PDF/ocr_dataset_1083.jsonl

{"citation_id": "19930086020", "source_url": "https://ntrs.nasa.gov/api/citations/19930086020/downloads/19930086020.pdf", "page_number": 2, "total_pages": 22, "image_filename": "19930086020_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:58:53.932095+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 11, "total_pages": 24, "image_filename": "19930085991_p11.jpg", "text": "NACA RM L9I28\n9\n\nThe calculated $C_{m\\alpha}$ values for the nose with and without fins for two center-of-gravity locations are plotted in figure 5. On the figure the finned nose is indicated as being statically stable for both center-of-gravity locations investigated. The empirical results obtained with model 1, however, indicated that only for the more forward center-of-gravity location did the finned nose damp applied rotation and descend in a stable nose-down attitude. It thus appears that dynamic stability must be given consideration either through calculations or through empirical corrections. For comparison with the calculated values, measured values of $C_{m\\alpha}$ obtained from force and moment tests of a large model similar to model 1 with and without fins are also plotted in figure 5. As can be seen, the calculated and measured values of $C_{m\\alpha}$ for the nose without fins were in fairly close agreement. Values of $C_{m\\alpha}$ measured for the nose section with the fins installed indicate that adding the fins had a greater stabilizing effectiveness than was indicated by the calculations, probably because interference effects of the fins on the flow over the nose section caused an additional increase in the stabilizing effectiveness of the fins.\n\nThe aforementioned work has been done on the basis of free-spinning-tunnel tests at airspeeds up to 60 miles per hour which, based on a scale range of about 1/10 to 1/23 for the various dynamic models tested, correspond to full-scale airspeeds up to 300 miles per hour. However, it is interesting to note that preliminary analysis of recent NACA higher-speed investigations has also indicated that fins were effective in stabilizing nose sections. In one instance (results unpublished), a smaller model of one of the free-spinning-tunnel nose sections was released with and without stabilizing fins in an atmospheric horizontal wind tunnel at sea-level airspeeds up to 150 miles per hour, simulating full-scale airspeeds up to 750 miles per hour (when compressibility effects were neglected). When stabilized with fins, the model descended to the floor of the tunnel in stable nose-forward flight; whereas without fins it turned away from a nose-first flight attitude. In another instance (results unpublished), duplicates of two of the free-spinning-tunnel models were fired with and without stabilizing fins at a Mach number of 1.2 (actual model speed) in a Langley free-flight apparatus, and the results obtained were similar to those obtained during the atmospheric horizontal wind-tunnel tests. In another investigation (reference 6), the Langley Pilotless Aircraft Research Division stabilized a large model of one of the free-spinning-tunnel nose sections with fins selected on the basis of the free-spinning-tunnel investigation and forcibly jettisoned the nose from an afterbody of a test rocket in flight at a Mach number of about 0.87. The nose traveled stably after leaving the afterbody.", "timestamp": "2026-07-22T06:58:54.899860+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 23, "total_pages": 55, "image_filename": "19930085966_p23.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:58:56.766466+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 10, "total_pages": 32, "image_filename": "19930085992_p10.jpg", "text": "8\nNACA RM L9E17\n\nmounted at the wing tip. In the following table these parameters have been compared for two different chordwise positions of weights 1 and 2:\n\n| Weight | Weight shape (see figs. 3(e) to 3(d)) | Weight location (percent chord) | $\\frac{W_w}{W}$ | $\\frac{I_w}{I_{RA}}$ | $\\frac{(v_1)_w}{v_1}$ | $\\frac{(f_e)_w}{f_e}$ | Run (see table I) |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1a | Streamlined | 47.3 | 0.0744 | 0.380 | 0.845 | 0.846 | 2 |\n| 2a | Nonstreamlined | 47.3 | .0755 | .403 | .874 | .754 | 3 |\n| 1b | Streamlined | 19.4 | .0754 | .509 | .763 | .824 | 5 |\n| 2b | Nonstreamlined | 19.8 | .0755 | .507 | .792 | .904 | 6 |\n\nExamination of the flutter speeds shows a maximum difference of $3\\frac{1}{2}$ percent between streamlined and nonstreamlined shapes. A radical change in the shape of weights located at the tip appears to have produced only a small change in flutter speed. A somewhat greater effect of shape on flutter frequency is noted with the changes occurring in opposite directions for the two different chordwise positions.\n\nThe effects of aerodynamic shape of weights for a wide range of spanwise positions are presented in figures 4 and 5 for weight 3 (fig. 3(g)) and weight 4 (fig. 3(h)). Comparison of the flutter speeds in figure 4 for streamlined and nonstreamlined shapes shows a difference of not more than 4 percent at any point along the span. Thus, a radical change in aerodynamic shape of weights at any spanwise location has produced only a small change in flutter speed, although for most spanwise positions as well as for the tip position the flutter speed was lower for the streamlined shape than for the nonstreamlined shape. Examination of figure 5 shows that with the exception of the tip position the flutter frequency differed by less than 3 percent between streamlined and nonstreamlined shapes at any point along the span.\n\nAlthough the effect on flutter of aerodynamic shape of the weights is shown to be small, it should be remarked that shape may be very significant in regard to such static aeroelastic instabilities", "timestamp": "2026-07-22T06:58:58.453234+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 17, "total_pages": 17, "image_filename": "19930085988_p17.jpg", "text": "CONFIDENTIAL\n\n$C_D$\n.3\n.2\n.1\n0\n.8 .9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9\nM\n\nPresent test\n\nFree-fall model\nreference 1\n\nCONFIDENTIAL\n\nNACA\n\nFigure 8.— Variation of total-drag coefficient with Mach number for wingless configuration and free-fall model. $C_D$ based on body frontal area.\n\nNACA-Langley - 10-27-49 - 575\nNACA RM L9H30\n16", "timestamp": "2026-07-22T06:58:59.327053+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 37, "total_pages": 47, "image_filename": "19930085919_p37.jpg", "text": "```markdown\n36\n\nCONFIDENTIAL\n\nLift coefficient, $C_L$\n\n$\\delta_e$, deg\n0\n-10\n-20\n-30\n-40\n\nElevon\n0.25c split flap\nShort fuselage\n100c\n\nAngle of attack, $\\alpha$, deg\n\nPitching-moment coefficient, $C_m$\n\n10 $\\delta_{sf}$, deg\n\nNACA\n\n(a) $C_L$ vs $\\alpha$ and $C_m$.\n\nFigure 12.- Effectiveness of the constant-chord elevon with the split flap deflected 45°. R, 4.2 x 10$^6$.\n\nNACA RM No. A9C21\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:59:01.122126+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 102, "total_pages": 114, "image_filename": "19930086061_p102.jpg", "text": "98\nNACA RM L9J07\n\n$$ \\frac{c_l c}{C_{L_{Cav}}.8} $$\n$$ \\frac{y}{b/2} \\text{, percent} $$\n(a) Angles of attack : 4.1°, 8.1°.\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 4.1 | 0.11 |\n| 8.1 | 0.22 |\n\n$$ \\frac{c_l c}{C_{L_{Cav}}} $$\n$$ \\frac{y}{b/2} \\text{, percent} $$\n(b) Angles of attack : 14.1°, 24.1°, 34.1°.\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 14.1 | 0.38 |\n| 24.1 | 0.63 |\n| 34.1 | 0.85 |\n\n$$ \\frac{c_l c}{C_{L_{Cav}}.8} $$\n$$ \\frac{y}{b/2} \\text{, percent} $$\n(c) Angles of attack : 39.1°, 44.1°, 50.1°.\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 39.1 | 0.94 |\n| 44.1 | 0.94 |\n| 50.1 | 0.57 |\n\nFigure 47.- Span load distribution of wing 3 at various angles of attack; $\\Psi = 20^\\circ$. Flagged symbols represent data taken with left semispan at $\\Psi = -20^\\circ$.", "timestamp": "2026-07-22T06:59:02.254024+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 92, "total_pages": 98, "image_filename": "19930086073_p92.jpg", "text": "90\n\n12\n10\n8\n6\n4\n2\n0\n0 .2 .4 .6 .8 1.0 1.2 1.4\nLift coefficient, $C_L$\nLift-drag ratio, L/D\nWing alone\nWing + body\nWing + body + vertical tail\n(a) Effect of body and vertical tail.\n\n12\n10\n8\n6\n4\n2\n0\n0 .2 .4 .6 .8 1.0 1.2 1.4\nLift coefficient, $C_L$\nLift-drag ratio, L/D\nFlap deflection, $\\delta_f$, deg\n0\n-20.7\n-10.8\n20.4\n45.4\nNACA\n(b) Effect of deflecting flaps ; wing + body.\n\nFigure 22.- Variation of lift-drag ratio with lift coefficient.\n\nNACA RM A9E04", "timestamp": "2026-07-22T06:59:05.722556+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 34, "total_pages": 92, "image_filename": "19930085951_p34.jpg", "text": "32\nNACA RM L9D29\n\n[CONFIDENTIAL]\n\nThrust coefficient, $C_T$, and power coefficient, $C_P$\n\nAdvance ratio, J\n\n(a) Thrust and power coefficients.\n\nFigure 10.- Characteristics of NACA 10-(3)(062)-045A propeller.\nRotational speed, 1500 rpm; $\\beta_{0.75R} = 45^\\circ$.\n\n[UNCLASSIFIED]", "timestamp": "2026-07-22T06:59:06.053712+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 55, "total_pages": 92, "image_filename": "19930085870_p55.jpg", "text": "```markdown\n56\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {○ CL\n                 {□ Cm\nWedge L.E.      {△ CL\n                 {◇ Cm\n.16\n.08\nCL\n0\n-.08\n-.16\n-.24\n.01\nCm\n0\n-.01\n\n.06\nElliptical L.E. {○ CD\n                 {□ L/D\nWedge L.E.      {△ CD\n                 {◇ L/D\n.04\nCD\n.02\n0\n-8\n-6\n-4\n-2\n0\n2\n4\n6\n8\n0\n6\n4\nL/D\n2\n0\nα, deg\n\n(d) Wing 4. w=1.017; R=860,000.\nFigure 7 .- Continued.\nCONFIDENTIAL\n\n[Figure: NACA logo]\n```", "timestamp": "2026-07-22T06:59:07.934974+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 65, "total_pages": 149, "image_filename": "19930083192_p65.jpg", "text": "NACA TN 1976\n61\n\nThe disturbed motions of an airplane in gusty air do not correlate with the maximum gust intensities experienced. Results indicate that the disturbed motions may change the loads in gusty air by 20 to 50 percent.\n\nCONCLUDING REMARKS\n\nThe available information on gust loads does not permit the determination of applied loads under actual operating conditions on an absolute basis. Lack of information on gust statistics, the behavior of airplanes in rough air, transient aerodynamics, and operating conditions make such an estimation of doubtful accuracy. Consequently, the procedure has been to transfer loads data from a reference airplane to new airplanes by using available knowledge of aircraft behavior. Statistical data on flight loads and gust experience under actual operating conditions are collected in order to set the level of loading, and \"transfer\" coefficients based on a knowledge of airplane reaction are then calculated and applied to new airplanes. No radical change in operating conditions or airplane characteristics are assumed to be introduced. In addition, it is implied that all airplanes are of the same general character (conventional) and that the relative loads for single isolated gusts are a measure of the relative loads in a sequence of gusts. The important features and implications in this procedure are that the same procedures must be used to evaluate loads data as are used in loads calculations. In addition, the procedure presupposes no drastic change in airplane configuration. The use of the same procedures of calculation tends to eliminate errors that are due to inability to predict airplane reactions on an absolute basis in that errors in the evaluation of loads data cancel those in the calculation of loads.\n\nExtensive changes in configuration may be a serious limitation since the procedure assumes in its general application that the effect of pitch on gust loads is the same for all aircraft and that all aircraft respond to the same gusts (as function of airplane size) to experience significant loads. When new configurations such as tailless airplanes or canard airplanes are considered, the load transfer must be based on the relative motions of the reference airplane and the new airplane and not on the motions of the new airplane alone. A second element, gust selection, has not yet been resolved but it would appear that tailless airplanes, because of small damping in pitch, and canard airplanes, because of the adverse initial pitching moment in a gust for swept-wing airplanes, might select some other ranges of gust sizes than those selected by conventional aircraft. So far, no information is available on this problem and it can only be assumed that available gust-structure data are adequate.", "timestamp": "2026-07-22T06:59:08.255416+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 31, "total_pages": 34, "image_filename": "19930085913_p31.jpg", "text": "30\nNACA RM L9F24\n\n[Figure: A graph plotting Frequency, cps against Distance of weight from root, percent l. The legend indicates circles for Natural frequency and squares for Flutter frequency. The graph contains curves labeled 1st, 2nd, and 3rd.]\n\n(e) Model C; $\\Lambda = 60^\\circ$; $e_w = -1$.\n\nFigure 3.— Continued.", "timestamp": "2026-07-22T06:59:09.646940+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 22, "total_pages": 30, "image_filename": "19930085975_p22.jpg", "text": "20\nNACA RM L9E10\n\nCONFIDENTIAL\n\nWing-tip helix angle, pb/2V, radians\n\n| $\\delta_{al}$ | $\\delta_{ar}$ |\n| :--- | :--- |\n| $\\diamond$ 8 | -8 |\n| $\\triangle$ 4 | -4 |\n| $\\nabla$ -4 | 4 |\n| $\\square$ -8 | 8 |\n\nMach number, M\n\nNACA\n\nFigure 9.- The variation with Mach number of the wing-tip helix angle $\\frac{pb}{2V}$ for various aileron deflections at several angles of attack. $\\Lambda_{c/4} = 3.6^\\circ$.", "timestamp": "2026-07-22T06:59:09.839728+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 18, "total_pages": 46, "image_filename": "19930085983_p18.jpg", "text": "```markdown\nCONFIDENTIAL\n\nNote: All dimensions given in feet unless otherwise specified.\n\nTypical section parallel to plane of symmetry\n\nAll sections have NACA a=1.0 mean-camber lines and 64A005 thickness distributions. (See Table I for section coordinates.)\n\nSpanwise camber distribution\n\n| Station | Percent semispan | Camber $y_c/c$ |\n| :--- | :--- | :--- |\n| $c_0$ | 0 | 0 |\n| $c_1$ | 20 | .0082 |\n| $c_2$ | 40 | .0108 |\n| $c_3$ | 60 | .0117 |\n| $c_4$ | 80 | .0115 |\n| $c_5$ | 100 | .0114 |\n\n[Figure: Diagram of a wing planform showing sections $c_0$ through $c_5$, fuselage area, and dimensions. $c_0=1.714$, $c_5=.429$. Vertical dimensions: .375, .750, 1.125, 1.500, 1.875. Angle $63^\\circ$.]\n\n[Graph: Plot of Spanwise station, percent semispan (0 to 100) vs Angle of twist, $a_1$, deg (-5 to 0). Curves for \"Model static twist\" and \"Theoretical twist ($C_L=.25, M=1.5$)\"]\n\nFigure 3.— Plan form of right half of wing showing spanwise variation of camber and twist and location of sections for which coordinates have been calculated.\n\nNACA\n\nNACA RM A9J27\n\n16\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:59:12.575750+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 23, "total_pages": 30, "image_filename": "19930085970_p23.jpg", "text": "NACA RM A59B09\nCONFIDENTIAL\n\nM\nO 0.50\n□ 0.70\n◇ 0.80\n△ 0.90\n▽ 0.95\n◁ 1.09\n▷ 1.14\n◁ 1.24\n▽ 1.51\n\nLift coefficient, $C_L$\n.6\n.4\n.2\n0\n-.2\n\nAngle of attack, $\\alpha$, deg (for M = .50)\n-4\n0\n4\n8\n\n[Figure: Graph showing variations of lift coefficient with angle of attack at various test Mach numbers.]\n\nFigure 5- Variations of lift coefficient with angle of attack at the various test Mach numbers.\n\nCONFIDENTIAL\n21", "timestamp": "2026-07-22T06:59:13.666729+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 22, "total_pages": 25, "image_filename": "19930085979_p22.jpg", "text": "NACA RM E9E12\n\nModel-mass flow, lb/sec\n32\n28\n24\n20\n16\n12\n\nTunnel-air\nvelocity\n(ft/sec)\nO 205\n□ 275\n◇ 355\n△ 435\n\n— Experimental\n--- Constant volume\n\n0 10 20 30 40 50 60 70 80 90\nModel-air-temperature rise, °F\n\n[Figure: Graph showing variation of model-mass flow with model-air-temperature rise for several tunnel-air velocities. Four sets of data points (circles, squares, diamonds, triangles) correspond to tunnel-air velocities of 205, 275, 355, and 435 ft/sec respectively. Each set includes experimental data (solid line) and constant volume prediction (dashed line). NACA logo in bottom right corner.]\n\nFigure 8. - Variation of model-mass flow with model-air-temperature rise for several tunnel-air velocities. Angle of attack, 0°; tunnel total temperature, 0° F.\n\n21", "timestamp": "2026-07-22T06:59:14.302586+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 8, "total_pages": 20, "image_filename": "19930085999_p8.jpg", "text": "6\nNACA RM E9I07\n\nA vibration signal under similar conditions is presented in\nfigure 3(b) in which the second strain-gage signal has been short-\ncircuited in order to show the absence of electrical interference\nin the instrumentation used for amplification, observation, and\nrecording of strain-gage signals.\n\nVibration was observed in the turbine blade with 0.03-inch\nfreedom of tip movement and also in the blade having 0.06-inch ampli-\ntude at the tip. An analysis of the data obtained from the turbine\nblade with the 0.03-inch amplitude of tip movement is presented in\nfigure 4. The symbols indicate the occurrence of vibration observed\nduring the investigation. The order lines indicate the loci of\npoints of which the frequency is a definite multiple of the turbine\nspeed. These lines show the frequency of any exciting force that\ncan occur at the definite multiples of turbine speed.\n\nFor example, the turbojet engine used in the investigation has\n14 combustion chambers. From consideration of the geometry of the\nengine, it is therefore possible that excitation forces exist at a\nfourteenth multiple of the turbine speed. The critical-speed dia-\ngram (fig. 4) shows that vibration was observed at orders of excita-\ntion, which could be attributed to interruption of the gas forces by\nthe combustion chambers and nozzle vanes, and at several orders of\nrotative speed, probably because of inequalities in mass flow.\n\nThe turbine blade vibrated in the first bending mode at approxi-\nmately 1150 cycles per second and in the first torsional mode at\nabout 1900 cycles per second. These mode frequencies were determined\nduring previous bench tests of turbine blades. Complex modes of\nvibration at higher frequencies were also excited, apparently by the\nhigher orders of intermittent gas loading. Only one of these complex-\nmode vibrations occurred within the service-cruising speed range,\napproximately 9600 to 11,500 rpm.\n\nThe critical speeds and frequencies obtained from operation of\na turbine blade with 0.06-inch amplitude of tip movement are pre-\nsented in figure 5 in the same manner as in figure 4. The turbine\nblade vibrated in the first bending mode at approximately 1150 cycles\nper second and in the first torsional mode at about 1800 cycles per\nsecond. Only one complex-mode vibration was of appreciable magnitude;\nthe excitation-force frequency coincided with the twenty-eighth order\nof turbine speed, which is also the second order of the 14 combustion\nchambers. No significant differences in vibratory-stress levels were\nobserved between this turbine blade and the blade with the lesser\namplitude of tip movement.", "timestamp": "2026-07-22T06:59:14.967033+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 45, "total_pages": 52, "image_filename": "19930085911_p45.jpg", "text": "44 CONFIDENTIAL NACA RM E9F22\n\nFree-stream static temperature, $T_0$, °R\n500\n450\n400\n\nFree-stream static pressure, $P_0$, lb/sq ft\n2000\n1500\n1000\n500\n\nFree-stream total pressure, $P_0$, lb/sq ft\n8000\n6000\n4000\n2000\n0\n\nFuel flow, $W_f$ x 3600, lb/hr\n8000\n6000\n4000\n2000\n0\n\n$W_f$\n$P_0$\n\nTime after release, $\\tau$, sec\n0 10 20 30 40 50\n\n(b) Independent test variables.\n\nFigure 10. - Continued. Time history of flight data and performance of ram-jet unit 16-A-5.\n\nCONFIDENTIAL\n\n1152", "timestamp": "2026-07-22T06:59:15.602532+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 14, "total_pages": 40, "image_filename": "19930085997_p14.jpg", "text": "12\nCONFIDENTIAL\nNACA RM A9I29\n\nREFERENCES\n\n1. Davis, Wallace F., Edwards, Sherman S., and Brajnikoff, George B.:\n   Experimental Investigation at Supersonic Speeds of Twin-Scoop\n   Duct Inlets of Equal Area. IV - Some Effects of Internal Duct\n   Shape Upon an Inlet Enclosing 37.2 Percent of the Forebody\n   Circumference. NACA RM A9A31, 1949.\n\n2. Davis, Wallace F., Brajnikoff, George B., Goldstein, David L., and\n   Spiegel, Joseph M.: An Experimental Investigation at Supersonic\n   Speeds of Annular Duct Inlets Situated in a Region of Appreciable\n   Boundary Layer. NACA RM A7G15, 1947.\n\n3. Allen, H. Julian: The Asymmetric Adjustable Supersonic Nozzle for\n   Wind-Tunnel Application. NACA RM A8E17, 1948.\n\n4. Ferri, Antonio, and Nucci, Louis M.: Preliminary Investigation of\n   a New Type of Supersonic Inlet. NACA RM L6J31, 1946.\n\n5. Keenan, Joseph H., and Kaye, Joseph: A Table of Thermodynamic\n   Properties of Air. Jour. App. Mech., vol. 10, no. 3, Sept. 1943,\n   pp. A-123-A130.\n\n6. Durand, W. F.: Aerodynamic Theory, vol. III, The Mechanics of\n   Viscous Fluids, sec. 23, J. Springer (Berlin), 1934, pp. 145-154.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:59:25.107934+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 18, "total_pages": 132, "image_filename": "19930085990_p18.jpg", "text": "16 CONFIDENTIAL NACA RM A9I01\n\nbeen neglected. The results of these calculations are presented in figure 49. The minimum power-off sinking speed was 46 feet per second and occurred at a forward speed of 175 miles per hour.\n\nSUMMARY OF RESULTS\n\nThe results of wind-tunnel tests at Mach numbers up to 0.95 of a semispan model of a hypothetical supersonic airplane with the horizontal tail mounted alternately in the extended wing-chord plane and 0.696 of the wing mean aerodynamic chord above the extended wing-chord plane have been presented. A summary of these results follows:\n\n1. At a lift coefficient of zero, the Mach number for drag divergence was about 0.92. There was a smooth increase of lift-curve slope with increasing Mach number up to a Mach number of 0.95.\n\n2. The contribution of the horizontal tail to the static longitudinal stability decreased with increasing Mach number. This decrease was due primarily to the increase with increasing Mach number in the rate of change with angle of attack of the effective angle of downwash at the tail. With the horizontal tail in the extended wing-chord plane, a further destabilizing effect was the decrease in dynamic-pressure ratio at the tail with increasing Mach number.\n\n3. With the horizontal tail in the extended wing-chord plane, the model was longitudinally unstable at Mach numbers above 0.87 at lift coefficients less than 0.3. With the horizontal tail 0.696 of the wing mean aerodynamic chord above the extended wing-chord plane, the model was longitudinally stable at all lift coefficients for all Mach numbers for which data were obtained.\n\n4. Either an all-movable stabilizer or a fixed stabilizer with a constant-chord elevator provided sufficient longitudinal control to balance the model throughout the test range of Mach numbers.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:59:26.725914+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 12, "total_pages": 24, "image_filename": "19930085991_p12.jpg", "text": "10\nNACA RM L9I28\n\nAlthough, as previously mentioned, a pilot jettisoned at transonic\nspeeds in a stabilized nose would not be subjected to excessive\naccelerations, at high supersonic speeds it is conceivable that even a\nstabilized nose may require the use of a controlled auxiliary propulsive\nforce to allow a gradual decrease in airspeed and thus prevent\ndecelerations high enough to endanger the pilot.\n\nCONCLUDING REMARKS\n\nAn empirical criterion based on investigations of five jettisonable\nnose configurations in the Langley 20-foot free-spinning tunnel has\nbeen developed which indicates the fin area required for stabilizing an\nairplane jettisonable nose section.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.", "timestamp": "2026-07-22T06:59:26.941442+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 24, "total_pages": 55, "image_filename": "19930085966_p24.jpg", "text": "NACA RM L9B17\nCONFIDENTIAL\n23\n\nBURNER MOUNTING LUG\nMAIN DISTRIBUTOR TUBE\nMAIN FEED LINE\nPILOT FLAME\n10 MM SPARK PLUG\nPILOT FEED LINE\nPILOT IGNITION JET\nCOMMON-RAIL\nIGNITER TUBE\nPILOT DISTRIBUTOR\nRAKE\nPILOT AIR INTAKE\nMAIN BOILER\nLARGE MAIN VAPOR JETS\nSMALL MAIN VAPOR JETS\nPILOT HOUSING\n\n[Figure: Diagram of a thin-plate-burner unit with nonvaporizing pilot burner.]\n\nNACA\nFigure 2.- Thin-plate-burner unit with nonvaporizing pilot burner.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:59:27.167275+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 93, "total_pages": 98, "image_filename": "19930086073_p93.jpg", "text": "NACA RM A9H04\n91\n\n$$C_{l\\beta}$$\n-0.004\n-0.002\n0\n0 .2 .4 .6 .8 1.0 1.2\n\n$$C_{n\\beta}$$\n.002\n0\n-0.002\n-0.004\n-0.006\n0 .2 .4 .6 .8 1.0 1.2\n\nWing alone\nWing + body\nWing + body + vertical tail\n\n$$C_{Y\\beta}$$\n-0.012\n-0.008\n-0.004\n0\n0 .2 .4 .6 .8 1.0 1.2\n\nLift coefficient, $$C_L$$\n\nFigure 23.— Effect of body and vertical tail on the stability derivatives.", "timestamp": "2026-07-22T06:59:33.066196+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 38, "total_pages": 47, "image_filename": "19930085919_p38.jpg", "text": "```markdown\nNACA RM No. A9G21\n\nCONFIDENTIAL\n\nRolling-moment coefficient, $C_l$\n\n$\\delta_e$, deg\n-40\n-30\n-20\n-10\n10\n\nAngle of attack, $\\alpha$, deg\n\n(b) $C_l$ vs $\\alpha$.\n\nFigure 12.- Concluded.\n\nCONFIDENTIAL\n\n37\n\n```", "timestamp": "2026-07-22T06:59:34.811080+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 35, "total_pages": 92, "image_filename": "19930085951_p35.jpg", "text": "NACA RM L9D29\n33\n\n[Figure: Graph with grid background. Y-axis (left): Efficiency, $\\eta$, scale 0 to 1.0. Y-axis (right): Mach number, M, scale 0 to 1.4. X-axis: Advance ratio, J, scale 1.4 to 3.2.\nThree curves are plotted:\n1. A curve labeled $\\eta$ starts near (2.0, 0.9), peaks slightly, then drops sharply around J=2.5.\n2. A line labeled \"Helical-tip Mach number\" rises from approx (2.0, 0.4) to (2.6, 0.45).\n3. A line labeled \"Air-stream Mach number\" rises from approx (2.0, 0.2) to (2.6, 0.28).\nA NACA logo is visible in the bottom right corner of the plot area.]\n\n(b) Efficiency.\nFigure 10.- Concluded. Rotational speed, 1500 rpm.", "timestamp": "2026-07-22T06:59:35.507891+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 56, "total_pages": 92, "image_filename": "19930085870_p56.jpg", "text": "```markdown\nNACA RM No. L9D07\n57\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O CL, □ Cm\nWedge L.E. {△ CL, ◇ Cm\n.16\n.08\nCL\n0\n-.08\n-.16\n-.24\n.01\nCm\n0\n-.01\n\n.06\nElliptical L.E. {O CD, □ L/D\nWedge L.E. {△ CD, ◇ L/D\n.04\nCD\n.02\n0\n-8\n-6\n-4\n-2\n0\n2\n4\n6\n8\nα, deg\n6\n4\nL/D\n2\n0\n[NACA logo]\n\n(e) Wing 5. w=1.156; R = 770,000.\nFigure 7. - Continued.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:59:35.631861+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 103, "total_pages": 114, "image_filename": "19930086061_p103.jpg", "text": "NACA RM L9J07\n99\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 4.1 | 0.10 |\n| 8.1 | 0.18 |\n\n$$ \\frac{y}{b/2} $$ , percent\n\n(a) Angles of attack : 4.1, 8.1.\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 14.1 | 0.34 |\n| 24.1 | 0.52 |\n| 34.1 | 0.64 |\n\n$$ \\frac{y}{b/2} $$ , percent\n\n(b) Angles of attack : 14.1°, 24.1°, 34.1°.\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 39.1 | 0.74 |\n| 44.1 | 0.80 |\n| 50.1 | 0.80 |\n\n$$ \\frac{y}{b/2} $$ , percent\n\n(c) Angles of attack : 39.1°, 44.1°, 50.1°.\n\nFigure 48.- Span load distribution of wing 3 at various angles of attack; $\\psi = 35^\\circ$. Flagged symbols represent data taken with left semispan at $\\psi = -35^\\circ$.", "timestamp": "2026-07-22T06:59:35.827633+00:00"}
{"citation_id": "19930086083", "source_url": "https://ntrs.nasa.gov/api/citations/19930086083/downloads/19930086083.pdf", "page_number": 1, "total_pages": 48, "image_filename": "19930086083_p1.jpg", "text": "```markdown\nFILE COPY\nNO 2\n\nRESTRICTED\n\nCopy\nRM L9F10\n\nNACA RM L9F10\n\nNACA\n\nRESEARCH MEMORANDUM\n\nPRELIMINARY AERODYNAMIC INVESTIGATION OF THE EFFECT\nOF CAMBER ON A 60° DELTA WING WITH\nROUND AND BEVELED LEADING EDGES\n\nBy John M. Riebe and Joseph E. Fikes\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\nRETURN TO THE ABOVE ADDRESS\nREQUEST FOR PUBLICATION SHOULD BE ADDRESSED\nAS FOLLOWS:\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STR., N. W.\nWASHINGTON 25, D. C.\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\n\nAUTHORITY J. W. CROWLEY\n\nDATE 12-14-53 CHANGE #1912 T.L.B.\n\nCLASSIFIED DOCUMENT\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, U.S.C. 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nAugust 16, 1949\n\nRESTRICTED\n\n194\n```", "timestamp": "2026-07-22T06:59:40.089179+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 23, "total_pages": 30, "image_filename": "19930085975_p23.jpg", "text": "NACA RM L9E10\n21\n\nCONFIDENTIAL\n\nWing-tip helix angle, pb/2V, radians\n\nMach number, M\n\n| $\\delta_{a_l}$ | $\\delta_{a_r}$ |\n| :--- | :--- |\n| $\\diamond$ 8 | -8 |\n| $\\triangle$ 4 | -4 |\n| $\\nabla$ -4 | 4 |\n| $\\square$ -8 | 8 |\n\nNACA\n\nFigure 10.- The variation with Mach number of the wing-tip helix angle $\\frac{pb}{2V}$ for various aileron deflections at several angles of attack. $\\Lambda_{c/4} = 32.6^\\circ$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:59:40.596441+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 66, "total_pages": 149, "image_filename": "19930083192_p66.jpg", "text": "62\nNACA TN 1976\n\nIn some generalized solutions of load prediction, a further restriction has been introduced by the substitution of wing loading for mass parameter as a significant variable. The use of an alleviation factor based on such a substitution is still further restricted by the assumption that all airplanes have the same relation between wing loading and mass parameter. Information presented on the load transfer coefficient indicates that the load transfer coefficient has appreciable spread when this substitution is made and the spread seems to be a function of airplane type and use since the data can be grouped according to land transports, flying boats, and personal airplanes. If wing loading is used, therefore, different relations should be considered according to airplane class and use.\n\nIn cases where the dynamic response may be a factor, it should be considered as an amplifying factor for the incremental loads whether for studies of large loads or repeated loads. Dynamic response should be considered on a relative basis since the uncertainties in calculations due to lack of knowledge and practical simplifications have not been resolved.\n\nOf the many related problems, that of gust-alleviating systems is of great importance. Such systems are of particular interest in reducing loads and improving riding comfort for high-speed aircraft. The two problems may not have compatible solutions since load reduction implies the reduction of stresses in all critical members without impairing other qualities of the airplane while the improvement in riding comfort implies the reduction of airplane motions and accelerations. Although the basic problem of reducing the load due to a single gust is amenable to analysis, the related problems of maintaining other airplane qualities and of determining alleviation in rough air introduce difficulties.\n\nThree fundamental systems of gust alleviation are available and consist of changing airplane characteristics, aeroelastic systems, and aeromechanical systems. The first method consists of emphasizing the pertinent characteristics of the airplane, such as increasing the wing loading to reduce accelerations, decreasing the slope of the lift curve (by use of vents or sweep and by changing the aspect ratio), adjusting the stability characteristics to obtain favorable response in rough air, or using free-floating flaps. Aeroelastic systems generally involve torsional deflections of the wing due to imposed loads. Such systems can be designed by incorporating slant hinges or flaps operated by wing deflections or by the adjustment of the torsional and bending rigidity. Aeroelastic systems have the advantage of being inherent in the construction of the airplane and the disadvantage that they form an additional source of elastic phenomena such as flutter or dynamic response due to a series of gusts. Aeromechanical systems cover combinations of servomechanisms, detectors, and aerodynamic controls. Some detector systems", "timestamp": "2026-07-22T06:59:41.691653+00:00"}
{"citation_id": "19930086020", "source_url": "https://ntrs.nasa.gov/api/citations/19930086020/downloads/19930086020.pdf", "page_number": 3, "total_pages": 22, "image_filename": "19930086020_p3.jpg", "text": "NACA RM A9J06 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nTHE EFFECTS OF SCALE AND TEST TECHNIQUE ON THE VALIDITY OF SMALL-SCALE MEASUREMENTS OF THE AERODYNAMIC CHARACTERISTICS OF A WING WITH THE LEADING EDGE SWEPT BACK $63^\\circ$\n\nBy L. Stewart Rolls\n\nSUMMARY\n\nThe lift and pitching-moment characteristics of two wings of the same plan form (aspect ratio 3.5, taper ratio 0.25, and leading-edge sweep angle $63^\\circ$) have been measured by the NACA wing-flow method in the Mach number range 0.52 to 1.11 and Reynolds number range 0.39 million to 0.81 million. One wing had a symmetrical airfoil section and no twist, while the other was cambered and twisted to support a uniform load distribution at a lift coefficient of 0.25 at a Mach number of 1.5.\n\nThe data are compared with the results from tests of similar models in the Ames 12-foot pressure wind tunnel at Reynolds numbers of approximately 2 million. The comparison shows appreciable discrepancy in the measured pitching-moment characteristics. Changes in the model configuration and test procedure were investigated, but no conclusive explanation of the discrepancy was developed. It is concluded that any attempt to determine the pitching-moment characteristics of highly swept-back wings is inadvisable at such small scale and at such low Reynolds numbers with semispan models.\n\nINTRODUCTION\n\nAs a continuation of a general investigation of the aerodynamic characteristics of a wing with the leading edge swept back $63^\\circ$, tests were conducted by the wing-flow method in order to obtain data bracketing a Mach number of 1.0. One of the models for the wing-flow tests had a symmetrical airfoil and no twist, while the other model was cambered and twisted to support a uniform load distribution at a lift coefficient of 0.25 at a Mach number of 1.5. The results of previous tests of the symmetrical wing are presented in references 1, 2, and 3, while the results of tests of the cambered and twisted wing are contained in references 4 and 5.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:59:41.831206+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 11, "total_pages": 32, "image_filename": "19930085992_p11.jpg", "text": "NACA RM L9EL7\n9\n\nas wing divergence. However, the static cases are not considered in this investigation.\n\nEffects of Spanwise Variation of Light Concentrated Weights\n\nFor these tests, the influence of aerodynamic shape was shown to be small as the spanwise location of the light concentrated weights was varied from root to tip. The general reduction in flutter speed, similar to that shown in reference 1 for weights having a comparable chordwise center-of-gravity location, may therefore be attributed wholly to the effect of the concentrated weights. In the present investigation a maximum reduction in flutter speed of 17 percent was obtained with weights which were approximately 5 percent of the weight of the wing. In reference 1 a maximum reduction of 13 percent is shown for weights which were approximately 60 percent of the weight of the wing. A comparison on the basis of weight alone with the results of reference 1 shows the reduction in flutter speed in the present cases to be of much larger magnitude than might be expected. That the effect is one of moment of inertia rather than one of mass is indicated by examination of figure 6, in which is shown the variation of natural frequencies with span position for weight 3. The bending frequencies appear to be relatively unchanged, indicating that the effect of mass is small; but in the torsional frequency there is noted a marked reduction, which can be attributed to the appreciable moment of inertia of the weight, for weight positions near the tip.\n\nThe flutter frequencies for weights 3 and 4 were not greatly affected by the variation in spanwise position of the weights (see figs. 5 and 6). As shown in figure 5, a maximum reduction of $17\\frac{1}{2}$ percent was found. In reference 1 the maximum reduction amounted to 59 percent for weights that were approximately 60 percent of the weight of the wing and had chordwise center-of-gravity positions comparable to those of weights 3 and 4.\n\nEffects of Moment of Inertia of Light Concentrated Weights\n\nThe effects of the moment of inertia of the weight on flutter speed and flutter frequency have been studied with the aid of tip weights 2c (fig. 3(e)) and 2d (fig. 3(f)), in addition to weight 2a. The results are presented in figures 7 and 8. As can be seen in these figures, an increase in moment of inertia produced a decrease both in flutter speed and flutter frequency. Comparison among the natural frequencies in figure 8 further shows that the main effect of", "timestamp": "2026-07-22T06:59:44.988825+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 25, "total_pages": 55, "image_filename": "19930085966_p25.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:59:47.331624+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 9, "total_pages": 20, "image_filename": "19930085999_p9.jpg", "text": "NACA RM E9I07\n\nThe results of a similar investigation of tightly mounted blades of the same design (reference 4) are shown in figure 6. Comparison between these results and those obtained with loosely mounted blades indicates that the vibration characteristics were nearly similar, both as to frequency and stress level.\n\nOn the basis of this comparison, any gain in damping effect produced by loose mounting appeared negligible. The vibration characteristics of the loosely mounted turbine blades differed from those of tightly mounted blades in one respect: the loosely mounted blades showed a tendency toward vibration of brief duration in first bending mode at many different turbine speeds in addition to those speeds coincident with the occurrence of a sustained vibration. The tightly mounted blades vibrated perceptibly only at definite turbine speeds coincident with the condition of sustained vibration associated with resonance.\n\nFurthermore, when vibration of appreciable magnitude occurred in the case of the loosely mounted blades, the vibration was sustained at a nearly constant value over a small but definite range of turbine speed. A condition of vibration in tightly mounted blades occurred only within a much more limited range of turbine speed. It was also observed that the loosely mounted blades would sometimes change momentarily from a complex-mode vibration to a first bending-mode vibration.\n\nSeveral factors must be considered in the analysis of the results. With the exception of the two experimental blades, all turbine blades were tightly mounted. This factor may have caused a relatively greater amount of compressive effect in the rim at the location of the loosely mounted blades than would have occurred in a turbine wheel completely fitted with loosely mounted blades. The differences in vibration characteristics observed for the two types of mounting, however, warrant the assumption that some degree of looseness existed in the experimental blade mounts during engine operation.\n\nAnother possible effect of the presence of tightly mounted turbine blades in the wheel is the transmission of vibration from a blade at resonance to other blades in a remote location, either through the base common to both, or in the manner exemplified by excitation of one tuning fork by another of the same natural frequency. Previous examination of a considerable number of turbine blades has shown, however, that natural frequency varies considerably among similar turbine blades. In addition, the method of dovetail insertion would tend to prevent any such vibration because of", "timestamp": "2026-07-22T06:59:48.770180+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 19, "total_pages": 46, "image_filename": "19930085983_p19.jpg", "text": "```markdown\nNACA RM A9I27\n\nCONFIDENTIAL\n\nLift coefficient, $C_L$\n\n$\\delta_u$, deg\n$\\circ$ 0\n$\\square$ -5\n$\\diamond$ -10\n$\\triangle$ -15\n$\\nabla$ -20\n$\\triangleright$ -25\n\n<!-- Image (103, 179, 915, 805) -->\n\nAngle of attack, $\\alpha$, deg\nfor $\\delta_u = 0^\\circ$\nPitching-moment coefficient, $C_m$\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 4.- The effect of elevon deflection on the aerodynamic characteristics of the wing-fuselage combination and on the elevon hinge-moment coefficients at a Mach number of 0.20.\n\nCONFIDENTIAL\n\n17\n```", "timestamp": "2026-07-22T07:00:00.044896+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 19, "total_pages": 132, "image_filename": "19930085990_p19.jpg", "text": "NACA RM A9I01 CONFIDENTIAL 17\n\nAPPENDIX\n\nThe following tables have been included to provide a convenient index to the figures presenting the results of this investigation:\n\nFORCE AND MOMENT CHARACTERISTICS\n\nWing Alone\n\n| Results presented | Flap deflection | Mach number | Reynolds number | Figure number |\n|-------------------|-----------------|-------------|-----------------|---------------|\n| $\\alpha$, $C_D$, & $C_m$ vs $C_L$ | $0^\\circ$ | 0.20 to 0.94 | $2 \\times 10^6$ | 3 |\n| $\\alpha$, $C_D$, & $C_m$ vs $C_L$ | $\\delta_n=30^\\circ, \\delta_f=50^\\circ$ | 0.20 | $3 \\times 10^6$ to $10 \\times 10^6$ | 4 |\n\nWing Alone With All Gaps Sealed\n\n| Results presented | Flap deflection | Mach number | Reynolds number | Figure number |\n|-------------------|-----------------|-------------|-----------------|---------------|\n| $C_L$ vs $\\alpha$ | — — — | 0.20 to 0.94 | $1 \\times 10^6$ | 5 |\n\nWing-Fuselage Combination\n\n| Results presented | Flap deflection | Mach number | Reynolds number | Figure number |\n|-------------------|-----------------|-------------|-----------------|---------------|\n| $C_L$ vs $\\alpha$ | $0^\\circ$ | 0.20 to 0.95 | $2 \\times 10^6$ | 6 |\n| $C_L$ vs $C_D$ | ↓ | ↓ | ↓ | 7 |\n| $C_L$ vs $C_m$ | ↓ | ↓ | ↓ | 8 |\n| $\\alpha$, $C_D$ & $C_m$ vs $C_L$ | $\\delta_n=30^\\circ, \\delta_f=50^\\circ$ | 0.20 | $2 \\times 10^6$ to $10 \\times 10^6$ | 9 |\n\nWing, Fuselage, and Horizontal Tail in Extended Wing-Chord Plane\n\n| Results presented | Flap deflection | Stabilizer angle | Mach number | Reynolds number | Figure number |\n|-------------------|-----------------|------------------|-------------|-----------------|---------------|\n| $C_L$ vs $\\alpha$ | $0^\\circ$ | $4^\\circ$ to $-10^\\circ$ | 0.20 to 0.95 | $2 \\times 10^6$ | 10(a) to 10(h) |\n| $C_L$ vs $C_D$ | ↓ | ↓ | ↓ | ↓ | 11(a) to 11(h) |\n| $C_L$ vs $C_m$ | ↓ | ↓ | ↓ | ↓ | 12(a) to 12(h) |\n| $C_L$ vs $\\alpha$ | $\\delta_n=30^\\circ, \\delta_f=50^\\circ$ | $4^\\circ$ to $-10^\\circ$ | 0.20 | $2 \\times 10^6$ to $10 \\times 10^6$ | 13(a) to 13(e) |\n| $C_L$ vs $C_D$ | ↓ | ↓ | ↓ | ↓ | 14(a) to 14(e) |\n| $C_L$ vs $C_m$ | ↓ | ↓ | ↓ | ↓ | 15(a) to 15(e) |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:00:00.234199+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 104, "total_pages": 114, "image_filename": "19930086061_p104.jpg", "text": "```markdown\n100\n\n<!-- Image (100, 48, 907, 833) -->\n\nFigure 49.- Effect of angle of attack on the local center-of-pressure location of wing 1; $\\psi = 0^\\circ$.\n\nNACA RM L9D07\n```", "timestamp": "2026-07-22T07:00:01.308352+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 57, "total_pages": 92, "image_filename": "19930085870_p57.jpg", "text": "58\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nElliptical L.E.: {O CL, □ Cm\nWedge L.E.: {△ CL, ◇ Cm\n.16\n.08\nCL\n0\n-.08\n-.16\n-.24\n.01\nCm\n0\n-.01\n\n.06\nElliptical L.E.: {O CD, □ L/D\nWedge L.E.: {△ CD, ◇ L/D\n.04\nCD\n.02\n0.8\n-6\n-4\n-2\n0\n2\n4\n6\n8\n0\n6\n4\nL/D\n2\n0\nα, deg\nNACA\n\n(f) Wing 6. w=1.252; R = 720,000.\nFigure 7. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:00:02.426345+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 36, "total_pages": 92, "image_filename": "19930085951_p36.jpg", "text": ".24\n.22\n.20\n.18\n.16\n.14\n.12\n.10\n.08\n.06\n.04\n.02\n0\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\nThrust coefficient, $C_T$\nAdvance ratio, J\n$\\beta_{0.75R}$ 20° 25° 30° 35° 40° 45°\nNACA\n(a) Thrust coefficient.\nFigure 11.— Characteristics of NACA 10-(3)(062)-045A propeller. Rotational speed, 1600 rpm.\nCONFIDENTIAL\nNACA RM L9D29\n34", "timestamp": "2026-07-22T07:00:03.286263+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 23, "total_pages": 25, "image_filename": "19930085979_p23.jpg", "text": "22\n\nRam-pressure recovery, $\\eta$\n1.00\n.96\n.92\n.88\n.84\n0 10 20 30 40 50 60 70 80 90\nModel-air-temperature rise, $^\\circ$F\n\n[Figure: Graph showing variation of ram-pressure recovery with model-air-temperature rise. Data points are plotted as open circles with a fitted line. NACA logo appears in lower right corner of plot area.]\n\nFigure 9. - Variation of ram-pressure recovery with model-air-temperature rise. Angle of attack, $0^\\circ$; plenum-chamber-gas temperature, $1000^\\circ$ F.\n\nNACA RM E5E12", "timestamp": "2026-07-22T07:00:05.755775+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 24, "total_pages": 30, "image_filename": "19930085975_p24.jpg", "text": "22\nNACA RM L9E10\n\nCONFIDENTIAL\n\n<!-- Image (109, 109, 686, 809) -->\n\nFigure 11.- The variation with Mach number of the wing-tip helix angle $\\frac{pb}{2V}$ for various aileron deflections at several angles of attack. $\\Lambda_{c/4} = 46.7^\\circ$.", "timestamp": "2026-07-22T07:00:09.252048+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 24, "total_pages": 30, "image_filename": "19930085970_p24.jpg", "text": "```markdown\nCONFIDENTIAL\n\nMean lift-curve slope, $dC_L/d\\alpha$, per deg\n\n| Configuration | Reynolds number | Source |\n| :--- | :--- | :--- |\n| — Wing and body | $.35-.52 \\times 10^6$ | 1- by $3\\frac{1}{2}$-ft wind tunnel |\n| — — Wing alone | $2.35 \\times 10^6$ | Reference 3 |\n| - - - - Wing alone | | Calculated |\n| $\\circ$ Wing and body | $.69 \\times 10^6$ | Reference 2 |\n| $\\square$ Wing and body | $8 \\times 10^6$ | Reference 4 |\n| $\\diamond$ Wing alone | $8 \\times 10^6$ | Reference 4 |\n\n[Figure: Graph plotting Mean lift-curve slope against Mach number, M. The x-axis ranges from 0 to 1.6. The y-axis ranges from 0 to .080. Various lines and points represent different configurations and sources as detailed in the legend.]\n\nMach number, M\n\nFigure 6.—Effect of Mach number on the mean lift-curve slope.\n\nNACA RM A52E09\n\nCONFIDENTIAL\n\n22\n```", "timestamp": "2026-07-22T07:00:15.516896+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 67, "total_pages": 149, "image_filename": "19930083192_p67.jpg", "text": "NACA TN 1976\n63\n\nconsidered are the use of an accelerometer or strain gage to measure\naccelerations or load and the use of a feeler vane or similar system to\ndetect gust-velocity or angle-of-attack changes. The servo-systems are\nassumed to operate the conventional controls of the airplane or to\noperate wing flaps or spoilers. The difficulties are those of possible\nfailures of complicated mechanisms, the question of oscillation of the\nsystem, the response of repeated gusts, and, finally, the need as air-\nplane speed increases of obtaining a servo-system that has a high fre-\nquency response (imposed frequencies up to 10 cps). The advantages\nare flexibility and utilization of many of the available mechanisms.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va., August 5, 1949", "timestamp": "2026-07-22T07:00:17.376761+00:00"}
{"citation_id": "19930086020", "source_url": "https://ntrs.nasa.gov/api/citations/19930086020/downloads/19930086020.pdf", "page_number": 4, "total_pages": 22, "image_filename": "19930086020_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM A9J06\n\nSince the pitching-moment data showed wide discrepancies when compared to data from tests at higher Reynolds numbers in the Ames 12-foot pressure wind tunnel, an attempt was made to isolate the cause of these discrepancies.\n\nSYMBOLS\n\n| | | |\n| :--- | :--- | :--- |\n| $C_L$ | lift coefficient | $\\left(\\frac{\\text{lift}}{qS'}\\right)$ |\n| $C_m$ | pitching-moment coefficient measured about 25-percent $\\bar{c}$ | $\\left(\\frac{\\text{pitching moment}}{qS'\\bar{c}}\\right)$ |\n| M | Mach number | $\\left(\\frac{V}{a}\\right)$ |\n| R | Reynolds number | $\\left(\\frac{\\rho V \\bar{c}}{\\mu}\\right)$ |\n| $S'$ | wing area of the semispan model, square feet | |\n| V | airspeed, feet per second | |\n| a | speed of sound, feet per second | |\n| b | wing span, perpendicular to plane of symmetry, feet | |\n| c | local chord, parallel to plane of symmetry, feet | |\n| $\\bar{c}$ | mean aerodynamic chord | $\\left(\\frac{\\int_0^{b/2} c^2 dy}{\\int_0^{b/2} c dy}\\right)$, feet |\n| q | dynamic pressure | $\\left(\\frac{1}{2}\\rho V^2\\right)$, pounds per square foot |\n| y | spanwise distance, feet | |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:00:18.208635+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 12, "total_pages": 32, "image_filename": "19930085992_p12.jpg", "text": "10\nNACA RM L9E17\n\nvariation in moment of inertia has been on the torsional degree of\nfreedom, with the bending frequencies remaining essentially unchanged.\nThis fact, together with the data observed in the variation in\nspanwise position of weights 3 and 4, indicates that even though a\nrelatively light concentrated weight is used, its moment of inertia\nmay be such that considerable influence is exerted on flutter speed\nand flutter frequency.\n\nCONCLUDING REMARKS\n\nIn a preliminary experimental program consisting of over\n20 flutter runs at low Mach numbers, results have been presented to\nshow some effects on flutter speed and flutter frequency of the\naerodynamic shape of concentrated weights. These parameters have been\ncompared for streamlined and nonstreamlined shapes of rigidly mounted\nweights that were varied over a wide range of spanwise positions on a\nstraight cantilever wing. In regard to shape, two general types of\nweights having similar mass and moment-of-inertia properties were\nemployed: one a streamlined body resembling in shape an external\nwing fuel tank and the other a chosen nonstreamlined body. Because of\nthe preliminary nature of this investigation, the following remarks\nare necessarily restricted to data on this wing and therefore cannot\nbe regarded as general.\n\nResults, concerning the main objective of the investigation, show\nthat both flutter speed and flutter frequency are relatively unaffected\nby radical changes in the aerodynamic shape of the concentrated\nweights.\n\nFurther observations in this investigation are possible on two\nother results which are considered to be logical, though perhaps\nincidental, outgrowths of the main objective. In regard to the first\nof these auxiliary results, the variation in spanwise position of\nrelatively light concentrated weights (approximately 5 percent of the\nweight of the wing) produces an effect on the flutter speed that is\nlarge when compared with the results of reference 1 for much heavier\nweights; the effect on flutter frequency, however, is small compared\nwith that found in reference 1 for heavier weights. In regard to the\nsecond of these other results, it is experimentally demonstrated that\nthe effect on flutter speed and flutter frequency of the moment of\ninertia of a relatively light concentrated weight may be large.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.", "timestamp": "2026-07-22T07:00:20.563235+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 15, "total_pages": 40, "image_filename": "19930085997_p15.jpg", "text": "NACA RM A9I29\n\nCONFIDENTIAL\n\nBoundary-layer scoop ④\nMain scoop entrance ①\nDuct survey position ②\n0.375\nPlane of symmetry\nSettling chamber ③\nBoundary-layer duct\nsurvey position ②\nVariable exit\nthroat\n1.250 dia.\n0.500 dia.\nTotal area of main scoops, $A_1=0.067$ sq in.\nTotal area of boundary-layer scoops,\n$A_4=0.081$ sq in.\n0.066\n32°30'\nRadial \"A\"\nReference plane\nStation\n0\n2.225\n2.881\n3.750\n3.875\n4.125\n4.412\n4.704\n4.900\n5.400\n5.800\n6.400\n6.900\n7.200\nPilot canopy\nReference plane\n6°\n2°\nConfiguration E\nConfiguration F\nAll dimensions are in inches.\nNACA\nPilot tubes\n(Stations 7500 & 9000)\nFigure 1. - Model dimensions.\nCONFIDENTIAL\n13", "timestamp": "2026-07-22T07:00:22.413301+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 26, "total_pages": 55, "image_filename": "19930085966_p26.jpg", "text": "NACA RM L9B17\n25\n\nCONFIDENTIAL\n\nDrilled rivet\nFlared section\n\nSection A-A\n\nA\nA\n\nNACA\n\nFigure 3.- Modified nonvaporizing pilot-burner housing.\nCONFIDENTIAL", "timestamp": "2026-07-22T07:00:22.583628+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 20, "total_pages": 46, "image_filename": "19930085983_p20.jpg", "text": "18\nCONFIDENTIAL\nNACA RM A9I27\n\n<!-- Image (167, 100, 769, 845) -->\n\n(b) $C_L$ vs $C_D$, $C_h$ vs $\\alpha$.\nFigure 4.- Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T07:00:27.129939+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 94, "total_pages": 98, "image_filename": "19930086073_p94.jpg", "text": "96\n\nRolling-moment coefficient, $C_l$\nYawing-moment coefficient, $C_n$\nSide-force coefficient, $C_Y$\nAngle of sideslip, $\\beta$, deg\n(a) Wing alone.\n\nRolling-moment coefficient, $C_l$\nYawing-moment coefficient, $C_n$\nSide-force coefficient, $C_Y$\nAngle of sideslip, $\\beta$, deg\n(b) Wing + body.\n\nRolling-moment coefficient, $C_l$\nYawing-moment coefficient, $C_n$\nSide-force coefficient, $C_Y$\nAngle of sideslip, $\\beta$, deg\n(c) Wing + body + vertical tail.\n\n$C_L$, 0.0\n0.4\n0.8\n1.1\n\nNACA\n\nFigure 24. — Variation of rolling-moment, yawing-moment, and side-force coefficients with sideslip at four lift coefficients.\n\nNACA RM A9E04", "timestamp": "2026-07-22T07:00:28.036825+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 39, "total_pages": 47, "image_filename": "19930085919_p39.jpg", "text": "38\nCONFIDENTIAL\nNACA RM No. 49C21\n\nLift coefficient, $C_L$\nAngle of attack, $\\alpha$, deg\nPitching-moment coefficient, $C_m$\n\n$\\delta_f$, deg $\\delta_e$, deg Drooped-nose flap\n$\\square$ -2.0 40\n$\\circ$ 0 0\n0.5$c$\nConstant-chord aileron\n\n$\\delta_f$, deg $\\delta_e$, deg Drooped-nose flap\n$\\diamond$ -2.0 40\n$\\circ$ 0 0\n0.5$c$\nConstant-chord aileron\n1.0$c$\n0.25$c$ split flap\n$\\delta_f$, 45°\n\nNACA\n\nFigure 13- Effect of the drooped-nose flap of 50-percent span on the lift and pitching-moment characteristics of the model with short fuselage. $R, 4.2 \\times 10^6$.", "timestamp": "2026-07-22T07:00:29.731518+00:00"}