted passage at low supersonic Mach numbers could enter the inlet. The expected result would be an increase in the total-pressure recovery attainable with the models. The maximum total-pressure ratios shown in figure 6 for models A and B having inlets with slots are greater than any attained with other similar scoop configurations. There is little difference, however, in the recovery attained by the two models. Of the slot sizes tested, the slots that were found to be best in reference 1 also produced the greatest recovery with models A and B; the effects of changes were small. The maximum total-pressure ratios attained were greater than those across a normal shock wave throughout the Mach number range of the tests and were within 2 percent of those attained with the nose inlets of reference 6 at Mach numbers of 1.7 and less.\n\nFigure 7 shows the variation of total-pressure ratio with mass-flow ratio for models A and B at several Mach numbers. These curves show that, although the maximum recovery with both models is nearly the same, greater mass-flow ratios can be attained with model A and a higher recovery of pressure is maintained at large flow ratios and over a wider range. Operation at the highest possible flow ratio is,\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:41:34.939136+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 68, "total_pages": 78, "image_filename": "19930082483_p68.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:41:41.688471+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 12, "total_pages": 22, "image_filename": "19930085922_p12.jpg", "text": "NACA RM No. L9C23\n\n[Figure: Photograph of a test wing model mounted on a sting in a wind tunnel test section. The model has a pointed nose, swept wings, and is attached to a vertical support structure. In the background, there are two small inset images showing close-ups of instrumentation or components. A label in the lower right corner of the image reads “NACA L-54508”.]\n\nFigure 3.- Photograph of the test wing, vertical fins off, mounted on the free-rolling sting in the Langley high-speed 7- by 10-foot tunnel test section.\n\n11", "timestamp": "2026-07-22T06:41:42.302092+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 51, "total_pages": 62, "image_filename": "19930082918_p51.jpg", "text": "62\nNACA TN 1940\n\n<!-- Image (103, 109, 785, 997) -->\n\nFigure 11.- Effect of aging on lattice parameter of low-carbon N-155 alloy solution-treated 10 hours at 2200° F and water-quenched.", "timestamp": "2026-07-22T06:41:44.411266+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 26, "total_pages": 26, "image_filename": "19930085890_p26.jpg", "text": "NACA - Langley Field, Va.\n\nBursting disk\n\nThermocouple\n\nInsulated heating wire\n\nInsulation\n\nStainless-steel bomb\n\nPressure-gage connection\n\nValve\n\n(a) Temperature sensitivity.\n\nNumber 6 electric detonating cap\n\nSmall separating tube\n\nBrass container\n\nValve\n\n(b) Detonation sensitivity.\n\nFigure 9. - Sensitivity apparatus.\n\nNACA RM No. E9C11\n\n25", "timestamp": "2026-07-22T06:41:44.633194+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 41, "total_pages": 60, "image_filename": "19930085862_p41.jpg", "text": "NACA RM No. L9A07\n39\n\n(a) Plain wing\n(b) Split flaps\n(c) Drooped-nose and split flaps and fences.\n\nFigure 11.- Variation of rolling-moment coefficient with aileron deflection for various flap configurations.", "timestamp": "2026-07-22T06:41:51.519070+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 12, "total_pages": 34, "image_filename": "19930085913_p12.jpg", "text": "```markdown\nNACA RM L9F24\n11\n\nTable I contains the phase-angle relationships between the torsional and bending stresses of the second and third natural frequencies and the flutter modes of vibration of each of the models tested. The phase angles pertaining to the flutter modes were read from records, portions of which are shown in figure 1. The phase angles pertaining to the natural modes were read from records not appearing in this paper. All the phase angles were obtained using the system illustrated in the section on apparatus. It should be kept in mind that these phase angles relate the wing stresses at the gage stations and not the deflections. In order to obtain a deflection curve, the spanwise stress or moment distribution must be known. Since the wings tested carried only two sets of gages, this spanwise distribution was not obtained. Thus, if the tip and root bending stresses are out of phase, it merely indicates that there is at least one inflection point in the mode shape but not necessarily a nodal point. On the records in figure 1, the various gage traces are marked with their appropriate attenuations. Since the amplitude of the gage trace is inversely proportional to its attenuation for a given stress and the amplitude is directly proportional to the stress, the bending moment at the root may be approximately related to the bending moment at the tip and the torque at the root to the torque at the tip. The relative bending moments and torques and the phase angles between bending and torsion at two stations on the wing can give no direct information as to the flutter mode but might be used as a check on results obtained analytically.\n\nThe large amount of coupling present when the weight was located on the leading edge of the wings is clearly illustrated by the sketches shown in figure 2. Note that the third natural frequency was of a torsional nature when the weight was near the root but changed to one of a second-bending nature as the weight neared the tip of the wing, and, conversely for the second natural frequency. Figure 2 also serves to illustrate that sweepback alone induced coupling.\n\nThe sketches of the models indicate that the roots were parallel to the air stream. An investigation of the effect of a change in root restraint on the models tested would be desirable.\n\nCONCLUDING REMARKS\n\nThe effect of sweepback on the flutter characteristics of a uniform cantilever wing carrying a concentrated weight has been experimentally investigated. The results as presented may be used in conjunction with analytical methods of predicting the flutter speed of sweptback wings carrying concentrated weights to indicate the validity of the methods used.\n```", "timestamp": "2026-07-22T06:41:53.418055+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 18, "total_pages": 52, "image_filename": "19930085911_p18.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 17\n\nThe conditions before heat addition in the combustion chamber, station 5, are calculated for a total-pressure drop across the flame holder of twice the dynamic pressure in front of the flame holder. Thus\n\n$$\n\\frac{P_5}{P_4} = \\frac{P_4 - 2q_4}{P_4} = 1 - \\frac{2q_4}{P_4}\n\\tag{10}\n$$\n\nbut\n\n$$\nq_4 = P_4 \\frac{\\gamma}{2} M_4^2 \\left(1 + \\frac{\\gamma-1}{2} M_4^2\\right)^{-\\frac{\\gamma}{\\gamma-1}}\n\\tag{11}\n$$\n\ntherefore\n\n$$\nP_5 = P_4 - \\frac{\\gamma M_4^2 P_4}{\\left(1 + \\frac{\\gamma-1}{2} M_4^2\\right)^{\\frac{\\gamma}{\\gamma-1}}}\n\\tag{12}\n$$\n\nIn order to determine $M_5$, it is necessary to determine\n\n$$\n\\frac{A_{cr,5}}{A_5} = \\left(\\frac{P_4}{P_5}\\right)\\left(\\frac{A_{cr,4}}{A_4}\\right)\n\\tag{13}\n$$\n\nThe Mach number $M_5$ can then be determined from equation (8).\n\nThe method used to determine the conditions after heat addition (station 6) involves the trial-and-error process of equations (14) and (15), which are derived in reference 4.\n\n$$\n\\left(1 + \\frac{W_f}{W_a}\\right)^2 \\frac{T_6}{T_5} = \\left(\\frac{M_6}{M_5}\\right)^2 \\left(\\frac{\\gamma_6}{\\gamma_5}\\right) \\frac{\\left(1 + \\gamma_5 M_5^2\\right)^2}{\\left(1 + \\gamma_6 M_6^2\\right)^2} \\frac{\\left(1 + \\frac{\\gamma_6-1}{2} M_6^2\\right)}{\\left(1 + \\frac{\\gamma_5-1}{2} M_5^2\\right)}\n\\tag{14}\n$$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:41:53.623685+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 46, "total_pages": 47, "image_filename": "19930083221_p46.jpg", "text": "44\nNACA TN No. 1824\n\n5. Kaplan, Carl: On Similarity Rules for Transonic Flows,\nNACA TN No. 1527, 1948.\n\n6. Lamb, Horace: Hydrodynamics. Dover Publication, N.Y., 1945.\n\n7. Heaslet, Max. A., and Lomax, Harvard: Two-Dimensional Unsteady\nLift Problems in Supersonic Flight. NACA TN No. 1621, 1948.\n\n8. Wagner, Herbert: Uber die Entstehung des dynamischen Auftriebes\nvon Tragflügeln. Z.f.a.M.M., Bd. 5, Heft 1, Feb. 1925,\nS. 17-35.\n\n9. Jones, Robert T.: The Unsteady Lift of a Wing of Finite Aspect\nRatio. NACA Rep. No. 681, 1940.\n\n10. Churchill, Ruel V.: Modern Operational Mathematics in Engineering.\nMcGraw-Hill Book Co., N.Y., 1944.\n\n11. Jahnke, E., and Emde, F.: Tables of Functions. Dover Publica-\ntions, N.Y., 1945.\n\n12. Evvard, John C.: Theoretical Distribution of Lift on Thin\nWings at Supersonic Speeds (An Extension). NACA TN No. 1585,\n1948.\n\n13. Heaslet, Max. A., and Lomax, Harvard: The Use of Source-Sink\nand Doublet Distributions Extended to the Solution of Arbitrary\nBoundary Value Problems in Supersonic Flow. NACA TN No. 1515,\n1948.\n\n14. Hayes, Wallace D.: Linearized Supersonic Flow. North American\nAviation, Inc. Rep. No. AL-222, June, 1947.\n\n15. von Kármán, Th.: The Problem of Resistance in Compressible\nFluids. (Fifth Volta Congress) Roma, Reale Accademia D'Italia,\n1936.\n\n16. Stewart, H. J., and Puckett, A. E.: Aerodynamic Performance of\nDelta Wings at Supersonic Speeds. Jour. Aero. Sci., vol. 14,\nno. 10, Oct. 1947, pp. 567-578.\n\n17. Heaslet, Max. A., and Lomax, Harvard: The Calculation of\nDownwash Behind Supersonic Wings With an Application to\nTriangular Plan Forms. NACA TN No. 1620, 1948.", "timestamp": "2026-07-22T06:41:55.663531+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 76, "total_pages": 114, "image_filename": "19930086061_p76.jpg", "text": "72\nNACA RM L9J07\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n-3\n-2\nP -1\n0\n1\nUpper\nLower\nRight semispan\n\n(a) $\\psi = 0^\\circ$\n\n-4\n-3\n-2\n-1\n0\n1\nLeft semispan\n-4\n-3\n-2\n-1\n0\n1\nUpper\nLower\nRight semispan\n\n(b) $\\psi = 10^\\circ$\n\nFigure 25.- Pressure distribution about wing 3 at various angles of yaw; $\\alpha = 39.1^\\circ$.", "timestamp": "2026-07-22T06:41:57.704117+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 46, "total_pages": 72, "image_filename": "19930085491_p46.jpg", "text": "Equation for fuselage ordinates:\n\n$$\n\\frac{r}{c} = \\left[ 1 - \\left( 1 - \\frac{2x}{l} \\right)^2 \\right]^{\\frac{3}{4}}\n$$\n\nFineness ratio: $\\frac{l}{2r_0} = 12.5$\n\nStreamwise airfoil section:  \nNACA 64A006\n\n[Figure: Diagram showing design dimensions of basic configuration WF-63, including fuselage outline, wing planform, and various dimensional annotations in inches. Key labeled dimensions include: x, 3.500\", l/2 = 4.250\", 6.715\", l = 8.50\", 2.286\", 1.143\", 63°, 21°, .680\", .960\", m.a.c. = 1.600\", .791\", .571\", 5.000\". Dashed lines indicate internal or reference geometry. NACA logo at bottom right.]\n\nCONFIDENTIAL (left margin, vertical)\n\nCONFIDENTIAL (right margin, vertical)\n\nNACA RM No. A8J04 (top right, vertical)\n\nFigure 4.—Design dimensions of basic configuration,WF-63.\n\nNACA (logo at bottom center)\n\n45 (bottom right corner)", "timestamp": "2026-07-22T06:41:59.442371+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 13, "total_pages": 22, "image_filename": "19930085922_p13.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:42:02.226242+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 9, "total_pages": 22, "image_filename": "19930085928_p9.jpg", "text": "8 CONFIDENTIAL NACA RM No. A9A31\n\nof course, desirable because the greatest mass of air per unit of entrance area flows through the system. As indicated on the curves, the maximum recovery at all conditions for either model occurred with a normal shock wave ahead of the inlet. With model A, this shock wave formed ahead of the inlet for all flow ratios at a Mach number of 1.36, and it moved into the inlet only at mass-flow ratios well above those for maximum recovery at the greater Mach numbers. The unexpected result is that, contrary to the pressure variation observed with the convergent-divergent nose inlets of references 3 and 7, the recovery decreased when the shock wave moved toward the throat of the inlet passage. This fact probably means that, although the pressure losses through the shock wave decreased, the total losses increased because of adverse effects of the boundary layer inside the ducts.\n\nSimilar flow characteristics were observed with model B having a slotted inlet. The normal shock wave existed upstream of the inlet for all Mach numbers at the mass-flow ratio for maximum pressure recovery, and it retreated into the duct only at high flow ratios at a Mach number of 2.01. Changing the slot area to enable the shock wave to enter the ducts reduced the recovery. The maximum total-pressure ratio of model B occurred at a cusp in the curve of the variation with mass-flow ratio. The following discussion is suggested as an explanation for the occurrence of this cusp. When a normal shock wave existed upstream of the inlet, the flow through the scoop entrances must have been subsonic and accelerated in the constricted passage. With the relatively large contraction ratio of model B, the flow probably was choked at the throat and expanded to supersonic velocity again in the subsequent expanding channel. It was finally reduced to subsonic velocity through a complex pattern of shock waves inside the diffuser. As the back pressure in the settling chamber was increased, these shock waves moved upstream toward the throat and the pressure rise through them was transmitted forward through the boundary layer. This increased pressure forced more boundary-layer air to flow out of the slots, and the amount increased as the shock waves moved toward the throat. With model B, a sudden rise in pressure recovery occurred when these shock losses formed at the throat, possibly because the balance between the pressure losses through the shock waves and the amount of low-energy air forced out of the slots fulfilled the conditions required for flow with the least pressure loss. With model A, the cusp did not occur, perhaps because the contraction and the slots were better proportioned.\n\nWith both models, the wide range of conditions for which the mass-flow ratio could be varied with only small changes in pressure ratio indicates that the slots and modified duct shape are not only\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:42:03.821445+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 42, "total_pages": 46, "image_filename": "19930085548_p42.jpg", "text": "NACA RM No. E8L30\n41\n\n1077\n\n<!-- Image (145, 289, 881, 661) -->\n\nFigure 19. - Effect of maximum cylinder pressure on indicated thermal efficiency at several fuel-air ratios for both experimental and calculated values of maximum cylinder pressure. Inlet-manifold temperature, 400° F; inlet-manifold pressure, 100 pounds per square inch absolute.", "timestamp": "2026-07-22T06:42:04.885098+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 78, "total_pages": 98, "image_filename": "19930086073_p78.jpg", "text": "```markdown\n76\n\nLift coefficient, $C_L$\n\nAngle of attack, $\\alpha$, deg\n\nAngle of sideslip, $\\beta$, deg\n$\\circ$ 0.0\n$\\square$ 6.0\n$\\diamond$ 12.0\n$\\triangle$ 15.9\n\n(a) $C_L$ vs $\\alpha$.\n\nFigure 17.— Wing plus body plus vertical tail at various angles of sideslip with controls neutral.\n\nNACA RM A9H04\n```", "timestamp": "2026-07-22T06:42:05.466147+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 69, "total_pages": 78, "image_filename": "19930082483_p69.jpg", "text": "256-1159 B\n1032\n\nNACA TN NO. 1807\n\nInlet total probe\n(used only for oper-\nating purposes)\n\nTurbine-inlet static-\npressure taps\n\nShielded-type inlet\ngas thermocouples\n\nNozzle-inlet\nbaffle segments\n\nHoneycomb-type flow\nstraightener\n\nTurbine-discharge,\nstatic-pressure taps\n\nStatic-pressure tap up-\nstream of rotor disk\n\n(3)\n(1)\n(3)\n(2)\n(1)\n(2)\n(1)\nSection A-A\n\nStatic-pressure tap\ndownstream of rotor disk\n\nGas-temperature ther-\nmocouple downstream of\nrotor disk\n\n(1) Tail-cone discharge,\nstatic-pressure taps\n(2) Tail-cone discharge,\ntotal-pressure Kiel\nprobes\n(3) Tail-cone discharge,\nunshielded thermo-\ncouples\n\nTurbine journal-thrust\nbearing surface\n\nGas-temperature thermocouple\nupstream of rotor disk\n\nTurbine journal-bearing surface\n\nTurbine rotor\n\nNACA\n\nFigure 7. - Schematic plan of turbine assembly and instrumentation used in investigation of effect of partial admission on\nperformance of gas turbine.\n\n67", "timestamp": "2026-07-22T06:42:07.835766+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 1, "total_pages": 55, "image_filename": "19930085966_p1.jpg", "text": "FILE COPY\nNO 4\n\nCONFIDENTIAL\n\nCopy 217\nRM L9B17\n\nNACA RM L9B17\n\nNACA\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL DETERMINATION OF THE SUBSONIC PERFORMANCE\nOF A RAM-JET UNIT CONTAINING THIN-PLATE BURNERS\n\nBy\n\nJohn R. Henry\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nAUTHORITY CROWLEY CHANGE#2335\nDATE 1-8-54\nT.C.F.\n\nRETURN TO THE [illegible]\nRESEARCH FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STR. N. W.\nWASHINGTON 25, D. C.\n\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nJune 29, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:42:08.304936+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 11, "total_pages": 47, "image_filename": "19930085919_p11.jpg", "text": "10 CONFIDENTIAL NACA RM No. A9C21\n\nlongitudinal-stability margin should be considered in choosing the type of control surface to be used for balance. With the negative deflection of the elevon limited to $40^\\circ$, the split flap deflected $45^\\circ$ and the center of gravity at 0.25$\\bar{c}$, the wing-fuselage combination could be balanced only for lift coefficients up to 0.46 (fig. 12). However, it appears possible to use a more rearward center of gravity and still to maintain adequate longitudinal stability at the lower lift coefficients. With a more rearward center of gravity, the wing-fuselage combination could be balanced at all lift coefficients with smaller elevon deflections, thus allowing more elevon effectiveness for lateral control.\n\nLeading-Edge Devices\n\nThe model was tested with both the drooped-nose flap and the extended-nose flap deflected $30^\\circ$, $35^\\circ$, $40^\\circ$, and $50^\\circ$. The optimum deflection for either flap was found to be about $40^\\circ$. As only slight differences were noted in the results for the several deflections, only the results for the $40^\\circ$ deflection are presented. The model was also investigated with each of the leading-edge flaps in various combinations with the constant-chord elevon undeflected and deflected negatively $20^\\circ$, and the 0.25-chord split flap undeflected and deflected $45^\\circ$ at the trailing edge of the wing.\n\nThe characteristics of the model with the drooped-nose flap of 50-percent wing span and of full wing span are presented in figures 13 and 14, respectively. The drooped-nose flap of 50-percent wing span decreased the lift at all angles of attack and failed to improve the pitching-moment characteristics of the model (fig. 13). However, the drooped-nose flap of full wing span gave slightly better results, increasing the lift coefficient at which longitudinal instability occurred about 0.15 with the split flap retracted and about 0.04 with the split flap extended (fig. 14). With the elevon deflected $-20^\\circ$, the split flap deflected $45^\\circ$ and the drooped-nose flap of full wing span deflected $40^\\circ$, the lift coefficient attained before the occurrence of longitudinal instability was greater than 1.0 (fig. 14).\n\nThe characteristics of the model with the extended-nose flap of 50-percent wing span and full wing span are presented in figures 15 and 16, respectively. The extended-nose flap of 50-percent wing span proved to be as ineffective as the drooped-nose flap of 50-percent wing span for increasing the lift coefficient of the model before the occurrence of longitudinal instability (figs. 13 and 15). However, with the split flap deflected $45^\\circ$ and the elevon deflected $-20^\\circ$, the extended-nose flap produced a more nearly linear variation\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:42:09.041288+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 28, "total_pages": 33, "image_filename": "19930085900_p28.jpg", "text": "```markdown\nNACA RM L9D20\n27\n\nCONFIDENTIAL\n\nO Chine jets normal to the center line (258 jets)\n□ Chine jets slanted aft at 45° to the center line (258 jets)\n△ Multiple-step jets normal to the center line (245 jets)\nV-steps pointed forward\n\nResistance, lb\nAverage air flow per jet, lb/sec\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:42:10.707411+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 29, "total_pages": 92, "image_filename": "19930085870_p29.jpg", "text": "NACA RM No. L9D07\nCONFIDENTIAL\n29\n\n[Figure: (a) Elliptical leading edge.]\n[Figure: (b) Wedge leading edge.]\n\n[Figure: (c) Wing series.]\n\nFigure 4.- Triangular-wing models.\nCONFIDENTIAL\nNACA\nL-58999", "timestamp": "2026-07-22T06:42:10.902633+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 1, "total_pages": 30, "image_filename": "19930085970_p1.jpg", "text": "Copy 294\nRM A9E09\n\nNACA RM A9E09\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION\nEMPLOYING A WING SWEPT BACK $63^\\circ$.- CHARACTERISTICS\nFOR SYMMETRICAL WING SECTIONS AT HIGH SUBSONIC\nAND MODERATE SUPERSONIC MACH NUMBERS\n\nBy Newton A. Mas\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCLASSIFIED DOCUMENT\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 50:31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nJuly 7, 1949\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: NACA RESEARCH ABSTRACT NO. 115\nEFFECTIVE DATE: MAY 8, 1957\nMHL\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:42:12.290667+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 42, "total_pages": 60, "image_filename": "19930085862_p42.jpg", "text": "40\nNACA RM No. L9A07\n\n<!-- Image (117, 109, 756, 883) -->\n\n(d) Leading-edge and split flaps.\n\n(e) Leading-edge and split flaps and fences.\n\n(f) Leading-edge flaps and fences.\n\nFigure 11.— Concluded.", "timestamp": "2026-07-22T06:42:13.869788+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 77, "total_pages": 114, "image_filename": "19930086061_p77.jpg", "text": "NACA RM L9J07\n73\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n\n-3\n-2\n-1 P\n0\n1\nRight semispan\n\nUpper\nLower\n\n20°\n\n(c) $\\psi = 20^\\circ$\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n\n-3\n-2\n-1 P\n0\n1\nRight semispan\n\nUpper\nLower\n\n35°\n\n(d) $\\psi = 35^\\circ$\n\nNACA\n\nFigure 25.- Concluded.", "timestamp": "2026-07-22T06:42:18.085853+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 52, "total_pages": 62, "image_filename": "19930082918_p52.jpg", "text": "NACA TN 1940\n63\n\n$$\n\\frac{p}{\\Delta p}\n$$\n\nAging time, hr\n\nFigure 12.- Root-mean-square lattice strains as a function of aging time at indicated temperatures for low-carbon M-155 alloy solution-treated 10 hours at 2200° F and water-quenched.", "timestamp": "2026-07-22T06:42:18.578552+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 8, "total_pages": 92, "image_filename": "19930085951_p8.jpg", "text": "6\nUNCLASSIFIED\nCONFIDENTIAL\nNACA RM L9D29\n\nthe shaft tension caused by the spinner-to-tip part of the blade rotating\nin the air stream. Tests were frequently repeated during the test program,\nand, for purposes of comparison, the data are considered accurate to\nwithin 1 percent and the faired envelopes are believed to be accurate to\nwithin much closer limits.\n\nRESULTS AND DISCUSSION\n\nFaired curves of thrust coefficient, power coefficient, and propeller\nefficiency plotted against advance ratio are presented in figures 8 to 16\nfor the two-blade NACA 10-(3)(062)-045A propeller and in figures 17 to 25\nfor the two-blade NACA 10-(3)(05)-045 propeller. Test points are shown on\nthe figures giving thrust and power coefficients. The variation of air-\nstream Mach number and helical-tip Mach number with advance ratio is shown\nin the figures giving propeller efficiency.\n\nEffect of blade-section thickness ratio on envelope efficiency.-\n\nFigure 26 presents a comparison of the envelope efficiencies of\nNACA propellers 10-(3)(08)-03, 10-(3)(08)-03R, and 10-(3)(12)-03 (refer-\nences 3, 5, and 8) at the various rotational speeds. The thinnest blade\nof this group (NACA 10-(3)(08)-03) maintains an envelope efficiency of over\n0.90 throughout the range of advance ratio of the tests at rotational\nspeeds of 1140, 1350, and 1600 rpm. At these rotational speeds and at\nthe design value of advance ratio (2.1) the NACA 10-(3)(08)-03 propeller\nis from 2 to 3 percent more efficient than the round-shank propeller\n(NACA 10-(3)(08)-03R) and from $1\\frac{1}{2}$ to $5\\frac{1}{2}$ percent more efficient than the\nNACA 10-(3)(12)-03 propeller. At the higher rotational speeds (2000 and\n2160 rpm) the envelope efficiencies of all three propellers are reduced.\nThe propeller with the thinnest blade sections suffers the least reduction\nin envelope efficiency, and the propeller with the thickest outboard blade\nsections suffers the greatest loss in envelope efficiency. At 2160 rpm\nand an advance ratio of 0.90 the envelope efficiencies of the\nNACA 10-(3)(08)-03R and NACA 10-(3)(12)-03 propellers are lower by 4 and\n12 percent, respectively, than the envelope efficiency of the\nNACA 10-(3)(08)-03 propeller. These differences in envelope efficiency\nat the higher rotational speeds may be explained by compressibility\neffects which generally lower the lift-drag ratios of thick blade sections\nat high section Mach numbers.\n\nIn figure 26(b) the envelope efficiencies of the NACA propellers in\nthe 0.03-solidity group are compared with the induced, or optimum,\nefficiency of a two-blade propeller with the Betz minimum induced-energy-\nloss loading. This curve of optimum efficiency was calculated by a method\nneglecting all profile-drag losses (reference 10) for a two-blade pro-\npeller operating at the same values of power coefficient as were obtained\nwith the NACA 10-(3)(08)-03 propeller. Although the power coefficients\nfor maximum efficiency are slightly different for the three propellers in\nUNCLASSIFIED\nCONFIDENTIAL", "timestamp": "2026-07-22T06:42:19.685005+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 47, "total_pages": 47, "image_filename": "19930083221_p47.jpg", "text": "NACA TN No. 1824\n45\n\n18. Jones, R.T.: Wing Plan Forms for High-Speed Flight.\n    NACA TN No. 1033, 1946.\n\n19. Robinson, A., and Young, A.D.: Note on the Application of the\n    Linearized Theory for Compressible Flow to Transonic Speeds.\n    The College of Aeronautics, Cranfield, Eng., Rep. no. 2,\n    Jan. 1947.\n\n20. Jones, R.T.: Properties of Low-Aspect-Ratio Pointed Wings\n    at Speeds Below and Above the Speed of Sound. NACA TN\n    No. 1032, 1946.\n\n21. Spreiter, John R.: Aerodynamic Properties of Slender Wing-\n    Body Combinations at Subsonic, Transonic, and Supersonic\n    Speeds. NACA TN No. 1662, 1948.", "timestamp": "2026-07-22T06:42:22.474361+00:00"}