National Advisory Committee for Aeronautics (NACA) - 183 Matching Results

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Accelerations in fighter-airplane crashes
From Introduction: "This report describes some measurements of these quantities obtained by crashing fighter aircraft under circumstances approximating those observed in service."
An active particle diffusion theory of flame quenching for laminar flames
An equation for quenching distance based on the destruction of chain carriers by the surface is derived. The equation expresses the quenching distance in terms of the diffusion coefficients and partial pressures of the chain carriers and gas phase molecules, the efficiency of the surface as a chain breaker, the total pressure of the mixture, and a constant which depends on the geometry of the quenching surface. Quenching distances measured by flashback for propane-air flames are shown to be consistent with the mechanism. The derived equation is used with the lean inflammability limit and a rate constant calculated from burning velocity data to estimate quenching distances for propane-air (hydrocarbon lean) flames satisfactorily.
Aerodynamic Characteristics at a Mach Number of 6.8 of Two Hypersonic Missile Configurations, One With Low-Aspect-Ratio Cruciform Fins and Trailing-Edge Flaps and One With a Flared Afterbody and All-Movable Controls
Report presenting an investigation to determine the aerodynamic characteristics in pitch at a Mach number of 6.8 of hypersonic missile configurations with cruciform trailing-edge flaps and all-movable control surfaces. Testing indicated that all-movable controls on the flared-afterbody model should be capable of producing much larger values of trim lift and of normal acceleration than the trailing-edge-flap configuration. Some of the configurations tested include body alone, body with 5 degree fins and trailing-edge flaps, and body with 10 degree flare and all-movable controls.
Aerodynamic Characteristics at a Mach Number of 6.8 of Two Hypersonic Missile Configurations, One With Low-Aspect-Ratio Cruciform Fins and Trailing-Edge Flaps and One With a Flared Afterbody and All-Movable Controls
Report discussing an investigation to determine the aerodynamic characteristics of hypersonic missile configurations with cruciform trailing-edge flaps with all-movable control surfaces. The all-movable controls were found to produce much larger values of trim lift and normal acceleration than the trailing-edge-flap configuration.
Aerodynamic characteristics at Mach numbers 2.36 and 2.87 of an airplane configuration having a cambered arrow wing with a 75 degree swept leading edge
From Introduction: "The results obtained in the wind-tunnel tests at Mach numbers 2.36 and 2.87 for several configurations utilizing this wing, including results on the wing alone are presented."
Aerodynamic characteristics of a 6-percent-thick symmetrical double-wedge airfoil at transonic speeds from tests by the NACA wing-flow method
From Introduction: "The investigation covered a range of Mach numbers from 0.66 to 1.12 and included measurements of angle of attack, pitching moment, normal force, and chord force. The drag at zero lift obtained in this investigation was reported in reference 1, but without the correction for tare of the end plate."
Aerodynamic Characteristics of a 45 Degree Swept-Back Wing With Aspect Ratio of 3.5 and NACA 2S-50(05)-50(05) Airfoil Sections
From Introduction: "The present paper presents the scale effect on the longitudinal aerodynamic characteristics, the aerodynamic characteristics in yaw, and the tuft studies for 0^o and 3.7^o yaw. The results of the effect of leading-edge and trailing-edge flaps on the aerodynamic characteristics of the wing will be presented in later reports."
Altitude-Wind-Tunnel Investigation of a 4000-Pound-Thrust Axial-Flow Turbojet Engine 5 - Analysis of Turbine Performance
"Performance characteristics of the turbine of a 4000-pound-thrust axial-flow turbojet engine was determined in investigations of the complete engine in the NACA Cleveland altitude wind tunnel. Characteristics are presented as functions of the total-pressure ratio across the turbine and of turbine speed and gas flow corrected to sea-level conditions. Three turbine nozzles of different areas were used to determine the area that gave optimum performance" (p. 1).
Altitude-Wind-Tunnel Investigation of a 4000-Pound-Thrust Axial-Flow Turbojet Engine 6: Combustion-Chamber Performance
"An analysis of the performance of the types A, B, and C combustion chambers of the 4000-pound-thrust axial-flow turbojet engine is presented. The data were obtained from investigations of the complete engine over a range of pressure altitudes from 5000 to 40,000 feet and ram pressure ratios from 1.00 to 1.86. The combustion-chamber pressure losses, the effect of the losses on cycle efficiency, and the combustion efficiency are discussed" (p. 1).
Altitude-Wind-Tunnel Investigation of Performance Characteristics of a J47D Prototype (RX1-1) Turbojet Engine With Fixed-Area Exhaust Nozzle
Report presenting an investigation to determine the overall performance of a prototype model of the J47D (RX1-1) turbojet engine with a fixed-area exhaust nozzle. Data was obtained for a range of engine speeds, altitudes, and Mach numbers. Results regarding the effect of those variables and generalization in terms of pumping characteristics are provided.
Analysis and Experimental Observation of Pressure Losses in Ram-Jet Combustion Chambers
From Introduction: "Some experimental data on flame-holder pressure losses have been presented (reference 1 to 4). A theoretical analysis that assume a sudden enlargement of flow area was made at the NACA Lewis laboratory to determine the effect of flame-holder open area and combustion-chamber-inlet Mach number on the pressure losses across flame holders. The results of this analysis were then compared with experimental data obtained with several different flame-holder designs."
An analysis of a highly compounded two-stroke-cycle compression-ignition engine
This report presents an analysis of a compound engine operating with manifold pressures ranging from 60 to 110 lb/sq in. absolute and discusses the effects of engine limits (peak cylinder pressure and turbine-inlet temperature) and component efficiency.
Analysis of Off-Design Operation of High Mach Number Supersonic Turbojet Engines
From Introduction: "Because of the lack of data on compressor characteristics at 130 or 145 percent of design equivalent rotational speed, several simplifying assumptions were made regarding the compressor characteristics during operation at equivalent rotational speeds above the design value (hereinafter called "compressor overspeeding"). In addition to results based on these assumptions, the effect of deviation from these assumptions is discussed."
Analysis of part-speed operation for high-pressure-ratio multistage axial-flow compressors
From Introduction: "For the analysis reported herein, the same hypothetical 12-stage compressor discussed in reference 6 was chosen, and the effects of inlet-stage stall characteristics and stage interactions were estimated by arbitrary adjustment of the stall characteristics and stage interactions were estimated by arbitrary adjustment of the stalled characteristics of the individual stages."
Analytical study of losses at off-design conditions for a fixed geometry turbine
From Introduction: "The purpose of this report is to present the results of the analytical investigation of the turbine of reference 1 to indicate the extent to which the various turbine losses affect the turbine efficiency over the range of performance."
Application of a Windshield-Display System to the Low-Altitude Bombing Problem
From Introduction: "The design and flight evaluation of an airborne target simulator for use in tracking studies of fighter-type airplanes equipped with optical gunsights have recently been reported (ref. 1). In this equipment the target airplane was represented by a movable dot of light projected on the windshield of the test airplane."
Application of Pulse Techniques to Strain Gages
Memorandum presenting pulse techniques applied to strain gages for increasing the output level and extending the usable range. Bonded and unbonded strain gages which normally operate with exciting potentials between 3.5 and 14 volts operated satisfactorily with 200-volt pulses of 1-microsecond duration and a repetition rate of 350 per second. Results regarding maximum allowable voltages, effective pulse duration, sensitivity, minimum readable signal, effects of cable capacitance and inductance, sensitivity to external noise, and linearity are provided.
A Brief Hydrodynamic Investigation of a 1/24-Scale Model of the DR-77 Seaplane
From Summary: "A limited investigation of a 1/24-scale dynamically similar model of the Navy Bureau of Aeronautics DR-77 design was conducted in Langley tank no. 2 to determine the calm-water take-off and the rough-water landing characteristics of the design with particular regard to the take-off resistance and the landing accelerations. During the take-off tests, resistance, trim, and rise were measured and photographs were taken to study spray. During the landing tests, motion-picture records and normal-acceleration records were obtained."
Buffeting forces on two-dimensional airfoils as affected by thickness and thickness distribution
Report presenting an experimental investigation to determine the effect of thickness and of thickness distribution on the fluctuations of normal-force and pitching-moment coefficients of airfoils. Seven symmetrical airfoils were tested. Two types of pressure pulsations were noted in the study: pulsations from an intermittent build-up and dropping of the pressure peak near the leading edge and pulsations attributable to shock-wave motion and unsteady air flow following the shock wave.
The Calculated and Experimental Incremental Loads and Moments Produced by Split Flaps of Various Spans and Spanwise Locations on a 45 Degrees Sweptback Wing of Aspect Ratio 8
Report presenting the incremental lift and pitching moments produced by 20-percent-chord split flaps of various spans at various spanwise positions on two 45 degree sweptback wings of aspect ratio 8.02 as obtained by pressure-distribution testing. Inboard flaps were found to be far more effective in producing lift than outboard flaps. Results regarding spanwise loading and pitching moment are provided.
Calculated Heats of Formation and Combustion of Boron Compounds (Boron, Hydrogen, Carbon, Silicon)
Memorandum presenting the calculation of the heats of formation and combustion for liquid and gaseous alkyl- and silyl-substituted boron compounds by a semitheoretical method. The results indicated that alkylation and more especially silylation reduce the heat of combustion. Results regarding alkyl- and silyldecaboranes, diborylmethane, diborylethane, and their alkyl derivatives, bipentaboranyl and bideocaboranyl, dipentaboranyl- and didecaboranylalkanes, and heats of combustion of isomers are provided.
Calculations on the forces and moments for an oscillating wing-aileron combination in two-dimensional potential flow at sonic speed
From Summary: "The linearized theory for compressible unsteady flow is used, as suggested in recent contributions to the subject, to obtain the velocity potential and the lift and moment for a thin harmonically oscillating, two-dimensional wing-aileron combination moving at sonic speed. The velocity potential is derived by considering the sonic case as the limit of the linearized supersonic theory. From the velocity potential explicit expressions for the lift and moment are developed for vertical translation and pitching of the wing and rotation of the aileron. The sonic results are compared and found to be consistent with previously obtained subsonic and supersonic results. Several figures are presented showing the variation of lift and moment with reduced frequency and Mach number and the influence of Mach number on some cases of bending-torsion flutter."
Characteristics of six propellers including the high-speed range
This investigation is part of an extensive experimental study that has been carried out at full scale in the NACA 20-foot tunnel, the purpose of which has been to furnish information in regard to the functioning of the propeller-cowling-nacelle unit under all conditions of take-off, climbing, and normal flight. This report presents the results of tests of six propellers in the normal and high-speed flight range and also includes a study of the take-off characteristics.
Comparison of Calculated and Experimental Temperatures and Coolant Pressure Losses for a Cascade of Small Air-Cooled Turbine Rotor Blades
From Summary: "Average spanwise blade temperatures and cooling-air pressure losses through a small (1.4-in, span, 0.7-in, chord) air-cooled turbine blade were calculated and are compared with experimental nonrotating cascade data. Two methods of calculating the blade spanwise metal temperature distributions are presented. The method which considered the effect of the length-to-diameter ratio of the coolant passage on the blade-to-coolant heat-transfer coefficient and assumed constant coolant properties based on the coolant bulk temperature gave the best agreement with experimental data."
A comparison of fuel sprays from several types of injection nozzles
This report presents the tests results of a series of tests made of the sprays from 14 fuel injection nozzles of 9 different types, the sprays being injected into air at atmospheric density and at 6 and 14 times atmospheric density. High-speed spark photographs of the sprays from each nozzle at each air density were taken at the rate of 2,000 per second, and from them were obtained the dimensions of the sprays and the rates of spray-tip penetration. The sprays were also injected against plasticine targets placed at different distances from the nozzles, and the impressions made in the plasticine were used as an indication of the distribution of the fuel within the spray. Cross-sectional sketches of the different types of sprays are given showing the relative sizes of the spray cores and envelopes. The characteristics of the sprays are compared and discussed with respect to their application to various types of engines.
Comparison of Injectors With a 200-Pound-Thrust Ammonia-Oxygen Engine
"Characteristic exhaust velocity was measured for a small range of mixture ratios with four different injectors. Performances of parallel-sheet, like-on-like, and triplet injectors were about the same, but a parallel-jet injector had a much lower performance. Performance values for ammonia-oxygen were slightly lower than for heptane-oxygen" (p. 1).
A Comparison of Several Systems of Boundary-Layer Removal Ahead of a Typical Conical External-Compression Side Inlet at Mach Numbers of 1.88 and 2.93
Report presenting an investigation at Mach numbers 1.88 and 2.93 to determine the performance characteristics of a conical external-compression side inlet model with a swept-leading-edge boundary-layer-removal scoop. Two other boundary-layer-removal systems were also investigated, which used a deflection wedge and cowl-lip scoops. Results regarding swept-scoop inlets, scoop performance, alternative boundary-layer-removal systems at Mach 1.88, and a comparison of the boundary layer removal systems are provided.
A comparison of the aerodynamic characteristics at transonic speeds of four wing-fuselage configurations as determined from different test techniques
Report presenting a comparison of the high-speed aerodynamic characteristics of a family of four wing-fuselage configurations with four different degrees of sweepback as determined from transonic-bump model tests in the 7- by 10-foot tunnel, sting-supported model tests in the 8-foot tunnel, and the 7-foot by 10-foot tunnel.
Compressible Laminar Boundary Layer Over a Yawed Infinite Cylinder With Heat Transfer and Arbitrary Prandtl Number
"The equations are presented for the development of the compressible laminar boundary layer over a yawed infinite cylinder. For compressible flow with a pressure gradient the chordwise and spanwise flows are not independent. Using the Stewartson transformation and a linear viscosity-temperature relation yields a set of three simultaneous ordinary differential equations in a form yielding similar solutions. These equations are solved for stagnation-line flow for surface temperatures from zero to twice the free-stream stagnation temperature and for a wide range of yaw angle and free-stream Mach number" (p. 1017).
Cooling characteristics of a 2-row radial engine
This report presents the results of cooling tests conducted on a calibrated GR-1535 Pratt and Whitney Wasp, Jr. Engine installed in a Vought X04U-2 airplane. The tests were made in the NACA full-scale tunnel at air speeds from 70 to 120 miles per hour, at engine speeds from 1,500 to 2,600 r.p.m., and at manifold pressures from 19 to 33 inches of mercury absolute. A Smith controllable propeller was used to facilitate obtaining the different combinations of engine speed, power, and manifold pressure.
Correlation of the Characteristics of Single-Cylinder and Flight Engines in Tests of High-Performance Fuels in an Air-Cooled Engine 1 - Cooling Characteristics
Variable charge-air flow, cooling-air pressure drop, and fuel-air ration investigations were conducted to determine the cooling characteristics of a full-scale air-cooled single cylinder on a CUE setup. The data are compared with similar data that were available for the same model multicylinder engine tested in flight in a four-engine airplane. The cylinder-head cooling correlations were the same for both the single-cylinder and the flight engine. The cooling correlations for the barrels differed slightly in that the barrel of the single-cylinder engine runs cooler than the barrel of te flight engine for the same head temperatures and engine conditions.
A description of the design of highly swept propeller blades
"A description of the two swept propellers investigated in the Langley 8-foot high-speed tunnel is presented, together with the discussions of the numerous assumptions and analyses on which the designs of these propellers are based. The blades are swept considerably along the entire blade radius and, in order to allow for reductions in the maximum stresses, are swept forward inboard and backward outboard. The blades have been designed on the basis of the blade-element method primarily to have subcritical efficiencies at the highest possible forward speed. The designs have been controlled primarily by the stresses in the blades" (p. 1).
Design and Experimental Performance of a 0.35 Hub-Tip Radius Ratio Transonic Axial-Flow-Compressor Rotor Designed for 40 Pounds Per Second Per Unit Frontal Area
Memorandum presenting an investigation to determine the feasibility of a high-performance transonic axial-flow compressor stage with a weight flow of 40 pounds per second per square foot of frontal area. A transonic axial-flow inlet stage with a hub-tip ratio of 0.35 and an axial Mach number of approximately 0.75 was designed and fabricated. Results regarding overall rotor performance, flow parameters, radial matching of blade-element sections, and comparison of blade-element parameters with design rules are provided.
Direct method of design and stress analysis of rotating disks with temperature gradient
A method is presented for the determination of the contour of disks, typified by those of aircraft gas turbines, to incorporate arbitrary elastic-stress distributions resulting from either centrifugal or combined centrifugal and thermal effects. The specified stress may be radial, tangential, or any combination of the two. Use is made of the finite-difference approach in solving the stress equations, the amount of computation necessary in the evolution of a design being greatly reduced by the judicious selection of point stations by the aid of a design chart. Use of the charts and of a preselected schedule of point stations is also applied to the direct problem of finding the elastic and plastic stress distribution in disks of a given design, thereby effecting a great reduction in the amount of calculation. Illustrative examples are presented to show computational procedures in the determination of a new design and in analyzing an existing design for elastic stress and for stresses resulting from plastic flow.
Effect of a hot-jet exhaust on pressure distributions and external drag of several afterbodies on a single-engine airplane model at transonic speeds
Report presenting an investigation of the jet effects on several afterbody shapes of a single-engine fighter-airplane model in the 16-foot transonic tunnel at a range of Mach numbers and angles of attack. The afterbody-geometry variables were boattail angle, afterbody length, and base area. Results regarding evaluation of support interference, afterbody-pressure-distribution characteristics, and drag characteristics are provided.
The effect of accelerating a hypothetical aircraft through the transonic range with controls fixed
Memorandum presenting Mach number histories of the motion experienced by a hypothetical, small, straight-wing aircraft accelerating at various rates through an assumed controls-fixed pitch-down balance change in the transonic range. Two approximate analytical solutions of the longitudinal equations of motion are developed which are based on certain simplifying assumptions indicated by the differential-analyzer results.
Effect of an Adjustable Supersonic Inlet on the Performance Up to Mach Number 2.0 of a J34 Turbojet Engine
"A J34 turbojet engine was investigated at free-stream Mach numbers of 0.12 and 1.6 to 2.0 to determine the effect of supersonic inlet operation on engine performance. With the exception of ideal jet thrust, the use of generalized engine parameters correlated the engine data satisfactorily when the exit nozzle was choked. Large total-pressure distortions did not affect compressor efficiency" (p. 1).
Effect of an All-Movable Wing-Tip Control on the Longitudinal Stability of 60 Degree Sweptback-Wing-Indented-Body Configuration Equipped With Fences at Transonic Speeds
Report presenting an investigation to obtain the effects of a 20-percent-semispan all-movable wing-tip control on the longitudinal stability characteristics of a twisted and cambered 60 degree sweptback-wing-indented-body configuration. Testing occurred over a range of angles of attack and Mach numbers. Results regarding the effects on drag coefficient and lift-drag ratio are also provided.
The Effect of Blade-Section Camber on the Static Characteristics of Three NACA Propellers
Report discussing static testing on 3 NACA two-blade propellers with design lift coefficients of 0, 0.3, and 0.5. Information about the effect of camber on the static-thrust figure of merit, the effect of camber on the stall-flutter speed, thrust coefficient, and blade flutter is provided.
The Effect of Boundary-Layer Control by Suction and Several High-Lift Devices on the Longitudinal Aerodynamic Characteristics of a 47.5 Degree Sweptback Wing-Fuselage Combination
"An investigation has been made in the Langley full-scale tunnel of a 47.5 degree sweptback wing-fuselage combination equipped for boundary-layer control by suction. The wing aspect ratio was 3.5, the taper ratio was 0.5, and the airfoil sections normal to the quarter-chord line were NACA 64(sub 1)-A112. The wing configurations tested included the wing with various combinations of extensible leading-edge and split flaps" (p. 1).
Effect of First-Stage Blade Design on Performance of Mark 25 Torpedo Power Plant
"The effect of rotor-blade length, inlet angle, and shrouding was investigated with four different nozzles in a single-stage modification of the Mark 25 aerial-torpedo power plant. The results obtained with the five special rotor configurations are compared with those of the standard first-stage rotor with each nozzle. Each nozzle-rotor combination was operated at nominal pressure ratios of 8, 15 (design), and 20 over a range of speeds from 6000 rpm to the design speed of 18,000 rpm" (p. 1).
The effect of several jet-engine air-inlet configurations on the low-speed static lateral stability characteristics of a 1/6-scale model of the MX-1764 airplane
Report presenting an investigation in the 300 mph 7- by 10-foot tunnel to determine the effect of wing-root leading-edge- and scoop-type jet-engine air-inlet configurations on the static lateral stability characteristics of a scale model of the MX-1764 airplane. The inlet configurations generally had only small effects on the lateral stability.
The Effect of Several Jet-Engine Air-Inlet Configurations on the Low-Speed Static Lateral Stability Characteristics of a 1/6- Scale of the MX-1764 Airplane
Memorandum presenting an investigation in the 7- by 10-foot tunnel to determine the effect of wing-root leading-edge- and scoop-type jet-engine air-inlet configurations on the static lateral stability characteristics of a scale model of the MX-1764 airplane. The addition of the inlet configurations to the model generally had only small effects on the lateral stability.
Effect of Turbine Axial Nozzle-Wheel Clearance on Performance of Mark 25 Torpedo Power Plant
"Investigations were made of the turbine from a Mark 25 torpedo to determine the performance of the unit with three different turbine nozzles at various axial nozzle-wheel clearances. Turbine efficiency with a reamed nondivergent nozzle that uses the axial clearance space for gas expansion was little affected by increasing the axial running clearance from 0.030 to 0.150 inch. Turbine efficiency with cast nozzles that expanded the gas inside the nozzle passage was found to be sensitive to increased axial nozzle-wheel clearance" (p. 1).
The effect of valve timing upon the performance of a supercharged engine at altitude and an unsupercharged engine at sea level
This investigation was conducted to determine the comparative effects of valve timing on the performance of an unsupercharged engine at sea level and a supercharged engine at altitude. The tests were conducted on the NACA universal test engine. The timing of the four valve events was varied over a wide range; the engine speeds were varied between 1,050 and 1,500 r.p.m.; the compression ratios were varied between 4.35:1 and 7.35:1. The conditions of exhaust pressure and carburetor pressure of a supercharged engine were simulated for altitudes between 0 and 18,000 feet. The results show that optimum valve timing for a supercharged engine at an altitude of 18,000 feet differs slightly from that for an unsupercharged engine at sea level. A small increase in power is obtained by using the optimum timing for 18,000 feet for altitudes above 5,000 feet. The timing of the intake opening and exhaust closing becomes more critical as the compression ratio is increased.
Effect of wing-tank location on the drag and trim of a swept-wing model as measured in flight at transonic speeds
Report presenting results of an exploratory free-flight investigation at zero lift of several rocket-powered drag research models equipped wing wing tanks at a range of Mach numbers. The tanks, which were slender bodies of revolution, were mounted on 34 degrees sweptback, nontapered wings of 2.7 aspect ratio. Results regarding drag and trim change are provided.
Effects of a J34 Turbojet Engine on Supersonic Diffuser Performance
Report presenting testing of a translating cone inlet with a variable bypass at Mach numbers 1.6, 1.8, and 2.0 with both a choked exit plug and a J34 turbojet engine. The main difference between the two options was increased inlet subcritical stability with the engine. Results regarding basic diffuser performance, inlet stability, buzz amplitude and frequency, and diffuser-exit profiles are provided.
The effects of a modified roll-command system on the flight-path stability and tracking accuracy of an automatic interceptor
From Summary: "In order to improve the tracking characteristics of an automatic interceptor, a revised roll-command computer was tested both on an analog computer and in flight. This report presents flight-test results which indicate a significant improvement in flight-path stability and tracking accuracy. The modified roll computer was designed on the basis of previous analytical studies."
Effects of Combinations of Aspect Ratio and Sweepback at High Subsonic Mach Numbers
Report discussing an investigation to determine the effects of sweepback and low aspect ratio on the aerodynamic characteristics of a wing at high subsonic Mach numbers. Tests were performed at aspect ratios of 2, 3, and 5 and sweepback angles of 0, 30, and 45 degrees. Generally, sweepback and low aspect ratio were found to both delay and lessen the effects of compressibility.
Effects of increasing Reynolds number from 2 x 10(exp 6) to 6 x 10(exp 6) on the aerodynamic characteristics at transonic speeds of a 45 degree swept wing with 6 degree leading-edge droop
Report presenting an investigation in the 16-foot and 8-foot transonic tunnel to determine the effects of Reynolds number and on a swept wing with camber. The wing had 45 degrees sweepback of the quarter-chord line, an aspect ratio of 4, a taper ratio of 0.6, and NACA 65A006 airfoil sections parallel to the plane of symmetry. Results regarding the effect of Reynolds number on the aerodynamic characteristics and effects of roughness strips are provided.
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