Wind tunnel calibration of the ''Arcasonde 1-A'' at simulated altitudes between 35 and 57 km
Subsonic wind tunnel calibration of Arcasonde 1-A atmospheric temperature sensing unit at simulated altitudes between 35 and 56 km
Haak, E. L.↗
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Results for “ALTITUDE SIMULATION”
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Spectral radiances of model rocket exhaust gases at simulated altitude measured with a rapid-scanning spectrometer.
Spectral radiance of model rocket exhaust gases measured by rapid scanning spectrometer at simulated altitude
EXHAUST GAS↗
Performance of a UTC FW-4S solid propellant rocket motor under the command effects of simulated altitude and rotational spin
One United Technology Center FW-4S solid-propellant rocket motor was fired at an average simulated altitude of 103,000 ft while spinning about its axial centerline at 180 rpm. The objectives of the test program were to determine motor altitude ballistic performance including the measurement of the nonaxial thrust vector and to demonstrate structural integrity of the motor case and nozzle. These objectives are presented and discussed.
Merryman, H. L.↗
Spectral radiance measurements of rocket model exhaust gases, at simulated altitude, using a high speed scanning spectrometer Final report, part 2
Spectral radiance measurements of scale model rocket engine exhaust gases at simulated altitude using high speed scanning spectrometer
SCANNING DEVICE↗
Design assurance test of the Thiokol Te-M-521-5 apogee kick motor tested in the spin mode at simulated altitude conditions
One Thiokol Chemical Corporation TE-M-521-5 solid-propellant apogee rocket motor was successfully fired at an average simulated altitude of about 108,000 ft while spinning at 46 rpm. The general program objectives were to verify compliance of motor performance with the manufacturer's specifications. Specific primary objectives were to determine vacuum ballistic performance of the motor after prefire vibration conditioning and temperature conditioning at 40F, altitude ignition characteristics, motor structural integrity, and motor temperature-time history during and after motor operation. Additional objectives were to measure the lateral (nonaxial) thrust component during motor operation and to measure radiation heat flux in the vicinity of the nozzle exit plane.
Cimino, A. A.↗
Ignition Delay Experiments with Small-scale Rocket Engine at Simulated Altitude Conditions Using Various Fuels with Nitric Acid Oxidants / Dezso J. Ladanyi
Ignition delay determinations of several fuels with nitric oxidants were made at simulated altitude conditions utilizing a small-scale rocket engine of approximately 50 pounds thrust. Included in the fuels were aniline, hydrazine hydrate, furfuryl alcohol, furfuryl mercaptan, turpentine, and mixtures of triethylamine with mixed xylidines and diallyaniline. Red fuming, white fuming, and anhydrous nitric acids were used with and without additives. A diallylaniline - triethylamine mixture and a red fuming nitric acid analyzing 3.5 percent water and 16 percent NO2 by weight was found to have a wide temperature-pressure ignition range, yielding average delays from 13 milliseconds at 110 degrees F to 55 milliseconds at -95 degrees F regardless of the initial ambient pressure that ranged from sea-level pressure altitude of 94,000 feet.
XYLIDINE↗
Effect of Fuel on Performance of a Single Combustor of an I-16 Turbojet Engine at Simulated Altitude Conditions
As part of a study of the effects of fuel composition on the combustor performance of a turbojet engine, an investigation was made in a single I-16 combustor with the standard I-16 injection nozzle, supplied by the engine manufacturer, at simulated altitude conditions. The 10 fuels investigated included hydrocarbons of the paraffin olefin, naphthene, and aromatic classes having a boiling range from 113 degrees to 655 degrees F. They were hot-acid octane, diisobutylene, methylcyclohexane, benzene, xylene, 62-octane gasoline, kerosene, solvent 2, and Diesel fuel oil. The fuels were tested at combustor conditions simulating I-16 turbojet operation at an altitude of 45,000 feet and at a rotor speed of 12,200 rpm. At these conditions the combustor-inlet air temperature, static pressure, and velocity were 60 degrees F., 12.3 inches of mercury absolute, and 112 feet per second respectively, and were held approximately constant for the investigation. The reproducibility of the data is shown by check runs taken each day during the investigation. The combustion in the exhaust elbow was visually observed for each fuel investigated.
Zettle, Eugene V↗
Simulated Altitude Performance of Combustors for the 24C Turbojet Engine. III - Performance of Rectangular-Slot Baskets: Performance of Rectangular-Slot Baskets - 3
The performance of an annular combustion chamber from a 24C turbojet engine was investigated over a range of simulated altitudes from 20,000 to 55,000 feet and corrected engine rotor speeds from 6000 to 13,000 rpm at a simulated ram-pressure ratio of 1.04. The purpose of the investigation was to determine the effects on the altitude operational limits, combustor-outlet gas temperature distribution, combustion efficiencies, and combustor inlet-to-outlet total-pressure drops of two changes in the 24C-4B basket air-passage arrangements that were designed to improve combustor-outlet temperature distribution. These changes were: (a) replacement of the downstream secondary air holes with large rectangular slots further upstream (rectangular-slot basket), and (b) enlargement of anticoking holes in the rectangular-slot basket (modified rectangular-slot basket). The results indicate that improved outlet-gas temperature distribution of each succeeding combustor basket investigated was attained at a sacrifice in the altitude limit of operation. The altitude limits of operation of the combustor with the original basket ranged from 34,000 feet at a corrected engine speed of 6000 rpm to a maximum of 52,000 feet at 12 ' 500 rpm. The altitude limits of the rectangular-slot basket were about 2000 feet lower throughout the engine speed range than those of the original basket. The altitude limits of the combustor with the modified rectangular-slot basket were about equivalent to those of the other baskets in the corrected-engine-speed range from 12,000 to 12,500 rpm but were about 10,000 feet lower than those of the original basket in the corrected-engine-speed range from 6000 to 9000 rpm. For the same inlet-air conditions, the combustion efficiencies were highest for the original basket and progressively lower for each of the other two baskets. The combustor inlet-to-outlet pressure drops of all three combustor baskets at the same operating conditions were within +/- 10 percent of the pressure drop of the original basket.
Tischler, Adelbert O.↗
Comparison of Rocket Performance using Exhaust Diffuser and Conventional Techniques for Altitude Simulation
A rocket engine with an exhaust-nozzle area ratio of 25 was operated at a constant chamber pressure of 600 pounds per square inch absolute over a range of oxidant-fuel ratios at an altitude pressure corresponding to approximately 47,000 feet. At this condition, the nozzle flow is slightly underexpanded as it leaves the nozzle. The altitude simulation was obtained first through the use of an exhaust diffuser coupled with the rocket engine and secondly, in an altitude test chamber where separate exhauster equipment provided the altitude pressure. A comparison of performance data from these two tests has established that a diffuser used with a rocket engine operating at near-design nozzle pressure ratio can be a valid means of obtaining altitude performance data for rocket engines.
Sivo, Joseph N.↗
Measurement of gaseous emissions from an afterburning turbojet engine at simulated altitude conditions
Gaseous emissions from a J85-GE-13 turbojet engine were measured over a range of fuel-air ratios from idle to full afterburning and simulated altitudes from near sea-level to 12,800 meters (42,000 ft). Without afterburning, carbon monoxide and unburned hydrocarbon emissions were highest at idle and lowest at takeoff; oxides of nitrogen exhibited the reverse trend. With afterburning, carbon monoxide and unburned hydrocarbon emissions were greater than for military power. Carbon monoxide emissions were altitude dependent. Oxides of nitrogen emissions were less at minimum afterburning than at military power. For power levels above minimum afterburning, the oxides of nitrogen emissions were both power level and altitude dependent.
Diehl, L. A.↗
Performance and Operational Characteristics of a Python Turbine-propeller Engine at Simulated Altitude Conditions / Carl L. Meyer and Lavern A. Johnson
The performance and operational characteristics of a Python turbine propeller engine have been investigated over a range of engine operating conditions at simulated altitudes from 5000 to 40,000 feet in the NACA Lewis altitude wind tunnel . For the performance phase of the investigation, a single cowl-inlet ram pressure ratio was maintained and independent control of engine speed and fuel flow was used to obtain a range of power at each engine speed. Engine performance data obtained at a given altitude could not be used to predict performance accurately at other altitudes by use of the standard air pressure and temperature generalizing factors. Specific fuel consumption based on shaft horsepower decreased as the shaft horsepower increased at a given engine speed. At a given altitude condition and turbine -inlet total temperature, there was an optimum engine speed at which maximum shaft horsepower was obtained. At a given engine speed and turbine-inlet total temperature, a greater portion of the total available energy was converted to propulsive power as the altitude increased. Windmilling starts were made at altitudes up to 40,000 feet; the minimum true airspeed required for successful windmilling starts increased as the altitude was raised. In general, the engine control prevented large overshoots or undershoots of engine speed during starts, accelerations, and decelerations.
Engines - Armstrong-Siddely Python Turbo-Propeller↗
Characterization of a Pressure-Fed LOX/LCH4 Reaction Control System Under Simulated Altitude and Thermal Vacuum Conditions
A liquid oxygen, liquid methane (LOX/LCH4) reaction control system (RCS) was tested at NASA Glenn Research Center's Plum Brook Station in the Spacecraft Propulsion Research Facility (B-2) under simulated altitude and thermal vacuum conditions. The RCS is a subsystem of the Integrated Cryogenic Propulsion Test Article (ICPTA) and was initially developed under Project Morpheus. Composed of two 28 lbf-thrust and two 7 lbf-thrust engines, the RCS is fed in parallel with the ICPTA main engine from four propellant tanks. 40 tests consisting of 1,010 individual thruster pulses were performed across 6 different test days. Major test objectives were focused on system dynamics, and included characterization of fluid transients, manifold priming, manifold thermal conditioning, thermodynamic vent system (TVS) performance, and main engine/RCS interaction. Peak surge pressures from valve opening and closing events were examined. It was determined that these events were impacted significantly by vapor cavity formation and collapse. In most cases the valve opening transient was more severe than the valve closing. Under thermal vacuum conditions it was shown that TVS operation is unnecessary to maintain liquid conditions at the thruster inlets. However, under higher heat leak environments the RCS can still be operated in a self-conditioning mode without overboard TVS venting, contingent upon the engines managing a range of potentially severe thermal transients. Lastly, during testing under cold thermal conditions the engines experienced significant ignition problems. Only after warming the thruster bodies with a gaseous nitrogen purge to an intermediate temperature was successful ignition demonstrated.
Atwell, Matthew J.↗
Experimental Evaluation of Rocket Exhaust Diffusers for Altitude Simulation
An experimental investigation of exhaust diffusers has been conducted to evaluate various methods of minimizing the overall pressure ratio (from chamber to ambient pressure) required to establish and maintain full expansion of the nozzle flow (altitude simulation). Exhaust-diffuser configurations investigated were (1) cylindrical diffusers, (2) diffusers with contraction, and (3) diffusers including a right-angle turn. Cylindrical diffusers were evaluated with primary nozzles of various area ratios and types, as well as two clustered configurations; the other diffusers were evaluated with individual nozzles of constant area ratio and varied type. Air was the working fluid, except for two check points obtained with JP-4 fuel and liquid-oxygen rocket engines and cylindrical diffusers. The minimum length-diameter ratio of cylindrical diffusers was about 6 for minimum pressure-ratio requirements. With cylindrical diffusers of adequate length, the pressure-ratio requirements were primarily a function of the ratio of diffuser to nozzle-throat areas and were essentially independent of primary-nozzle type (including two clustered configurations) or area ratio. The two check points obtained with rocket engines indicated the pressure-ratio requirements at given ratios of diffuser to nozzle-throat areas were lowered, as compared with the requirements with air, as a result of the reduced ratio of specific heats. The minimum length-diameter ratio of the contraction throat of convergent-divergent diffusers was also about 6 for minimum pressure-ratio requirements. With adequate contraction-throat length, the pressure-ratio requirements of such diffusers were appreciably below those of comparable cylindrical diffusers when used with conical and cutoff-isentropic nozzles, but not when used with a bell nozzle. Minimum pressure-ratio requirements of a diffuser including a simple long-radius right-angle turn at maximum diffuser area, obtained with the center of radius of the turn a minimum of 2 diffuser diameters downstream of the nozzle exit, were not appreciably above those of a comparable optimum cylindrical diffuser. A diffuser including a long-radius right-angle turn at a contraction minimum area had somewhat lower pressure-ratio requirements than the aforementioned simple turn.
Sivo, Joseph N.↗
Preliminary investigation of the combustion of a 50 percent pentaborane - 50 percent JP-4 fuel blend in a turbojet combustor at simulated altitude conditions
A preliminary investigation was conducted to determine the combustion characteristics of a fuel composed of 50 percent pentaborane and 50 percent JP-4 (MIL-F-5624A) by weight in a turbojet combustor. A combustor designed to fit the housing of a J33-A-23 turbojet engine was selected for convenience. The fuel was evaluated at two engine conditions simulating altitudes of 40,000 and 57,000 feet, an engine speed of 85 percent of rated rpm, and a flight Mach number of 0.6. The pentaborane blend was initially evaluated in combustors developed for pure pentaborane and diborane reported in NACA RM E53B18 and RM E52L15. The performance of the blend was unsatisfactory in these combustors. A new combustor was then developed which provided combustor efficiencies measured from 91 to 101 percent as compared with efficiencies of 92 to 94 percent previously obtained for pentaborane at comparable conditions. Additional refinements of design details are needed to obtain lower oxide deposits and a more uniform outlet temperature profile; however, the combustor is believed to incorporate some of the design principles required to obtain satisfactory over-all performance with the fuel blend investigated.
FUELS - JP-4-*PENTABORANE↗
Simulated Altitude Performance of Combustor of Westinghouse 19XB-1 Jet-Propulsion Engine
A 19XB-1 combustor was operated under conditions simulating zero-ram operation of the 19XB-1 turbojet engine at various altitudes and engine speeds. The combustion efficiencies and the altitude operational limits were determined; data were also obtained on the character of the combustion, the pressure drop through the combustor, and the combustor-outlet temperature and velocity profiles. At altitudes about 10,000 feet below the operational limits, the flames were yellow and steady and the temperature rise through the combustor increased with fuel-air ratio throughout the range of fuel-air ratios investigated. At altitudes near the operational limits, the flames were blue and flickering and the combustor was sluggish in its response to changes in fuel flow. At these high altitudes, the temperature rise through the combustor increased very slowly as the fuel flow was increased and attained a maximum at a fuel-air ratio much leaner than the over-all stoichiometric; further increases in fuel flow resulted in decreased values of combustor temperature rise and increased resonance until a rich-limit blow-out occurred. The approximate operational ceiling of the engine as determined by the combustor, using AN-F-28, Amendment-3, fuel, was 30,400 feet at a simulated engine speed of 7500 rpm and increased as the engine speed was increased. At an engine speed of 16,000 rpm, the operational ceiling was approximately 48,000 feet. Throughout the range of simulated altitudes and engine speeds investigated, the combustion efficiency increased with increasing engine speed and with decreasing altitude. The combustion efficiency varied from over 99 percent at operating conditions simulating high engine speed and low altitude operation to less than 50 percent at conditions simulating operation at altitudes near the operational limits. The isothermal total pressure drop through the combustor was 1.82 times as great as the inlet dynamic pressure. As expected from theoretical considerations, a straight-line correlation was obtained when the ratio of the combustor total pressure drop to the combustor-inlet dynamic pressure was plotted as a function of the ratio of the combustor-inlet air density to the combustor-outlet gas density. The combustor-outlet temperature profiles were, in general, more uniform for runs in which the temperature rise was low and the combustion efficiency was high. Inspection of the combustor basket after 36 hours of operation showed very little deterioration and no appreciable carbon deposits.
Childs, J. Howard↗
SCALE MODEL TESTING OF 90 DEG TURN EJECTOR SYSTEMS FOR ROCKET ENGINE ALTITUDE SIMULATION
Scale model testing of 90-deg turn ejector systems for rocket engine simulation
ROCKET ENGINE↗
Exhaust emissions survey of a turbofan engine for flame holder swirl type augmentors at simulated altitude flight conditions
Emissions of carbon dioxide, total oxides of nitrogen, unburned hydrocarbons, and carbon monoxide from an F100 afterburning two spool turbofan engine at simulated flight conditions are reported. Tests were run at Mach 0.8 at altitudes of 10.97 and 13.71 km (36,000 and 45,000 ft), and at Mach 1.2 at 13.71 km (45,000 ft). Emission measurements were made from intermediate power (nonafterburning) through maximum afterburning, using a single point gas sample probe traversed across the horizontal diameter of the exhaust nozzle. The data show that emissions vary with flight speed, altitude, power level, and radial position across the nozzle. Carbon monoxide emissions were low for intermediate and partial afterburning power. Unburned hydrocarbons were near zero for most of the simulated flight conditions. At maximum afterburning, there were regions of NOx deficiency in regions of high CO. The results suggest that the low NOx levels observed in the tests are a result of interaction with high CO in the thermal converter. CO2 emissions were proportional to local fuel air ratio for all test conditions.
Moss, J. E., Jr.↗
Simulated Altitude Performance of Combustors for the Westinghouse 24C Jet Engine I-24C-2 Combustor
A Westinghouse 24C-2 combustor was investigated at conditions simulating operation of the 24C Jet engine at zero ram over ranges of altitude and engine speed. The investigation was conducted to determine the altitude operational limits, that is, the maximum altitude for various engine speeds at which an average combustor-outlet gas temperature sufficient for operation of the jet engine could be obtained. Information was also obtained regarding the character of the flames, the combustion efficiency, the combustor-outlet gas temperature and velocity distributions, the extent of afterburning, the flow characteristics of the fuel manifolds, the combustor inlet-to-outlet total-pressure drop, and the durability of the combustor basket. The results of the investigation indicated that the altitude operational limits for zero ram decreased from 12,000 feet at an engine speed of 4000 rpm to a minimum of 9000 feet at 6000 rpm, and thence increased to 49,000 feet at 12,000 rpm.. At altitudes below the operational limits, flames were essentially steady, but, at altitudes above the operational limits, flames were often cycling and either blew out or caused violent explosions and vibrations. At conditions on the altitude operational limits the type of combustion varied from steady to cycling with increasing fuel-air ratio and the reverse occurred with decreasing fuel-air ratio. In the region of operation investigated, the combustion efficiency ranged from 75 to 95 percent at altitudes below the operational limits and dropped to 55 percent or less at some altitudes above the operational limits. The deviations in the local combustor-outlet gas temperatures were within +20 to -30 percent of the mean combustor temperature rise for inlet-air temperatures at the low end of the range investigated, but became more uneven (up to +/-100 percent) with increasing inlet-air temperatures. The distribution of the combustor-outlet gas velocity followed a similar trend. Practically no afterburning downstream of the combustor outlet occurred. At conditions of high inlet-air temperature several factors indicated that fuel vapor or air formed in the fuel manifolds and adversely affected combustion. The combustor inlet-to-outlet total-pressure drop can be correlated as a function of the ratio of the combustion-air inlet density to outlet density and of the inlet dynamic pressure. The walls of the combustor basket were warped and burned during 29 hours of operation.
Manganiello, Eugene J.↗