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  Partner: UNT Libraries Government Documents Department
 Collection: National Advisory Committee for Aeronautics Collection
Flow and Force Equations for a Body Revolving in a Fluid
A general method for finding the steady flow velocity relative to a body in plane curvilinear motion, whence the pressure is found by Bernoulli's energy principle is described. Integration of the pressure supplies basic formulas for the zonal forces and moments on the revolving body. The application of the steady flow method for calculating the velocity and pressure at all points of the flow inside and outside an ellipsoid and some of its limiting forms is presented and graphs those quantities for the latter forms. In some useful cases experimental pressures are plotted for comparison with theoretical. The pressure, and thence the zonal force and moment, on hulls in plane curvilinear flight are calculated. General equations for the resultant fluid forces and moments on trisymmetrical bodies moving through a perfect fluid are derived. Formulas for potential coefficients and inertia coefficients for an ellipsoid and its limiting forms are presented.
The Inertia Coefficients of an Airship in a Frictionless Fluid
The apparent inertia of an airship hull is examined. The exact solution of the aerodynamical problem is studied for hulls of various shapes with special attention given to the case of an ellipsoidal hull. So that the results for the ellipsoidal hull may be readily adapted to other cases, they are expressed in terms of the area and perimeter of the largest cross section perpendicular to the direction of motion by means of a formula involving a coefficient kappa which varies only slowly when the shape of the hull is changed, being 0.637 for a circular or elliptic disk, 0.5 for a sphere, and about 0.25 for a spheroid of fineness ratio. The case of rotation of an airship hull is investigated and a coefficient is defined with the same advantages as the corresponding coefficient for rectilinear motion.
Empirical relation between induced velocity, thrust, and rate of descent of a helicopter rotor as determined by wind-tunnel tests on four model rotors
The empirical relation between the induced velocity, thrust, and rate of vertical descent of a helicopter rotor was calculated from wind tunnel force tests on four model rotors by the application of blade-element theory to the measured values of the thrust, torque, blade angle, and equivalent free-stream rate of descent. The model tests covered the useful range of C(sub t)/sigma(sub e) (where C(sub t) is the thrust coefficient and sigma(sub e) is the effective solidity) and the range of vertical descent from hovering to descent velocities slightly greater than those for autorotation. The three bladed models, each of which had an effective solidity of 0.05 and NACA 0015 blade airfoil sections, were as follows: (1) constant-chord, untwisted blades of 3-ft radius; (2) untwisted blades of 3-ft radius having a 3/1 taper; (3) constant-chord blades of 3-ft radius having a linear twist of 12 degrees (washout) from axis of rotation to tip; and (4) constant-chord, untwisted blades of 2-ft radius. Because of the incorporation of a correction for blade dynamic twist and the use of a method of measuring the approximate equivalent free-stream velocity, it is believed that the data obtained from this program are more applicable to free-flight calculations than the data from previous model tests.
Development of a supersonic area rule and an application to the design of a wing-body combination having high lift-to-drag ratios
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Evolution of the helicopter
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Wing pressure distributions over the lift range of the convair xf-92a delta-wing airplane at subsonic and transonic speeds
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Determination of rate, area, and distribution of impingement of waterdrops on various airfoils from trajectories obtained on the differential analyzer
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Moments of cambered round bodies
Results are presented for the moments and position of force centers of a series of cambered round bodies derived from a torpedo-like body of revolution. The effects of placing fins on the rear of the body of revolution are also included.
Compilation and Analysis of US Turbojet and Ram-Jet Engine Characteristics
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The factors that determine the minimum speed of an airplane
The author argues that because of a general misunderstanding of the principles of flight at low speed, there are a large number of airplanes that could be made to fly several miles per hour slower than at present by making slight modifications. In order to show how greatly the wing section affects the minimum speed, curves are plotted against various loadings. The disposition of wings on the airplane slightly affects the lift coefficient, and a few such cases are discussed. Another factor that has an effect on minimum speed is the extra lift exerted by the slip stream on the wings. Also discussed are procedures to be followed by the pilot, especially with regard to stick movements during low speed flight. Also covered are stalling, yaw, rolling moments, lateral control, and the effectiveness of ailerons and rudders.
The testing of aviation engines
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Material compatibility with gaseous fluorine
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The Design of Airplane-engine Superchargers
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Droplet Impingement and Ingestion by Supersonic Nose Inlet in Subsonic Tunnel Conditions
The amount of water in cloud droplet form ingested by a full-scale supersonic nose inlet with conical centerbody was measured in the NACA Lewis icing tunnel. Local and total water impingement rates on the cowl and centerbody surfaces were also obtained. All measurements were made with a dye-tracer technique. The range of operating and meteorological conditions studied was: angles of attack of 0 deg and 4.2 deg, volume-median droplet diameters from about 11 to 20 microns, and ratios of inlet to free-stream velocity from about 0.4 to 1.8. Although the inlet was designed for supersonic (Mach 2.0) operation of the aircraft, the tunnel measurements were confined to a free-stream velocity of 156 knots (Mach 0.237). The data are extendable to other subsonic speeds and droplet sizes by dimensionless impingement parameters. Impingement and ingestion efficiencies are functions of the ratio of inlet to free-stream velocity as well as droplet size. For the model and range of conditions studied, progressively increasing the inlet velocity ratio from less than to greater than 1.0 increased the centerbody impingement efficiency and shifted the cowl impingement region from the inner- to outer-cowl surfaces, respectively. The ratio of water ingested by the inlet plane to that contained in a free-stream tube of cross section equal to that at the inlet plane also increased with increasing inlet velocity ratio. Theoretically calculated values of inlet water (or droplet) ingestion are in good agreement with experiment for annular inlet configurations.
A Dye-Tracer Technique for Experimentally Obtaining Impingement Characteristics of Arbitrary Bodies and a Method for Determining Droplet Size Distribution
A dye-tracer technique has been developed whereby the quantity of dyed water collected on a blotter-wrapped body exposed to an air stream containing a dyed-water spray cloud can be colorimetrically determined in order to obtain local collection efficiencies, total collection efficiency, and rearward extent of impingement on the body. In addition, a method has been developed whereby the impingement characteristics obtained experimentally for a body can be related to theoretical impingement data for the same body in order to determine the droplet size distribution of the impinging cloud. Several cylinders, a ribbon, and an aspirating device to measure cloud liquid-water content were used in the studies presented herein for the purpose of evaluating the dye-tracer technique. Although the experimental techniques used in the dye-tracer technique require careful control, the methods presented herein should be applicable for any wind tunnel provided the humidity of the air stream can be maintained near saturation.
Effect of Ice and Frost Formations on Drag of NACA 65(sub 1) -212 Airfoil for Various Modes of Thermal Ice Protection
The effects of primary and. runback icing and frost formations on the drag of an 8-foot-chord NACA 651-212 airfoil section were investigated over a range of angles of attack from 20 to 80 and airspeeds up to 260 miles per hour for icing conditions with liquid-water contents ranging from 0.25 to 1.4 grams per cubic meter and datum air temperatures of -30 to 30 F. The results showed that glaze-ice formations, either primary or runback, on the upper surface near the leading edge of the airfoil caused large and rapid increases in drag, especially at datum air temperatures approaching 32 F and in the presence of high rates of water catch. Ice formations at lower temperatures (rime ice) did not appreciably increase the drag coefficient over the initial (standard roughness) drag coefficient. Cyclic de-icing of the primary Ice formations on the airfoil leading-edge section permitted the drag coefficient to return almost to the bare airfoil drag value. Runback icing on the lower surface did not present a serious drag problem except when heavy spanwise ridges of runback ice occurred aft of the heatable area. Frost formations caused rapid and large increases in drag with incipient stalling of the airfoil.
Effect of Ice Formations on Section Drag of Swept NACA 63A-009 Airfoil with Partial-Span Leading-Edge Slat for Various Modes of Thermal Ice Protection
The effects of primary and runback ice formations on the section drag of a 36 deg swept NACA 63A-009 airfoil section with a partial-span leading-edge slat were studied over a range of angles of attack from 2 to 8 deg and airspeeds up to 260 miles per hour for icing conditions with liquid-water contents ranging from 0.39 to 1.23 grams per cubic meter and datum air temperatures from 10 to 25 F. The results with slat retracted showed that glaze-ice formations caused large and rapid increases in section drag coefficient and that the rate of change in section drag coefficient for the swept 63A-009 airfoil was about 2-1 times that for an unswept 651-212 airfoil. Removal of the primary ice formations by cyclic de-icing caused the drag to return almost to the bare-airfoil drag value. A comprehensive study of the slat icing and de-icing characteristics was prevented by limitations of the heating system and wake interference caused by the slat tracks and hot-gas supply duct to the slat. In general, the studies showed that icing on a thin swept airfoil will result in more detrimental aerodynamic characteristics than on a thick unswept airfoil.
Determination of rate, area, and distribution of impingement of of waterdrops on various airfoils from trajectories obtained on the differential analyzer
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Effectiveness of Thermal-Pneumatic Airfoil-Ice-Protection System
Icing and drag investigations were conducted in the NACA Lewis icing research tunnel employing a combination thermal-pneumatic de-icer mounted on a 42-inch-chord NACA 0018 airfoil. The de-icer consisted of a 3-inch-wide electrically heated strip symmetrically located about the leading edge with inflatable tubes on the upper and lower airfoil surfaces aft of the heated area. The entire de-icer extended to approximately 25 percent of chord. A maximum power density of 9.25 watts per square inch was required for marginal ice protection on the airfoil leading edge at an air temperature of 00 F and an airspeed of 300 miles per hour. Drag measurements indicated, that without icing, the de-icer installation increased the section drag to approximately 140 percent of that of the bare airfoil; with the tubes inflated, this value increased to a maximum of approximately 620 percent. A 2-minute tube-inflation cycle prevented excessive ice formation on the inflatable area although small scattered residual Ice formations remained after inflation and were removed intermittently during later cycles. Effects of the time lag of heater temperatures after initial application of power and the insulating effect of ice formations on heater temperatures were also determined.
Effects of surface roughness and extreme cooling on boundary-layer transition for 15 deg cone-cylinder in free flight at Mach numbers to 7.6
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Effects of Ice Formations on Airplane Performance in Level Cruising Flight
A flight investigation in natural icing conditions was conducted by the NACA to determine the effect of ice accretion on airplane performance. The maximum loss in propeller efficiency encountered due to ice formation on the propeller blades was 19 percent. During 87 percent of the propeller icing encounters, losses of 10 percent or less were observed. Ice formations on all of the components of the airplane except the propellers during one icing encounter resulted in an increase in parasite drag of the airplane of 81 percent. The control response of the airplane in this condition was marginal.
Effects of extreme surface cooling on boundary-layer transition
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Correlations Among Ice Measurements, Impingement Rates Icing Conditions, and Drag Coefficients for Unswept NACA 65A004 Airfoil
An empirical relation has been obtained by which the change in drag coefficient caused by ice formations on an unswept NACA 65AO04 airfoil section can be determined from the following icing and operating conditions: icing time, airspeed, air total temperature, liquid-water content, cloud droplet impingement efficiencies, airfoil chord length, and angles of attack. The correlation was obtained by use of measured ice heights and ice angles. These measurements were obtained from a variety of ice formations, which were carefully photographed, cross-sectioned, and weighed. Ice weights increased at a constant rate with icing time in a rime icing condition and at progressively increasing rates in glaze icing conditions. Initial rates of ice collection agreed reasonably well with values predicted from droplet impingement data. Experimental droplet impingement rates obtained on this airfoil section agreed with previous theoretical calculations for angles of attack of 40 or less. Disagreement at higher angles of attack was attributed to flow separation from the upper surface of the experimental airfoil model.
Wind tunnel force tests in wing systems through large angles of attack
Force tests on a systematic series of wing systems over a range of angle of attack from minus forty-five degrees to plus ninety degrees are covered in this report. The investigation was made on monoplane and biplane wing models to determine the effects of variations of tip shape, aspect ratio, flap setting, stagger, gap, decalage, sweepback, and airfoil profile.
Wind-Tunnel Investigation at a Mach Number of 2.01 of the Aerodynamic Characteristics in Combined Angles of Attack and Sideslip of Several Hypersonic Missile Configurations with Various Canard Controls
An investigation of the aerodynamic characteristics of several hypersonic missile configurations with various canard controls for an angle-of-attack range from 0 deg to about 28 deg at sideslip angles of about 0 deg and 4 deg at a Mach number of 2.01 has been made in the Langley 4- by 4-foot supersonic pressure tunnel. The configurations tested were a body alone which had a ratio of length to diameter of 10, the body with a 10 deg flare, the body with cruciform fins of 5 deg or 15 deg apex angle, and a flare-stabilized rocket model with a modified Von Karman nose. Various canard surfaces for pitch control only were tested on the body with the 10 deg flare and on the body with both sets of fins. The results indicated that the addition of a flared afterbody or cruciform fins produced configurations which were longitudinally and directionally stable. The body with 5 deg fins should be capable of producing higher normal accelerations than the flared body. A l l of the canard surfaces were effective longitudinal controls which produced net positive increments of normal force and pitching moments which progressively decreased with increasing angle of attack.
Tabulated Pressure Data for a Series of Controls on a 40 Deg Sweptback Wing at Mach Numbers of 1.61 and 2.01
An investigation has been made at Mach numbers of 1.61 and 2.01 and Reynolds numbers of 1.7 x l0(exp 6) and 3.6 x l0(exp 6) to determine the pressure distributions over a swept wing with a series of 14 control configurations. The wing had 40 deg of sweep of the quarter-chord line, an aspect ratio of 3.1, and a taper ratio of 0.4. Measurements were made at angles of attack from 0 deg to +/- 15 deg for control deflections from -60 deg to 60 deg. This report contains tabulated pressure data for the complete range of test conditions.
Notes on the N.A.C.A. control force recorder
Emphasized here is the desirability of using recording instruments in the investigation of the characteristics of airplanes with particular reference to the National Advisory Committee for Aeronautics (NACA) control force recorder. Given here are photographs, records, and a description of the instrument developed by NACA for investigations on different types of aircraft. Described here is an instrument for recording control forces. At present, this control force recorder registers only the forces exerted on the stick. However, attachments are being designed to enable the forces on the rudder bar also to be recorded. The instrument in its final form will consist of three parts, namely, the recorder, the controller for the stick, and the controller for the rudder. The first two are in use now. The theory of operation is simple. In the controller, which is slipped over and fastened to the stick, are small electrical resistances which vary with the force applied to the handle. The recording apparatus then consists of suitable variable resistances properly connected to galvanometers whose deflections are proportional to the forces applied to the stick.
Notes on the theory of the accelerometer
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Aerodynamic Effects Caused by Icing of an Unswept NACA 65A004 Airfoil
The effects of ice formations on the section lift, drag, and pitching-moment coefficients of an unswept NACA 65A004 airfoil section of 6-foot chord were studied.. The magnitude of the aerodynamic penalties was primarily a function of the shape and size of the ice formation near the leading edge of the airfoil. The exact size and shape of the ice formations were determined photographically and found to be complex functions of the operating and icing conditions. In general, icing of the airfoil at angles of attack less than 40 caused large increases in section drag coefficients (as much as 350 percent in 8 minutes of heavy glaze icing), reductions in section lift coefficients (up to 13 percent), and changes in the pitching-moment coefficient from diving toward climbing moments. At angles of attack greater than 40 the aerodynamic characteristics depended mainly on the ice type. The section drag coefficients generally were reduced by the addition of rime ice (by as much as 45 percent in 8 minutes of icing). In glaze icing, however, the drag increased at these angles of attack. The section lift coefficients were variably affected by rime-ice formations; however, in glaze icing, lift increases at high angles of attack amounted to as much as 9 percent for an icing time of 8 minutes. Pitching-moment-coefficient changes in icing conditions were somewhat erratic and depended on the icing condition. Rotation of the iced airfoil to angles of attack other than that at which icing occurred caused sufficiently large changes in the pitching-moment coefficient that, in flight, rapid corrections in trim might be required in order to avoid a hazardous situation.
Comparison of Several Methods of Cyclic De-Icing of a Gas-Heated Airfoil
Several methods of cyclic de-icing of a gas-heated airfoil were investigated to determine ice-removal characteristics and heating requirements. The cyclic de-icing system with a spanwise ice-free parting strip in the stagnation region and a constant-temperature gas-supply duct gave the quickest and most reliable ice removal. Heating requirements for the several methods of cyclic de-icing are compared, and the savings over continuous ice prevention are shown. Data are presented to show the relation of surface temperature, rate of surface heating, and heating time to the removal of ice.
Compatibility of Metals with Liquid Fluorine at High Pressures and Flow Velocities
No abstract available.
Comparison of Heat Transfer from Airfoil in Natural and Simulated Icing Conditions
An investigation of the heat transfer from an airfoil in clear air and in simulated icing conditions was conducted in the NACA Lewis 6- by 9-foot icing-research tunnel in order to determine the validity of heat-transfer data as obtained in the tunnel. This investiation was made on the same model NACA 65,2-016 airfoil section used in a previous flight study, under similar heating, icing, and operating conditions. The effect of tunnel turbulence, in clear air and in icingwas indicated by the forward movement of transition from laminar to turbulent heat transfer. An analysis of the flight results showed the convective heat transfer in icing to be considerably different from that measured in clear air and. only slightly different from that obtained in the icing-research tunnel during simulated icing.
Wind-tunnel investigation of several factors affecting the performance of a high-speed pursuit airplane with air-cooled radial engine
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Preliminary wind-tunnel investigation of the effect of area suction on the laminar boundary layer over an NACA 64A010 airfoil
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Preliminary Wind-Tunnel Investigation of the Performance of Republic F-105 Wing-Root Inlet Configurations at Various Angles of Attack and a Mach Number of 2.01
A 1/13-scale model of the forebody of the Republic F-105 with twin-duct wing-root inlets was tested in the Langley 4- by 4-foot supersonic pressure tunnel through a range of angle of attack from -4 deg to 15 deg at a Mach number of 2.01 and a Reynolds number of approximately 3.4 x 10(exp 6) per foot. The tests were made with four configurations which incorporated varying amounts of sweep and stagger of the inlet leading edges, modifications to the areas of the boundary-layer diverter floor plate, and modifications to the area of the boundary-layer diverter bleed slots. The highest overall pressure recovery at an angle of attack of 0 deg (average total-pressure recovery, 0.84 mass-flow ratio, 0.98) was achieved with configuration having an inlet leading-edge sweep angle of 58 deg with no leading-edge stagger. Stagger was found to improve the angle-of- attack performance, but at a sacrifice in inlet efficiency for an angle of attack of 0 deg. The boundary-layer diverter floor height, of the order of one boundary-layer thickness, was satisfactory for bypassing the fuselage boundary layer. The boundary-layer diverter-plate bleed slots were effective in increasing the total-pressure recovery of the inlet. The total-pressure-recovery contour plots, taken at the compressor-face station, indicate the existence of high-velocity "cores" throughout the inlet operating range.
Analysis of a Pneumatic Probe for Measuring Exhaust-Gas Temperatures with Some Preliminary Experimental Results
A pneumatic probe based on continuity of mass flow through two restrictions separated by a cooling chamber was constructed to measure gas temperature at and beyond the limit of thermocouples. This probe consisted of a subsonic flat-plate orifice for the first restriction and a sonic-flow converging-diverging nozzle for the second restriction. The effect of variation in gas constants on the calibration is examined for common engine-exhaust gases. A high-temperature wind tunnel that allowed calibration of the probe at temperatures up to 2000 deg R and. Mach numbers up to 0.8 is described. Agreement to better than 30 deg R between pneumatic probe indication and the indication of a rake of radiation shielded thermocouples indicates that extrapolation of the calibration to higher temperatures is possible with fair accuracy.
Adaptation of a Cascade Impactor to Flight Measurement of Droplet Size in Clouds
A cascade impactor, an instrument for obtaining: the size distribution of droplets borne in a low-velocity air stream, was adapted for flight cloud droplet-size studies. The air containing the droplets was slowed down from flight speed by a diffuser to the inlet-air velocity of the impactor. The droplets that enter the impactor impinge on four slides coated with magnesium oxide. Each slide catches a different size range. The relation between the size of droplet impressions and the droplet size was evaluated so that the droplet-size distributions may be found from these slides. The magnesium oxide coating provides a permanent record. of the droplet impression that is not affected by droplet evaporation after the. droplets have impinged.
Analogy Between Mass and Heat Transfer with Turbulent Flow
An analysis of combined heat and mass transfer from a flat plate has been made in terms of Prandtl t s simplified physical concept of the turbulent boundary layer. The results of the analysis show that for conditions of reasonably small heat and mass transfer, the ratio of the mass-and heat-transfer coefficients is dependent on the Reynolds number of the boundary layer, the Prandtl number of the medium of diffusion, and the Schmidt number of the diffusing fluid in the medium of diffusion. For the particular case of water evaporating into air, the ratio of mass-transfer coefficient to heat-transfer coefficient is found to be slightly greater than unity.
Analytical Investigation of Icing Limit for Diamond-Shaped Airfoil in Transonic and Supersonic Flow
Calculations have been made for the icing limit of a diamond airfoil at zero angle of attack in terms of the stream Mach number, stream temperature, and pressure altitude. The icing limit is defined as a wetted-surface temperature of 320 F and is related to the stream conditions by the method of Hardy. The results show that the point most likely to ice on the airfoil lies immediately behind the shoulder and is subject to possible icing at Mach numbers as high as 1.4.
The art of writing scientific reports
As the purpose of the report is to transmit as smoothly and as easily as possible, certain facts and ideas, to the average person likely to read it, it should be written in a full and simple enough manner to be comprehended by the least tutored, and still not be boring to the more learned readers.
Analysis of Meteorological Data Obtained During Flight in a Supercooled Stratiform Cloud of High Liquid-Water Content
Flight icing-rate data obtained in a dense and. abnormally deep supercooled stratiform cloud system indicated the existence of liquid-water contents generally exceeding values in amount and extent previously reported over the midwestern sections of the United States. Additional information obtained during descent through a part of the cloud system indicated liquid-water contents that significantly exceeded theoretical values, especially near the middle of the cloud layer.. The growth of cloud droplets to sizes that resulted in sedimentation from the upper portions of the cloud is considered to be a possible cause of the high water contents near the center of the cloud layer. Flight measurements of the vertical temperature distribution in the cloud layer indicated a rate of change of temperature with altitude exceeding that of the moist adiabatic lapse rate. This excessive rate of change is considered to have contributed to the severity of the condition.
The comparison of well-known and new wing sections tested in the variable density wind tunnel
Three groups of airfoils have been tested in the variable density wind tunnel. The first group contains three airfoils. The second group is a systematic series of twenty-seven airfoils. The third group consists of several frequently used wing sections.
Experimental investigation of effect of spike- tip and cowl-lip blunting on the internal performance of a two-cone cylindrical-cowl inlet at mach number 4.95
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Experimental investigation of air-cooled turbine blades in turbojet engine. 2: Rotor blades with 15 fins in cooling-air passages
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Experimental Investigation of Effects of Combustion-chamber Length and Inlet Total Temperature, Total Pressure, and Velocity on Afterburner Performance
Afterburner combustion efficiency - effects of chamber length, temperature, pressure and velocity.
Experimental investigation of diffuser pressure ratio control with shock positioning limit on 28 inch ram-jet engine
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Experimental performance of a 5000-pound- thrust rocket chamber using a 20-percent- fluorine - 80-percent-oxygen mixture with RP-1
Increased performance of rocket engine using fluorine-oxygen mixture with RP-1 fuel.
Experimental investigation of air-cooled turbine blades in turbojet engine. 1: Rotor blades with 10 tubes in cooling-air passages
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Experimental investigation of extreme internal flow turning at the cowl lip of an axisymetric inlet at a Mach number of 2.95
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Experimental Sea-level Static Investigation of a Short Afterburner
Sea-level static testing of turbojet engine afterburner.