From Introduction: "This report describes some measurements of these quantities obtained by crashing fighter aircraft under circumstances approximating those observed in service."
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.
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.
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.
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."
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."
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."
"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).
"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).
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.
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."
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.
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."
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."
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."
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."
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.
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."
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.
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.
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.
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."
"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).
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.
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.
"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).
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.
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.
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.
"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).
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.
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.
"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).
"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).
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.
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.
"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).
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.
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.
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."
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.
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.
Memorandum presenting an investigation of the effect of the addition of approximately 6 percent nickel, cobalt, or platinum on some properties of molybdenum disilicide. These additions resulted in appreciably lowering the modulus-of-rupture strength from that of unalloyed molybdenum disilicide. Results regarding density and particle-size analysis, chemical analysis of bars, metallographic and x-ray analysis, density and resistivity, modulus-of-rupture strengths, thermal shock tests, and oxidation resistance are provided.
Report presenting the effect of control effectiveness of systematically varying the size and location of trailing-edge flaps on a 45 degree sweptback wing at a Mach number of 1.9. The wing model had an aspect ratio of 2.5, a taper ratio of 0.625, and 6-percent-thick hexagonal airfoil sections. Results regarding the rolling-moment and yawing-moment coefficients, lift, rolling-moment, and pitching-moment effectiveness parameters, and flap characteristics are provided.
Report presenting an investigation of a silicone-diester blend (SD-17) in various bench studies and in a turbopropeller engine to determine its suitability as a lubricant for aircraft engines. The performance of the fluid was satisfactory during more than 17 hours of operation in a T-38 engine with power levels from 1745 to 2400 horsepower. Results regarding bench studies and turbopropeller engine study are provided.
"The static aeroelastic divergence characteristics of a delta-plan-form model of the canard control surface of a proposed air-to-ground missile have been studied both analytically and experimentally in the Mach number range from 0.6 to 3.0. The experiments indicated that divergence occurred at a nearly constant value of dynamic pressure at Mach numbers up to 1.2. At higher Mach numbers somewhat higher values of dynamic pressure were required to produce divergence" (p. 1).
Report presenting measurements made at a Mach number of 0.25 of the aerodynamic characteristics of a wing with an aspect ratio of 3 combined separately with three geometrically similar bodies of revolution with the same fineness ratio but differing sizes. Results regarding wings, bodies of revolution, wing in the presence of the bodies, wing combined with the bodies of revolution, and effect of wing incidence changes are provided.
Report presenting an investigation to develop screech instrumentation and study the mechanism of screech in a simulated afterburner. Results regarding the screech mechanism, effect of system variables on screech, speculation regarding screech mechanism, and suggested techniques for elimination or control of screech are provided.
Report presenting zero-lift experimental flutter data in a range of Mach numbers on a 1.5-percent-thick, untapered, unswept, solid steel wing with hexagonal airfoil sections and wing length-to-chord ratios of 1.61 and 1.73. The wing was tested as a cantilever with and without a half-body-of-revolution fuselage. Results regarding the calculated flutter speeds and frequencies, effect of fuselage, variation of wing flutter characteristics with Mach number, results of analytic solutions, and variation of wing flutter characteristics with mass ratio are provided.
Report presenting the propulsive and aerodynamic characteristics of a flat ramjet engine suitable for use on a helicopter rotor and compared with previous tests of an equivalent engine with a circular cross section. The results indicate that the flat engine has higher values of propulsive thrust plus power-off drag than the circular engine. The power-off and drag characteristics indicate that the flat engine has lift-drag ratios about 3 times those obtained with the circular engine.
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