PDF/ocr_dataset_1080.jsonl

{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 25, "total_pages": 92, "image_filename": "19930085951_p25.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n\nNACA RM L9D29\n\n(a) Blade sections at 0.3 radius.\nNACA 10-(3)(08)-03\nNACA 10-(3)(08)-03R\nNACA 10-(3)(12)-03\n\nNACA 10-(3)(08)-03\nNACA 10-(3)(08)-03R\nNACA 10-(3)(12)-03\n\n(b) Blade sections at 0.7 radius.\n\nFigure 5.— Comparison of blade sections at two radii for NACA propellers having a solidity\nof 0.03 per blade at the 0.7 radius.\n\nCONFIDENTIAL\nUNCLASSIFIED\n\n23", "timestamp": "2026-07-22T06:53:32.126999+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 28, "total_pages": 47, "image_filename": "19930085919_p28.jpg", "text": "NACA RM No. A9G21\n\nCONFIDENTIAL\n\n[Figure: Three wing configurations: Plain wing, Plain wing + long fuselage, Plain wing + short fuselage]\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\nAngle of attack, $\\alpha$, deg\n\nPitching-moment coefficient, $C_m$\n\nNACA\n\nFigure 7.—Lift, drag, and pitching-moment characteristics of the plain wing and of the wing-fuselage combinations. $R, 4.2 \\times 10^6$.\n\nCONFIDENTIAL\n\n27", "timestamp": "2026-07-22T06:53:33.280802+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 59, "total_pages": 60, "image_filename": "19930085862_p59.jpg", "text": "NACA RM No. L9A07\n57\n\n<!-- Image (163, 110, 856, 875) -->\n\nFigure 24.- Variation of rolling-moment coefficient with span of 0.10c projection step spoilers on wing equipped with drooped-nose flaps and trailing-edge split flaps.", "timestamp": "2026-07-22T06:53:33.674259+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 15, "total_pages": 55, "image_filename": "19930085966_p15.jpg", "text": "14 CONFIDENTIAL NACA RM L9B17\n\n$$\n\\eta_d = \\frac{\\Delta p_{1-2}}{(p_{t1} - p_1) - (p_{t2} - p_2)}\n$$\n\nA plot of these values against the air-bleed pressure coefficient gave the curve of figure 19 with a point scatter of ±1 percent-units. The points in figure 19 were from runs with no combustion using screen resistances. It is apparent that the diffuser efficiency is a function of the pressure and mass-flow conditions at the bleed and is not affected by combustion as such. The low-pressure-coefficient end of the curve, which drops below 99 percent efficiency, is for very lean mixtures (fuel-air ratios of 0.022 or less) so that most of the combustion data were taken with a diffuser efficiency of 99 percent. The high efficiency is due to complete removal of the boundary layer at the diffuser inlet, thus preventing the occurrence of boundary layers of sufficient thickness to separate within the diffuser length. It can be concluded that the 20-percent drop below the ideal efficiency must be charged almost entirely to combustion-chamber momentum, friction, and turbulence losses.\n\nThe combustion efficiency is a function of many variables including the type of burner, the combustion-chamber configuration, the pressure and temperature of the intake air, the fuel and air distribution, the fuel flow, the combustion-chamber inlet-air velocity, the fuel-air ratio, and the type of fuel. The fuel flow and air flow (and consequently the combustion-chamber inlet velocity and fuel-air ratio) were the principal variables in the test program. The relation of the measured combustion efficiency to these variables is illustrated by figure 20, which is a family of curves of constant combustion efficiency plotted on coordinates of fuel and air flows. Superimposed on the principal coordinates are curves of constant combustion-chamber inlet velocity, temperature rise through the combustion chamber, and fuel-air ratio. All the variables plotted in figure 20 are interrelated; however, it is possible to draw some general conclusions concerning the combustion efficiency. It is apparent from the efficiency curves that the lean-mixture tests were not extended to high enough combustion-chamber inlet velocities to obtain marked decreases in combustion efficiency due to approaching the blow-out condition. This effect would have caused a decreasing negative slope of the lower-value efficiency curves with increasing air flow at constant fuel flow.\n\nIt is also apparent that it was somewhat more efficient to burn a given quantity of fuel at high fuel-air ratios or low air flows. It is difficult to deduce the reason for this effect; it is possible that\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:37.538845+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 11, "total_pages": 25, "image_filename": "19930085979_p11.jpg", "text": "10\nNACA RM E9E12\n\nFor all icing conditions investigated, icing of the accessory housing occurred at bleedbacks less than approximately 6.0 percent for a plenum-chamber-gas temperature of $1000^\\circ$ F. The largest ice formation observed on the accessory housing was an ice cap 4 inches in diameter and $3\\frac{1}{2}$ inches thick with some ice extending back 7 inches along the skin. This ice accretion formed during a 40-minute icing period in which the bleedback was reduced from an initial value of 5.6 percent to a final value of 4.9 percent at a constant plenum-chamber-gas temperature of $1000^\\circ$ F. The icing condition was at a tunnel total temperature of $0^\\circ$ F and liquid-water content of 0.75 gram per cubic meter. No ice formations were observed on the duct walls for the conditions investigated.\n\nThe importance of accessory-housing ice protection depends largely on the particular jet engine to be used. If protection is essential, the accessory housing cover may be heated either by hot gases or electrically.\n\nSUMMARY OF RESULTS\n\nThe following results were obtained from an aerodynamic and icing investigation in the icing research tunnel of a two-thirds-scale model of a turbojet-engine nacelle with a short straight air inlet utilizing a hot-gas bleedback system for ice prevention:\n\n1. The optimum temperature distribution was obtained at a bleedback of 4.9 percent and was independent of tunnel-air velocity. This value of bleedback resulted in an average model-dry-air-temperature rise of $50^\\circ$ F with a maximum local temperature deviation of $10^\\circ$ F at the inlet screen for a plenum-chamber-gas temperature of $1000^\\circ$ F.\n\n2. The use of plenum-chamber-gas pressure other than optimum resulted in increased temperature gradients across the inlet. For pressures higher than optimum, low-temperature regions existed near the duct walls; for pressures lower than optimum, low-temperature regions occurred at accessory-housing skin.\n\n3. The introduction of cold gas under pressure through the orifices decreased the ram-pressure recovery linearly with increasing bleedback.\n\n4. Hot-gas bleedback decreased the ram-pressure recovery linearly with increasing model-air temperature rise.", "timestamp": "2026-07-22T06:53:39.914731+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 23, "total_pages": 34, "image_filename": "19930085913_p23.jpg", "text": "22\nNACA RM L9F24\n\n| | 1 | 2 | 3 | 4 | | 1 | 2 | 3 | 4 | | 1 | 2 | 3 | 4 |\n|---|---|---|---|---|---|---|---|---|---|---|---|---|---|---|\n| | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 |\n| 49 | | | | | 50 | | | | | 51 | | | | |\n| | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 |\n| 52 | | | | | 53 | | | | | 54 | | | | |\n| | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 |\n| 55 | | | | | 56 | | | | | 57 | | | | |\n| | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 |\n| 58 | | | | | 59 | | | | | 60 | | | | |\n| | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 | | 50 | 50 | 20 | 20 |\n| 61 | | | | | 62 | | | | | 63 | | | | |\n\n[NACA]\n\n(d) Model B; $\\Lambda = 45^\\circ$, $e_w = 0$.\nFigure 1.— Continued.", "timestamp": "2026-07-22T06:53:42.182301+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 94, "total_pages": 114, "image_filename": "19930086061_p94.jpg", "text": "90\nNACA RM L9J07\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n\n$$ \\frac{y}{b/2} \\text{, percent} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\circ$ 4.1 | 0.14 |\n| $\\square$ 8.1 | 0.29 |\n\n(a) Angles of attack: 4.1°, 8.1°\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n\n$$ \\frac{y}{b/2} \\text{, percent} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\diamond$ 14.1 | 0.42 |\n| $\\triangle$ 24.1 | 0.59 |\n| $\\triangledown$ 34.1 | 0.75 |\n\n(b) Angles of attack: 14.1°, 24.1°, 34.1°\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n\n$$ \\frac{y}{b/2} \\text{, percent} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\square$ 39.1 | 0.70 |\n| $\\circ$ 44.1 | 0.58 |\n\n(c) Angles of attack: 39.1°, 44.1°\n\nFigure 39.- Span load distribution of wing 1 at various angles of attack; $\\psi = 20^\\circ$. Flagged symbols represent data taken with left semispan at $\\psi = -20^\\circ$.", "timestamp": "2026-07-22T06:53:44.127816+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 57, "total_pages": 149, "image_filename": "19930083192_p57.jpg", "text": "NACA TN 1976\n53\n\nregulations, or the introduction of newer airplanes, then the prediction may be satisfactory. Such changes as, for example, the shift from prewar to postwar operations may lead to some uncertainty as to the applicability of a prediction. Predictions assume that the trend in the data has not been affected by new factors. (No method yet is available for evaluating or detecting the effect of changes in the data now used.)\n\nOther questions that occur in statistical analyses arise in connection with the existence or inclusion of data from two different regions or populations as part of the same sample, the presence of absolute and physical boundaries that cannot be exceeded, the effect of extraneous variables, and the effect of gradual transitions from one region to another. The presence of such limitations can have serious results as far as the significance of the results is concerned and such effects cannot always be evaluated or separated. An example of the effect of different regions might be the combination of acceleration data involving flight below maximum lift coefficient and flight above maximum lift coefficient. The aerodynamic characteristics of the airplane and its behavior in the two regions are entirely different. Thus, a sample that combines such data does not offer an adequate means of predicting future expectations of a given acceleration except for the conditions of the original data, since any derived curves are a composite of two independent sets and therefore the prediction can only apply to the same combination.\n\nRESULTS\n\nV-G data.- From the individual V-G records that are summarized in table XIX, frequency distributions of the maximum acceleration increment $\\Delta n_{\\text{max}}$, the maximum speed $V_{\\text{max}}$, and the speed $V_o$ at which maximum acceleration increment was experienced have been compiled for the wartime and prewar periods. The statistical parameters: the mean, the standard deviation, and the skewness are given in table XX, which also includes the values of $V_{o_{\\text{max}}}$, $\\Delta n_{\\text{max}}$, and $V_{\\text{max}}$ recorded during the period under consideration.\n\nThe parameters of the Pearson type III curves given in table XX have been used to obtain the flight miles to exceed the specified values of a selected quantity. The transformation from probability to flight miles was performed through use of a nominal cruising speed and the average flight hours per V-G record. Table XXI gives the flight miles to exceed: (1) the limit load-factor increment, (2) the placard speed of the airplane, and (3) an acceleration increment corresponding to encountering a gust with an effective gust velocity of 37.5K feet per second at the most probable value of $V_o$, $V_p$. The probable speed $V_p$ has been included in table XXI as a fraction of the high speed in level flight $V_L$.", "timestamp": "2026-07-22T06:53:46.235816+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 37, "total_pages": 52, "image_filename": "19930085911_p37.jpg", "text": "36\nCONFIDENTIAL\nNACA RM E9F22\n\n<!-- Image (116, 135, 836, 821) -->\n\n(d) Combustion-chamber-inlet variables.\n\nFigure 8. - Continued. Time history of flight data and performance of ram-jet unit 16-A-3.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:47.677468+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 6, "total_pages": 32, "image_filename": "19930085992_p6.jpg", "text": "4\nNACA RM L9E17\n\n$f_{h_2}$ second bending natural frequency, cycles per second\n$f_t$ first torsion natural frequency, cycles per second\n$f_e$ experimental flutter frequency, cycles per second\n$V_i$ indicated airspeed at flutter, feet per second\n$V$ true airspeed at flutter, feet per second\n$q$ dynamic pressure at flutter, pounds per square foot\n$g_{h_1}$ structural damping coefficient, first bending\n$g_{h_2}$ structural damping coefficient, second bending\n$g_\\alpha$ structural damping coefficient, first torsion\n\nSubscript:\nw refers to the corresponding properties or parameters\nof wing carrying concentrated weights\n\nAPPARATUS\n\nThe entire series of approximately 20 flutter tests was made on\na single uniform wing. The model selected for testing, built of\nmagnesium alloy, was 40 inches long with an 8-inch chord and had an\nNACA 16-004 airfoil section. As shown in figure 1, the model was\nmounted rigidly to the top of the test section as a cantilever beam\nso that the flutter produced may be considered to correspond to a\nsymmetrical mode. The chordwise slots shown along the trailing edge\nin figure 1 were cut to a depth of approximately $2\\frac{3}{4}$ inches at every\ninch along the span in an effort to move the elastic axis forward\nand hence keep the divergence speed above the expected flutter speed", "timestamp": "2026-07-22T06:53:48.132886+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 13, "total_pages": 30, "image_filename": "19930085975_p13.jpg", "text": "CONFIDENTIAL\n\nTunnel ceiling\n\nBalance strut\n\nStrut fairing\n\nSting fairing\n\nShaft locking screws\n\nTunnel center line\n\nSting support\n\nTest wing\n\nAngle-of-attack changing block\n\nRoller bearing\n\nBall thrust bearing\n\nTunnel floor\n\nCONFIDENTIAL\n\nNACA\n\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-tunnel test section.\n\nNACA RM L9E10\n\n11", "timestamp": "2026-07-22T06:53:50.598178+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 64, "total_pages": 72, "image_filename": "19930085491_p64.jpg", "text": "Laminar flow area\nTurbulent flow area\nSeparated flow area\nFuselage area\nDirection of flow\n\n(a) $C_L=0; R=0.31$ million\n(b) $C_L=0; R=0.62$ million\n\n(c) $C_L=0.21; R=0.62$ million\n(d) $C_L=0.28; R=0.62$ million\n\nNACA\n\nFigure 13.- Boundary-layer flow patterns on 63° swept-back wing.\n\nCONFIDENTIAL\nNACA RM No. A8J04\nCONFIDENTIAL\n63", "timestamp": "2026-07-22T06:53:52.059858+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 6, "total_pages": 17, "image_filename": "19930085988_p6.jpg", "text": "NACA RM L9H30 CONFIDENTIAL 5\n\nFreely-falling-body tests of a configuration having $45^\\circ$ sweptback wings located at two longitudinal stations on the body of reference 1 (from which the present body shape was derived) were reported in reference 3. The reference wings were nontapered, of aspect ratio 4.1, and had NACA 65-series sections of 6.36-percent thickness ratio in the free-stream direction. The wing-plus-interference drag coefficients have been determined from the total-drag-coefficient curves of references 1 and 3 and are compared with the present test results in figure 7. The station of the 0.5-root-chord point of the wings relative to the station of maximum diameter was 1.5 diameters forward and rearward for the reference tests and 0.6 diameter forward for the present tests.\n\nThe comparison indicates that the wing-plus-interference drag of the present configuration might be significantly reduced by a rearward shifting of the wing. Evidence of an unfavorable interference effect is indicated below M = 1 by the base-drag-coefficient curves in figure 5 wherein the addition of the wing and removal of two fins increased the base-drag coefficient by 0.002 at M = 0.95.\n\nBase-pressure coefficients for the body-alone and wing-body configurations are shown over the Mach number range in figure 6. The differences in configuration between the two test models had little effect on the results above a Mach number of 1. Below M = 1 however there appears to be a marked quantitative difference due to a configuration change although the qualitative agreement remains good.\n\nTests of a similar body at low Reynolds numbers but with artificial transition at the nose were reported in reference 4. The base-pressure coefficient was indicated to be -0.035 at M = 1.5 which compares favorably with the present results.\n\nTotal-drag coefficient, referred to body frontal area, against Mach number is given in figure 8 for the wingless configuration. For comparison, the results of reference 1 are included. When proper allowance is made for the effect of the fins and of the differences in body shape near the tail, reasonable agreement is indicated at supersonic speeds.\n\nCONCLUSIONS\n\nThe zero-lift drag of a transonic research model with and without tapered wings sweptback $45^\\circ$ has been measured at supersonic, transonic, and high subsonic speeds and at high Reynolds numbers in flight tests of rocket-powered models. Within the limit of the investigation the results indicated the following:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:55.603503+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 7, "total_pages": 40, "image_filename": "19930085997_p7.jpg", "text": "NACA RM A9I29 CONFIDENTIAL 5\n\nimpossible to simulate full-scale ducting in the suction system with a model of this scale, more precise measurements of these quantities were not made. These data are probably most valuable for their qualitative significance.\n\nMeasurements of the total pressure in the settling chamber were made at three equally spaced circumferential locations at the position indicated in figure 1. The area ratio between the total scoop-entrance area and the cross-sectional area of the settling chamber was 0.07 which, for isentropic diffusion from sonic velocity, would correspond to a settling-chamber Mach number of about 0.04. With this degree of diffusion, it was considered unnecessary to attempt a further survey of the total pressure other than that afforded by the three pitot tubes. This assumption was substantiated by the fact that the difference in total pressure measured by each of the three tubes was within 2 percent of the average of those tubes at every rate of mass flow.\n\nIn order to determine the effect of the subsonic diffusers upon the values of total pressure measured in the settling chamber of the model, the total-pressure distribution at position 2 was obtained. The average computed Mach number at this position was approximately 0.50 at mass-flow ratios at which the normal shock wave was ahead of the inlets. Measurements of the total pressure were made in both ducts with the model at an angle of attack of $0^\\circ$. The measurement locations are shown in figure 4. As indicated in the figure, each location was numbered and the duct cross section at this position was divided to obtain a weighted average of the total-pressure measurements. Properly weighted values of $H_2/H_0$ would have been based upon the mass flow through the area divisions shown in figure 4. Since measurements of these mass flows were impossible because of the model scale, the weighted averages were based upon the areas.\n\nRESULTS AND DISCUSSION\n\nThe results will be discussed in two parts. In the first section, test results at an angle of attack of $0^\\circ$ are discussed. In this section, model modifications designed to improve the pressure recovery throughout the Mach number range are investigated, and the effects of variations in the parameters $m_1/m_0$ and $m_4/m_0$ upon the total-pressure recovery of the best configuration are discussed. Also, in this section, the performance of the best model is analyzed with respect to the total-pressure distribution within the ducts, interaction between duct systems, and estimated energy expended in boundary-layer removal. Finally, in the second section, the effects of variations in angle of attack are treated, and modifications designed to improve the pressure recovery at angle of attack are discussed.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:56.377196+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 10, "total_pages": 132, "image_filename": "19930085990_p10.jpg", "text": "8\nCONFIDENTIAL\nNACA RM A9I01\n\nAn index of the figures presenting the results of this investigation\nis given in the appendix.\n\nRESULTS AND DISCUSSION\n\nForce and Moment Characteristics\n\nThe lift, drag, and pitching-moment characteristics of the model\nand its components are presented in figures 3 through 26.\n\nWing alone.- The effects of Reynolds number and of Mach number on\nthe lift, drag, and pitching-moment characteristics of the wing have been\nreported in reference 1. Data from that reference for a Reynolds number\nof 2,000,000 at Mach numbers from 0.20 to 0.94 are reproduced herein in\nfigure 3. The data of this figure indicate no large or erratic effects\nof compressibility up to a Mach number of 0.94. The wing lift-curve\nslope was 0.062 at a Mach number of 0.20 and increased to 0.095 at a Mach\nnumber of 0.94. The total movement of the aerodynamic center at zero\nlift was only about 7 percent of the wing mean aerodynamic chord over the\ntest Mach number range.\n\nThe force and moment characteristics of the wing with various\ncombinations of leading-edge and trailing-edge flap deflections have been\nreported in reference 2. The data of this reference indicate that a\nleading-edge flap deflection of 30° and a trailing-edge flap deflection\nof 50° were optimum for maximum lift. Data obtained with this combination\nof flap deflections are presented herein in figure 4 for a Mach number of\n0.20 and Reynolds numbers from 3,000,000 to 10,000,000. These data show\nthat deflection of the flaps increased the maximum lift of the wing from\n0.76 to 1.40 and that the aerodynamic characteristics of the wing with\nthe flaps deflected were little affected by increase of Reynolds number to\n10,000,000.\n\nThe variation with angle of attack of the lift coefficient of the\nwing with the gaps sealed and faired is presented in figure 5 for a Reynolds\nnumber of 1,000,000 for Mach numbers up to 0.94. Since the wing and tail\nwere geometrically similar and the mean aerodynamic chord of the tail was\none-half that of the wing, these data may be considered to represent the\nlift characteristics of the isolated tail and may be applied as the character-\nistics of the tail on the model at a Reynolds number of 2,000,000, based\non the wing mean aerodynamic chord, if corrections are made for the down-\nwash and reduction in the dynamic pressure at the tail.\n\nWing-fuselage combination.- The force and moment characteristics of\nthe wing-fuselage combination with the flaps neutral are shown in figures\n6, 7, and 8. Comparison of these data with those of figure 3 reveals that\naddition of the fuselage caused an increase in the drag, a reduction in\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:53:58.948483+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 6, "total_pages": 24, "image_filename": "19930085991_p6.jpg", "text": "```markdown\n4\nNACA RM L9I28\n\nparallel with and perpendicular to the air stream. The behavior of the\nmodel in the air stream was observed after each launching, after which\nthe air-stream velocity was lowered and the model caught in a safety\nnet and retrieved for the next launching. A photograph showing the\ntest section of the free-spinning tunnel with an airplane model spinning\nin the tunnel is shown in figure 2.\n\nVarious combinations of fin installations and center-of-gravity\nlocations were investigated to determine arrangements which would make\neach model descend in a stable nose-down manner. It is recognized that\nthe use of fins on an airplane jettisonable nose section will require\nthat the airplane either be constructed so that satisfactory flight\ncharacteristics can be obtained with the fins installed or that the fins\nbe initially retracted and be extended immediately as the nose separates\nfrom the rest of the airplane. The type of fin arrangement found during\nthe tests to be most effective in stabilizing the models, and hence used\nin the present study, consisted of four or three fins placed on the side\nof the nose section, generally at $90^\\circ$ or $120^\\circ$ intervals, respectively,\naround the periphery of the nose section at the break-off station.\nSketches of the various types of fin arrangements tested are shown in\nfigure 3. Arrangements d and e in figure 3 illustrate methods of\nmounting fins on a protuberance. Arrangements g, h, and i in\nfigure 3 simulate curved retractable fins. A fairly wide range of fin\naspect ratios, 0.4 to 2.0 (based on the span and area of each fin), was\ncovered during the tests. For all the tests in which curved fins were\ninstalled on the models, four fins were used with two being curved in\neach direction in order to avoid unbalanced rolling moments such as\nmight occur if the number of fins curved in each direction were not\nequal. In order to obtain a direct comparison of the relative\nstabilizing effectiveness of curved and flat fins, some of the tests\nwith curved fins were made with the fins installed at $90^\\circ$ intervals on\nthe nose periphery (arrangement g in fig. 3) in such a manner that they\nhad the same profile shape and projected area in a radial plane as a\ncorresponding flat-fin arrangement (arrangement c in fig. 3). The span\nof all the curved fins tested was small enough so that they would not\noverlap when retracted against the fuselage.\n\nA factor indicating the relative effectiveness of a given fin design\nwas determined for each condition tested. This factor, hereinafter\ncalled the fin-stabilization factor $\\left(\\frac{S_p L_T}{S_p L}\\right)$, was determined by the method\nillustrated in figure 1 and is the ratio of the smallest projected fin\narea in any plane parallel to the longitudinal axis multiplied by the\nprojected distance between the centroid of this area and the model\ncenter of gravity to the projected area of the model (excluding\nprotuberances) in the plane of smallest projected fin area multiplied\n```", "timestamp": "2026-07-22T06:53:59.541425+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 95, "total_pages": 114, "image_filename": "19930086061_p95.jpg", "text": "NACA RM L9J07\n91\n\n$$ \\frac{c_l c}{C_L c_{av}} $$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| $\\circ$ 4.1 | 0.11 |\n| $\\square$ 8.1 | 0.21 |\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| $\\diamond$ 14.1 | 0.33 |\n| $\\triangle$ 24.1 | 0.45 |\n| $\\triangleright$ 34.1 | 0.51 |\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| $\\square$ 39.1 | 0.54 |\n| $\\diamond$ 44.1 | 0.53 |\n\n(c) Angles of attack: 39.1°, 44.1°.\n\nFigure 40.- Span load distribution of wing 1 at various angles of attack; $\\psi = 35^\\circ$. Flagged symbols represent data taken with left semispan at $\\psi = -35^\\circ$.", "timestamp": "2026-07-22T06:54:05.803553+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 14, "total_pages": 30, "image_filename": "19930085975_p14.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:54:08.162084+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 26, "total_pages": 92, "image_filename": "19930085951_p26.jpg", "text": "```markdown\n24\nNACA RM L9D29\n\nUNCLASSIFIED\nCONFIDENTIAL\n\nDeveloped plan form\n\nLeading edge\n\nNACA 10-(3)(05)-045 propeller\nNACA 10-(3)(062)-045 A propeller\nNACA 10-(3)(062)-045 propeller\nNACA 10-(3)(08)-045 propeller\n\nBlade thickness ratio, h/b\nBlade width ratio, b/D\nBlade angle, $\\beta$, deg\nBlade-section design lift coefficient, $C_{ld}$\n\nFraction of tip radius, r/R\n\n[Figure: Blade-form curves for NACA propellers having a solidity of 0.045 per blade at the 0.7 radius.]\n\nFigure 6.- Blade-form curves for NACA propellers having a solidity of 0.045 per blade at the 0.7 radius.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:54:11.692302+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 16, "total_pages": 55, "image_filename": "19930085966_p16.jpg", "text": "NACA RM L9B17 CONFIDENTIAL 15\n\nthe hotter pilot flame at the lower air flows improved the evaporation and ignition of the main boiler fuel, and it is also possible that the lower air flows allowed the main boiler fuel jets to penetrate further into the air stream producing a more homogeneous mixture. Burning a given amount of fuel at low air flows also means burning at high combustion temperatures, as is indicated in figure 20. It is possible that this process was more efficient because of the beneficial effect of high temperature on combustion.\n\nThe variable which affected the efficiency to the greatest degree is shown by figure 20 to be the fuel flow, especially at the highest air flows. There may have been some loss in efficiency with increasing fuel flows due to exceeding the evaporative capacity of the boilers although this effect should have been minimized because the pilot fuel flow was increased proportionally with the main-boiler fuel flow. There undoubtedly is a change in fuel, fuel-air, and air distributions with increasing fuel flow at constant air flow. A locally enriched burning mixture should create an increased local resistance to air flow thus enriching the mixture further until a static pressure equilibrium with the surrounding air stream is reached. If the local region is at stoichiometric mixture before the enrichment takes place, the excess fuel may never burn with air from the surrounding regions and thus the over-all combustion efficiency drops. It is believed that such phenomena took place in the region of each of the five burners, causing five retarded regions surrounded by regions of high mass flow rates. At station 5 thermocouple and pressure measurements taken in preliminary tests indicated that the combustion had spread between the five regions forming a central core of hot gases surrounded by a relatively cool annulus of high mass flow rate adjacent to the wall. These phenomena were substantiated by visual observation.\n\nBecause the thin-plate-burner configuration has the characteristic of decreasing combustion efficiency with increasing fuel flow, it is quite evident that operation under high-thrust-output conditions, for instance a fuel-air of 0.06 and a combustion-chamber inlet velocity of 150 feet per second, is not feasible.\n\nIt is possible that longer combustion chambers might have provided better mixing and more complete combustion. This is substantiated in figure 21 by the slope of the curves of static pressure along the combustion chamber. The curves indicate that for the higher fuel flows combustion was still proceeding at the end of the combustion chamber, whereas for the lowest fuel flow shown the slope of the curve near the end of the combustion chamber is of the same order of magnitude as that which would result from the friction pressure drop alone.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:54:12.107077+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 65, "total_pages": 72, "image_filename": "19930085491_p65.jpg", "text": "64\n\nCONFIDENTIAL\n\nWF-57\n$C_L=0$\n\nWF-60\n$C_L=0$\n\nWF-63\n$C_L=0$\n\nWF-67\n$C_L=0$\n\nWF-70\n$C_L=0$\n\n(a) Non-lifting wings\n\nLaminar flow area\nTurbulent flow area\nSeparated flow area\nFuselage area\nDirection of flow\n\nWF-57\n$C_L=.27$\n\nWF-60\n$C_L=.26$\n\nWF-63\n$C_L=.21$\n\nWF-67\n$C_L=.22$\n\nWF-70\n$C_L=.21$\n\n(b) Lifting wings\n\nNACA\n\nFigure 14.- Effect of sweep on boundary-layer flow at a Reynolds number of 0.62 million.\n\nCONFIDENTIAL\n\nNACA RM NO. A8J04", "timestamp": "2026-07-22T06:54:13.985638+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 13, "total_pages": 46, "image_filename": "19930085983_p13.jpg", "text": "NACA RM A9I27 CONFIDENTIAL 11\n\n4. Madden, Robert T.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back 63°.- Characteristics at a Mach Number of 1.53 Including Effect of Small Variations of Sweep. NACA RM A8J04, 1949.\n\n5. Madden, Robert T.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back 63°.- Investigation at a Mach Number of 1.53 to Determine the Effects of Cambering and Twisting the Wing for Uniform Load at a Lift Coefficient of 0.25. NACA RM A9C07, 1949.\n\n6. Jones, J. Lloyd, and Demele, Fred A.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back 63°.- Characteristics Throughout the Subsonic Speed Range with the Wing Cambered and Twisted for a Uniform Load at a Lift Coefficient of 0.25. NACA RM A9D25, 1949.\n\n7. Jones, Robert T.: Estimated Lift-Drag Ratios at Supersonic Speed. NACA TN 1350, 1947.\n\n8. Glauert, H.: Wind Tunnel Interference on Wings, Bodies and Aircrews. R. & M. No. 1566, British A. R. C., 1933.\n\n9. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnels with Consideration of the Effect of Compressibility. NACA RM A7B28, 1947.\n\n10. Swanson, Robert S., and Priddy, E. LaVerne: Lifting-Surface-Theory Values of the Damping in Roll and of the Parameter Used in Estimating Aileron Stick Forces. NACA ARR L5F23, 1945.\n\n11. Polhamus, Edward C.: A Simple Method of Estimating the Subsonic Lift and Damping in Roll of Sweptback Wings. NACA TN 1862, 1949.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:54:15.621827+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 16, "total_pages": 30, "image_filename": "19930085970_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:54:15.868561+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 60, "total_pages": 60, "image_filename": "19930085862_p60.jpg", "text": "58\nNACA RM No. L9A07\n\n<!-- Image (89, 175, 862, 796) -->\n\n(a) Plain wing.\n\n(b) Leading-edge and split flaps.\n\nFigure 25.- Comparison of aileron and spoiler effectiveness.", "timestamp": "2026-07-22T06:54:19.244131+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 12, "total_pages": 25, "image_filename": "19930085979_p12.jpg", "text": "NACA RM E9E12\n\n5. For a constant tunnel-air velocity, the mass flow through the model decreased linearly with increasing model-air-temperature rise.\n\n6. Satisfactory agreement between the measured and calculated heat requirements for icing conditions was obtained.\n\n7. Icing protection for the accessory housing could be obtained at bleedbacks greater than 6.0 percent for a plenum-chamber-gas temperature of $1000^\\circ$ F.\n\nLewis Flight Propulsion Laboratory,  \nNational Advisory Committee for Aeronautics,  \nCleveland, Ohio.\n\nREFERENCES\n\n1. Callaghan, Edmund E., Ruggeri, Robert S., and Krebs, Richard P.: Experimental Investigation of Hot-Gas Bleedback for Ice Protection of Turbojet Engines. I - Nacelle with Offset Air Inlet. NACA RM E8D13, 1948.\n\n2. Callaghan, Edmund E., and Ruggeri, Robert S.: Experimental Investigation of Hot-Gas Bleedback for Ice Protection of Turbojet Engines. II - Nacelle with Long Straight Air Inlet. NACA RM E9C16, 1949.", "timestamp": "2026-07-22T06:54:21.867108+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 27, "total_pages": 92, "image_filename": "19930085951_p27.jpg", "text": "UNCLASSIFIED\nCONFIDENTIAL\n\nNACA RM L9D29\n\n(a) Blade sections at 0.3 radius.\nNACA 10-(3)(05)-045\nNACA 10-(3)(062)-045 A\nNACA 10-(3)(08)-045\nNACA 10-(3)(062)-045\n\nNACA 10-(3)(05)-045\nNACA 10-(3)(062)-045 A\nNACA 10-(3)(062)-045\nNACA 10-(3)(08)-045\n\n(b) Blade sections at 0.7 radius.\nNACA\n\nFigure 7.- Comparison of blade sections at two radii for NACA propellers having a solidity of 0.045 per blade at the 0.7 radius.\n\nUNCLASSIFIED\nCONFIDENTIAL\n\nCR", "timestamp": "2026-07-22T06:54:29.489934+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 7, "total_pages": 17, "image_filename": "19930085988_p7.jpg", "text": "6\nCONFIDENTIAL\nNACA RM L9H30\n\n1. The drag coefficient at supersonic speeds was approximately 0.015 for the body and 0.027 for the body-plus-wing configuration.\n\n2. The drag coefficient at subsonic speeds was approximately 0.008 for the body and 0.013 for the body-plus-wing configuration.\n\n3. The force-break Mach number was 0.98 for the body and 0.95 for the body-plus-wing configuration.\n\n4. The base contributed very little to the total drag of the test models, but the base-pressure data indicated a possible interference effect in that the addition of the wing and removal of two stabilizing fins increased the base-drag coefficient by 0.002 at a Mach number of 0.95.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nREFERENCES\n\n1. Thompson, Jim Rogers, and Mathews, Charles W.: Total Drag of a Body of Fineness Ratio 12 and Its Stabilizing Tail Surfaces Measured during Free Fall at Transonic Speeds. NACA CB L6D08, 1946.\n\n2. Alexander, Sidney R., and Nelson, Robert L.: Flight Tests to Determine the Effect of Taper on the Zero-Lift Drag of Wings at Low Supersonic Speeds. NACA RM L7E26, 1947.\n\n3. Mathews, Charles W., and Thompson, Jim Rogers: Comparison of the Transonic Drag Characteristics of Two Wing-Body Combinations Differing Only in the Location of the $45^\\circ$ Sweptback Wing. NACA RM L7I01, 1947.\n\n4. Chapman, Dean R., and Perkins, Edward W.: Experimental Investigation of the Effects of Viscosity on the Drag of Bodies of Revolution at a Mach Number of 1.5. NACA RM A7A31a, 1947.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:54:29.938123+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 96, "total_pages": 114, "image_filename": "19930086061_p96.jpg", "text": "92\nNACA RM L9J07\n\n$$\n\\frac{c_l c}{C_{L}C_{av}}\n$$\nTheoretical load distribution\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 4.1 | 0.14 |\n| 8.1 | 0.32 |\n\n$$\n\\frac{y}{b/2} \\text{ , percent}\n$$\n(a) Angles of attack: 4.1°, 8.1°.\n\n$$\n\\frac{c_l c}{C_{L}C_{av}}\n$$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 14.1 | 0.54 |\n| 24.1 | 0.85 |\n| 32.1 | 1.06 |\n\n$$\n\\frac{y}{b/2} \\text{ , percent}\n$$\n(b) Angles of attack: 14.1°, 24.1°, 32.1°.\n\n$$\n\\frac{c_l c}{C_{L}C_{av}}\n$$\n\n| $\\alpha$, deg | $C_L$ |\n| :--- | :--- |\n| 36.1 | 1.16 |\n| 44.1 | 0.75 |\n\n$$\n\\frac{y}{b/2} \\text{ , percent}\n$$\n(c) Angles of attack: 36.1°, 44.1°.\n\nFigure 41.- Span load distribution of wing 2 at various angles of attack; $\\psi = 0^\\circ$. All data are taken over left semispan.", "timestamp": "2026-07-22T06:54:30.884141+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 66, "total_pages": 72, "image_filename": "19930085491_p66.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 65\n\n[Figure: (a) Wind off.]\n\n[Figure: (b) Wind on. NACA A-12340]\n\nFigure 15.— Schlieren patterns common to all photographs.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:54:31.653943+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 7, "total_pages": 24, "image_filename": "19930085991_p7.jpg", "text": "NACA RM L9I28\n\nby the length of the model. For the fin arrangements in which four flat fins were installed on the nose-section periphery at $90^\\circ$ intervals and for the curved-fin arrangement in which the fins had the same location, profile shape, and projected area in a radial plane as the flat-fin arrangement, the plane of the smallest projected fin area was a plane which made a $45^\\circ$ angle with a plane through either pair of opposite fins. (See fig. 1.) For the remaining flat- and curved-fin arrangements, the plane of the smallest projected fin area was determined graphically for each condition tested. When a fin was mounted on a protuberance (arrangements d and e in fig. 3), it was arbitrarily considered to have the same projected area forward of the break-off station as did the fins at the other periphery intervals. The fin-stabilization factor was plotted against the center-of-gravity location for each condition tested, with different symbols being used to indicate whether or not the model descended in a stable nose-down attitude.\n\nRESULTS AND DISCUSSION\n\nA brief résumé of the results of Langley 20-foot free-spinning-tunnel tests of models simulating possible airplane jettisonable nose sections without stabilizing fins is included in table I. As shown in the table, some of the models descended with tumbling motions about their lateral or normal axis; whereas others trimmed at a high angle of attack and rolled about their longitudinal axis. The latter condition was obtained only with those models which had a canopy portion or other protuberance. The protuberances apparently excited a rolling moment which developed into an equilibrium rotation. When a suitable arrangement of stabilizing fins and center-of-gravity location was used, the models descended in a stable nose-down attitude without rolling. The results indicated that curved and flat fins having projected areas of the same magnitude and direction were equally effective in stabilizing a nose section.\n\nThe plot of fin-stabilization factor against the center-of-gravity location was examined and it was seen that for all the results except those for model 2, a boundary could be drawn which fairly well separated the regions for which stable nose-down descent was and was not obtained. It was noted that all the other models differed from model 2 primarily in that their fuselages extended forward almost to a point at their front end; for model 2, then, the fin-stabilization factors and center-of-gravity locations were recalculated by using an assumed altered body shape in which the model's profile lines extended forward until they too intersected at a point. The recalculated fin-stabilization factors were plotted and the conditions for stable", "timestamp": "2026-07-22T06:54:31.976695+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 24, "total_pages": 34, "image_filename": "19930085913_p24.jpg", "text": "NACA RM L9F24\n23\n\n| | | |\n| :--- | :--- | :--- |\n| 1 2 3 4 | 1 2 3 4 | 1 2 3 4 |\n| 50 50 20 20 | 50 50 20 20 | 50 50 20 20 |\n| 64 | 65 | 66 |\n| 50 50 30 30 | 50 50 20 20 | 30 30 20 20 |\n| 67 | 68 | 69 |\n| 30 30 20 20 | 30 50 20 20 | 50 50 20 20 |\n| 70 | 71 | 72 |\n| 50 50 30 30 | 30 30 30 30 | 30 30 30 30 |\n| 73 | 74 | Record taken at<br>limiting tunnel<br>velocity<br>75 Not flutter |\n| 30 30 30 30 | 30 50 30 30 | 30 50 30 30 |\n| 76 | 77 | 78 |\n| 30 50 20 20 | 20 30 20 20 | 30 30 20 20 |\n| 79 | 80 | 81 |\n\nNACA\n\n(e) Model C; $\\Lambda = 60^\\circ$; $e_w = -1$.\n\nFigure 1.— Continued.", "timestamp": "2026-07-22T06:54:32.248875+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 29, "total_pages": 47, "image_filename": "19930085919_p29.jpg", "text": "```markdown\n28\n\nCONFIDENTIAL\n\nLift coefficient, $C_L$\n\nPlain wing\n$R \\times 10^{-6}$ 2.5 4.2 5.9 7.2\n$\\circ$ $\\square$ $\\diamond$ $\\triangle$\n\nPlain wing + long fuselage\n$R \\times 10^{-6}$ 2.5 4.2 5.9 7.2\n$\\circ$ $\\square$ $\\diamond$ $\\triangle$\n\nDrag coefficient, $C_D$\n.08 .16 .24 .32 .40 .48 .56 .64\n\nAngle of attack, $\\alpha$, deg\n-8 0 8 16 24\n\nPitching-moment coefficient, $C_m$\n.08 0 -.08\n\nNACA\n\nFigure 8- Reynolds number effects on the lift, drag and pitching-moment characteristics.\n\nCONFIDENTIAL\n\nNACA RM No. A5C21\n```", "timestamp": "2026-07-22T06:54:34.309939+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 17, "total_pages": 55, "image_filename": "19930085966_p17.jpg", "text": "16 CONFIDENTIAL NACA RM L9B17\n\nCONCLUDING REMARKS\n\nThe thin-plate-burner configuration produced combustion efficiencies ranging between 56 and 72 percent in the ranges of variables covered by the tests. At the higher test fuel flows the burners exhibited marked decreases in efficiency with increasing fuel flows, which is believed to be caused by increasing maldistribution with increasing fuel flows. Because of this characteristic, operation of the thin-plate-burner configuration under high-thrust-output conditions is not considered feasible.\n\nThe ram-jet unit produced approximately constant thrust coefficients with variation of simulated flight Mach number for curves of constant combustion-chamber total-temperature ratio, variable combustion-chamber inlet Mach number, and constant nozzle exit area. Estimates of thrust coefficients at supersonic flight speeds for combustion-chamber performances limited to those obtained in the tests produced values regarded as too low to be practical.\n\nThe cycle-efficiency and propulsive-efficiency product of the ram-jet unit was approximately 80 percent of that for a no-pressure-loss unit under the same conditions of operation.\n\nThe performance of the intake diffuser, which had an area ratio of 2.14 to 1 and an equivalent conical angle of expansion of $16^\\circ$ was a unique function of the inlet-boundary-layer thickness. Over 99-percent diffuser efficiency was obtained when the boundary layer at the inlet was completely eliminated.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:54:35.424410+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 48, "total_pages": 92, "image_filename": "19930085870_p48.jpg", "text": "NACA RM No. L9D07\n49\n\nCONFIDENTIAL\n\nElliptical L.E. $\\begin{cases} \\bigcirc & C_L \\\\ \\square & C_m \\end{cases}$\nWedge L.E. $\\begin{cases} \\triangle & C_L \\\\ \\diamond & C_m \\end{cases}$\n\n| $C_L$ | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:54:38.117677+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 15, "total_pages": 30, "image_filename": "19930085975_p15.jpg", "text": "CONFIDENTIAL\n\nNACA RM L58J10\n\n[Figure: Test wing mounted on the free-roll sting in the Langley high-speed 7- by 10-foot tunnel.]\n\nNACA\nL-54508\n\nFigure 3.— Test wing mounted on the free-roll sting in the Langley high-speed 7- by 10-foot tunnel.\n\nCONFIDENTIAL\n\n13", "timestamp": "2026-07-22T06:54:38.627788+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 11, "total_pages": 132, "image_filename": "19930085990_p11.jpg", "text": "NACA RM A9I01 CONFIDENTIAL 9\n\nthe maximum lift at Mach numbers less than 0.80, and a forward movement of the aerodynamic center at low lift coefficients. The lift, drag, and pitching-moment characteristics of the wing-fuselage combination with the wing flaps deflected are presented in figure 9. Comparison of these data with those of figure 4 indicates that the addition of the fuselage caused a decrease in the maximum lift coefficient from 1.40 to 1.34 and an increase of $1^\\circ$ in the angle of attack for zero lift. The characteristics of the wing-fuselage combination were little affected by a change in Reynolds number from 6,000,000 to 10,000,000, but an increase from 2,000,000 to 6,000,000 resulted in a sizable decrease in the drag.\n\nWing, fuselage, and horizontal tail in the extended wing-chord plane.— Lift, drag, and pitching-moment characteristics of the complete semispan model with the horizontal tail mounted in the extended wing-chord plane are presented in figures 10, 11, and 12 for Mach numbers up to 0.95 and stabilizer angle settings from $4^\\circ$ to $-10^\\circ$. At a Mach number of 0.20, the aerodynamic center was shifted from 14 percent to 41 percent of the wing mean aerodynamic chord due to the addition of the tail. (See fig. 12(a).) As the Mach number was increased, the stabilizing effect of the horizontal tail was diminished to the extent that at a Mach number of 0.95 the horizontal tail made little or no contribution to the stability of the model at lift coefficients between $\\pm 0.3$. As will be discussed later, this decrease in the contribution of the tail to the stability was due to an increase in $\\partial \\epsilon / \\partial \\alpha$ and to a decrease in the dynamic-pressure ratio at the tail as the Mach number was increased. With a stabilizer angle setting of $0^\\circ$ and in a range of lift coefficients of about $\\pm 0.30$, the complete model was neutrally stable about the quarter point of the wing mean aerodynamic chord at a Mach number of about 0.87 and longitudinally unstable at higher Mach numbers. At lift coefficients greater than 0.30 stability existed at all test Mach numbers. The all-movable stabilizer provided sufficient longitudinal control to balance the airplane model at all Mach numbers up to 0.95 and at all angles of attack up to the stall. The value of $(\\partial C_m / \\partial C_L)_{C_L=0}$ was approximately $-0.036$ at a Mach number of 0.20 and increased slightly with increasing Mach number. (See fig. 12.)\n\nThe lift, drag, and pitching-moment characteristics of the complete semispan model with the wing flaps deflected are presented in figures 13, 14, and 15 for a Mach number of 0.20 and Reynolds numbers of 2,000,000, 3,000,000, and 10,000,000. At lift coefficients from zero to the maximum the complete model was longitudinally stable about the quarter point of the mean aerodynamic chord.\n\nWing, fuselage, and horizontal tail above the extended wing-chord plane.— To investigate the improvement in longitudinal stability and control afforded by raising the horizontal tail above the wing wake, tests were conducted with the model tail mounted 13 inches (0.696 wing mean aerodynamic chord) above the extended wing-chord plane.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:54:41.539213+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 67, "total_pages": 72, "image_filename": "19930085491_p67.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:54:42.455135+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 28, "total_pages": 92, "image_filename": "19930085951_p28.jpg", "text": "CONFIDENTIAL\nUNCLASSIFIED\n\nThrust coefficient, $C_T$\n\nAdvance ratio, $J$\n\n$\\beta_{0.75R}$ 20° 25° 30° 35° 40° 45° 50° NACA 55°\n\n(a) Thrust coefficient.\n\nFigure 8.— Characteristics of NACA 10-(3)(062)-045A propeller. Rotational speed, 1140 rpm.\n\nCONFIDENTIAL\n\n26\n\nNACA RM L9D29", "timestamp": "2026-07-22T06:54:43.026330+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 17, "total_pages": 30, "image_filename": "19930085970_p17.jpg", "text": "NACA RM A9E09 CONFIDENTIAL 15\n\nReynolds number, R, millions\nMach number, M\n\nFigure 3.- Variation with Mach number of Reynolds number based on the mean aerodynamic chord of the wing.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:54:48.821923+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 58, "total_pages": 149, "image_filename": "19930083192_p58.jpg", "text": "54\nNACA TN 1976\n\nTime-history data.- Time histories from the special investigations of the postwar period have been evaluated to obtain the flight miles to exceed the placard speed, the probable speed of flight for each sample, and the distribution of airspeed. Because of the limited duration of the samples, each sample was read to obtain the maximum speed during a fixed interval of time (6 to 10 minutes) and the resulting frequency distributions were then used to obtain the flight miles to exceed the placard speed. The results are included in table XXI and in figure 48. Figure 48 presents the percent of total flight time spent at speeds equal to or greater than any selected fraction of the high speed $V_L$. The probable speed noted in table XXI for airplanes E and F is the probable speed of flight and is not, as in the case of other airplanes, the probable speed at which maximum acceleration will occur. A simple analysis in which the velocity-acceleration envelopes were synthesized from the speed-frequency-distribution data and the gust-frequency data of reference 9 indicated that the probable speed for maximum acceleration was about $0.03V_L$ greater than the probable speed of flight.\n\nThe data obtained have also been evaluated according to flight condition to obtain curves of the flight hours required in climb, cruise, and descent to exceed given values of speed. Figure 49 is typical of the results obtained for airplane E.\n\nDisturbed motions.- From the data obtained in the special investigations a statistical study has been made of other quantities of interest in gust-load studies, such as the effects of disturbances on the relative frequency distribution and the fraction of the time spent in rough air.\n\nSome question has arisen as to the effect of the selected datum on the frequency distributions obtained in continuous rough air. This question is partly answered by the tests reported in reference 9 on one of the roughest flights with the XC-35 airplane where the relative gust-frequency distribution was determined by using a disturbed datum and an arbitrary 1 g datum. The results of the study are given in figure 50.\n\nCamera records of the pilot's instrument panel in the XC-35 airplane have been evaluated to obtain the relative frequency of occurrence of the maximum total variations in the angular displacements of the XC-35 airplane during separate traverses through clouds. The results are summarized in figure 51 as the relative frequency of exceeding selected values of yawing, rolling, and the pitching displacements and the rate of turn. It should be noted in considering this figure that the quantity plotted is the sum of the maximum positive and the maximum negative values recorded during each cloud traverse.\n\nPath ratio.- The available data for prewar transport operation given in reference 9 indicate that the path ratio (the miles of rough air divided by the total miles flown) varied from 0.24 to 0.006 for", "timestamp": "2026-07-22T06:54:52.631372+00:00"}
{"citation_id": "19930085997", "source_url": "https://ntrs.nasa.gov/api/citations/19930085997/downloads/19930085997.pdf", "page_number": 8, "total_pages": 40, "image_filename": "19930085997_p8.jpg", "text": "6\nCONFIDENTIAL\nNACA RM A9I29\n\nAngle of Attack of $0^\\circ$\n\nEffect of Mach number on pressure recovery.- Maximum total-pressure ratios $(H_5/H_0)_{max}$ as a function of the free-stream Mach number are presented in figure 5 for the four inlet configurations, A, B, C, and D, that were tested at an angle of attack of $0^\\circ$. From an examination of the curves in this figure, it is apparent that the changes in the inlet, shown in figure 3, improved the pressure recovery. Examination of schlieren photographs of the flow about inlet configuration A indicated that an expansion region originating at the leading edge of the boundary-layer scoop extended into the flow immediately ahead of the main inlet. This expansion was caused by improper alinement of the outer surface of the boundary-layer lip with the flow on the pilot enclosure. In order to eliminate this expansion in configuration B, the lip of the boundary-layer scoop was shaped to form a $5^\\circ$ angle with the surface of the cockpit enclosure, and the height of the boundary-layer scoop was increased to insure complete removal of the boundary layer. With this modification, the expansion ahead of the main scoop was replaced by an oblique shock wave and a greater pressure recovery resulted. By further increasing the angle formed by the outer surface of the boundary-layer scoop and the cockpit enclosure in configuration C, this oblique shock wave was strengthened and the total-pressure recovery was improved. However, the pressure recovery at the highest Mach number tested was still the same as that of configuration B. The decrease in pressure recovery at $M_0$ equal to 2.01 apparently was caused by the fact that the intersection of the stronger oblique shock wave and the normal shock wave was inboard of the lip of the scoop. Thus, the air that entered the scoop near the outer lip suffered larger losses in total pressure through a strong normal shock wave than was experienced by the air that passed through both an oblique and normal shock wave.\n\nHighest total-pressure recovery was obtained with configuration D. In figure 3 this configuration is shown to be similar to C, except that the leading edge of the boundary-layer scoop was extended farther ahead of the main inlet. The purpose of this modification was to retain the oblique shock strength of configuration C and to enable the oblique shock wave to extend across the inlet at a Mach number of 2.0. Schlieren photographs and line drawings of the shock-wave patterns ahead of the inlets of this configuration are shown in figure 6. At a free-stream Mach number of 2.0 the oblique shock wave from the leading edge of the boundary-layer scoop is shown to intersect the normal shock wave at a point slightly outboard of the scoop. Thus, all the air that entered the inlets underwent compression through the oblique shock wave before encountering the normal shock wave.\n\nSince the flow in the settling chamber of the model was diffused to a Mach number much lower than the usual intake Mach number at the compressor of a turbojet engine, the measured values of $(H_5/H_0)_{max}$ include diffusion losses that would not occur in an airplane.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:54:53.429056+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 30, "total_pages": 47, "image_filename": "19930085919_p30.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 29\n\nLift coefficient, $C_L$\n\n$\\delta_f, 45^\\circ$\n\n$\\delta_f, 60^\\circ$\n\n$\\delta_f, 75^\\circ$\n\nAngle of attack, $\\alpha$, deg Pitching-moment coefficient, $C_m$\n\n(a) $C_L$ vs $\\alpha$ and $C_m$.\n\nFigure 9- Effect of the split flaps in several positions on the lift, drag, and pitching-moment characteristics of the model. $R, 4.2 \\times 10^6$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:54:54.328820+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 13, "total_pages": 25, "image_filename": "19930085979_p13.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:54:56.724793+00:00"}
{"citation_id": "19930085999", "source_url": "https://ntrs.nasa.gov/api/citations/19930085999/downloads/19930085999.pdf", "page_number": 1, "total_pages": 20, "image_filename": "19930085999_p1.jpg", "text": "RESTRICTED\nCopy\nRM E9I07\n\nNACA RM E9I07\n\nNACA\n\nRESEARCH MEMORANDUM\n\nVIBRATION OF LOOSELY MOUNTED TURBINE BLADES DURING SERVICE\nOPERATION OF A TURBOJET ENGINE WITH CENTRIFUGAL COMPRESSOR\nAND STRAIGHT-FLOW COMBUSTION CHAMBERS\n\nBy W. C. Morgan, R. H. Kemp\nand S. S. Manson\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nAUTHORITY CROWLEY CHANGE #1924\nDATE 12-14-53\nT.C.F.\n\nCLASSIFIED DOCUMENT\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\nWASHINGTON\nNovember 3, 1949\n\nRESTRICTED", "timestamp": "2026-07-22T06:54:56.988706+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 8, "total_pages": 24, "image_filename": "19930085991_p8.jpg", "text": "6\nNACA RM L9I28\n\nand unstable descent for model 2 then fell in the same regions obtained\nby drawing the boundary line for the other four models. The plot\nand boundary line are presented in figure 4.\n\nAs can be seen in figure 4, the results of the study indicate that\nit is difficult to achieve nose-down stability of an inherently unstable\nnose section by merely moving the center of gravity forward; the results\nalso indicate that if the center of gravity is too far rearward even\nvery large fins might not make a nose section stable.\n\nIn addition to the fin-stabilization factor and the center-of-\ngravity location, it is expected that other factors, such as mass\ndistribution, fineness ratio, and body lines, may affect the boundary in\nfigure 4 somewhat but are apparently of only secondary importance. The\nfin-stabilization factor is proportional to the static longitudinal\nstability factor $C_{m\\alpha}$, except for omission of the $C_{L\\alpha}$ term which\nnormally is greatly influenced by aspect ratio. The present empirical\nresults, however, did not indicate an appreciable effect on the boundary\nof varied fin aspect ratio within the range investigated. From these\nresults, it appears that the boundary may be used as an empirical\ncriterion to indicate the fin area required to stabilize an airplane\njettisonable nose section having a pointed front, and from the inter-\npretation of results obtained with model 2, it appears that the boundary\nmay also be used to obtain an indication of the fin area required for\nstabilizing a nose section with other than a pointed front.\n\nAnother possible method of approach to the problem of selecting\nsuitable stabilizing fins for a specific nose design might consist of\ncalculating the instability of the nose section and the stabilizing\neffect of the fins. In such a method, it will probably be necessary to\nconsider both static- and dynamic-stability parameters or use some\nempirical correction to allow for the dynamic-stability effects. In\norder to illustrate a possible approach, brief static-stability\ncalculations have been made for model 1 of the present investigation,\nwith and without a set of four triangular stabilizing fins of\narrangement a installed. The span of the fins considered was 27 percent\nof the nose length and the aspect ratio was 2.\n\nThe instability of the nose section without fins was calculated by\nthe equation\n\n$$C_{m\\alpha} = \\frac{2}{S_p L} \\int_0^L \\frac{dA}{dx} (\\cos^2\\epsilon)(a - x)dx$$", "timestamp": "2026-07-22T06:54:57.778029+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 18, "total_pages": 55, "image_filename": "19930085966_p18.jpg", "text": "NACA RM L9B17 CONFIDENTIAL 17\n\nAPPENDIX\n\nESTIMATION OF THRUST COEFFICIENTS FOR\nSUPERSONIC FLIGHT SPEEDS\n\nThe range of simulated flight Mach number obtained in the tests\nwas limited and the relation of the low Mach number data to possible\nhigh Mach number performance was not obvious; therefore, an estimate\nbased on the subsonic test-stand data was made of the thrust-coefficient\nvariation with flight Mach number. The combustion-chamber performance\nin terms of Mach numbers and pressure and temperature ratios was held\nto those values obtained in the test regardless of the flight Mach\nnumber. It was believed reasonable to restrict the combustion-chamber\ninlet velocity by limiting the inlet Mach number to test values since\nmost ram-jet burners depreciate in performance if the air velocity is\nincreased beyond certain values. This limitation would be imposed\nphysically by regulation of the nozzle exit area. The limitation of\nthe combustion-chamber temperature ratios to test values is considered\nconservative since the higher levels of pressures and temperatures\nassociated with higher flight Mach numbers are favorable to combustion.\nIt was further assumed that the friction and turbulence loss characteris-\ntics of the combustion chamber remained unchanged.\n\nIn attempting to derive expressions relating the temperatures and\npressures before and after combustion in a tube of constant cross-sectional\narea the problem arises as to how to account for friction and turbulence\nlosses. Actually friction and turbulence losses occur along the entire\nlength of the chamber, the amount of loss over any one section depending\non the chamber and burner design. To attempt to write such an exact\nfriction loss distribution into combustion equations would be extremely\ndifficult. For the purpose of this presentation it will be assumed that\nthe over-all loss can be represented in two parts, the first part being\nproportional to the dynamic pressure before combustion and the second part\nexpressed as being proportional to the dynamic pressure after combustion,\nthus the sum of $K_3q_3$ and $K_4q_4$ is equal to the total loss. On this\nbasis the following expressions can be written:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:01.089331+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 16, "total_pages": 30, "image_filename": "19930085975_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:55:02.089205+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 8, "total_pages": 17, "image_filename": "19930085988_p8.jpg", "text": "NACA RM L9H30\n7\n\nCONFIDENTIAL\nTABLE I\nBODY AND WING COORDINATES FOR TEST MODELS\n\n[Figure: Diagram of a body shape with X and r coordinates indicated]\nBody coordinates in inches\n\n| Body coordinates | | | |\n| :--- | :--- | :--- | :--- |\n| *130-inch transonic model* | | | |\n| X | r | X | r |\n| 0.000 | 0.000 | 54.600 | 6.135 |\n| 0.780 | 0.360 | 62.400 | 6.339 |\n| 1.170 | 0.465 | 70.200 | 6.462 |\n| 1.950 | 0.668 | 78.000 | 6.500 |\n| 3.900 | 1.126 | 85.800 | 6.442 |\n| 7.800 | 1.880 | 93.620 | 6.276 |\n| 11.700 | 2.517 | 101.400 | 5.993 |\n| 15.600 | 3.075 | 109.200 | 5.556 |\n| 23.400 | 4.046 | 117.000 | 4.880 |\n| 31.200 | 4.820 | 124.800 | 3.940 |\n| 39.000 | 5.405 | 130.000 | 3.231 |\n| 46.800 | 5.836 | | |\n| *Nose radius = .078 inch* | | | |\n\n[Figure: Diagram of a wing shape with X and y coordinates indicated]\nWing coordinates in percent chord\n\n| Wing coordinates | | | |\n| :--- | :--- | :--- | :--- |\n| *NACA 65A006* | | | |\n| X | y | X | y |\n| 0.00 | 0.000 | 40.00 | 2.996 |\n| 0.50 | 0.464 | 45.00 | 2.992 |\n| 0.75 | 0.563 | 50.00 | 2.925 |\n| 1.25 | 0.710 | 55.00 | 2.793 |\n| 2.50 | 0.991 | 60.00 | 2.602 |\n| 5.00 | 1.313 | 65.00 | 2.364 |\n| 7.50 | 1.591 | 70.00 | 2.087 |\n| 10.00 | 1.824 | 75.00 | 1.775 |\n| 15.00 | 2.194 | 80.00 | 1.437 |\n| 20.00 | 2.474 | 85.00 | 1.083 |\n| 25.00 | 2.687 | 90.00 | 0.727 |\n| 30.00 | 2.842 | 95.00 | 0.370 |\n| 35.00 | 2.945 | 100.00 | 0.013 |\n| *L.E. radius = .229% c T.E. radius = .014% c* | | | |\n\n[Figure: NACA logo]\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:55:03.811582+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 97, "total_pages": 114, "image_filename": "19930086061_p97.jpg", "text": "```markdown\nNACA RM L9J07\n93\n\n<!-- Image (148, 103, 899, 876) -->\n\nFigure 42.- Span load distribution of wing 2 at various angles of attack; $\\psi = 10^\\circ$. Flagged symbols represent data taken with left semispan at $\\psi = -10^\\circ$.\n```", "timestamp": "2026-07-22T06:55:05.488937+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 49, "total_pages": 92, "image_filename": "19930085870_p49.jpg", "text": "```markdown\n50\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\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(i) Wing 9. w=1.280; R = 700,000.\nFigure 6.-Continued.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:55:06.940186+00:00"}