ALBERT

Beschreibung: An investigation to increase the compressor surge-limit pressure ratio of the XJ40-WE-6 turbojet engine at high equivalent speeds was conducted at the NACA Lewis altitude wind tunnel. This report evaluates the compressor modifications which were restricted to (1) twisting rotor blades (in place) to change blade section angles and (2) inserting new stator diaphragms with different blade angles. Such configuration changes could be incorporated quickly and easily in existing engines at overhaul depots. It was found that slight improvements in the compressor surge limit were possible by compressor blade adjustment. However, some of the modifications also reduced the engine air flow and hence penalized the thrust. The use of a mixer assembly at the compressor outlet improved the surge limit with no appreciable thrust penalty.

Beschreibung: An investigation was conducted at simulated high-altitude flight conditions to evaluate the use of compressor evaporative cooling as a means of turbojet-engine thrust augmentation. Comparison of the performance of the engine with water-alcohol injection at the compressor inlet, at the sixth stage of the compressor, and at the sixth and ninth stages was made. From consideration of the thrust increases achieved, the interstage injection of the coolant was considered more desirable preferred over the combined sixth- and ninth-stage injection because of its relative simplicity. A maximum augmented net-thrust ratio of 1.106 and a maximum augmented jet-thrust ratio of 1.062 were obtained at an augmented liquid ratio of 2.98 and an engine-inlet temperature of 80 F. At lower inlet temperatures (-40 to 40 F), the maximum augmented net-thrust ratios ranged from 1.040 to 1.076 and the maximum augmented jet-thrust ratios ranged from 1.027 to 1.048, depending upon the inlet temperature. The relatively small increase in performance at the lower inlet-air temperatures can be partially attributed to the inadequate evaporation of the water-alcohol mixture, but the more significant limitation was believed to be caused by the negative influence of the liquid coolant on engine- component performance. In general, it is concluded that the effectiveness of the injection of a coolant into the compressor as a means of thrust augmentation is considerably influenced by the design characteristics of the components of the engine being used.

Beschreibung: An investigation was made of the performance of nine conical cooling-air ejectors at primary jet pressure ratios from 1 to 10, secondary pressure ratios to 4.0, and a temperature ratio of unity. This phase of the investigation was limited to conical ejectors having shroud exit to primary nozzle exit diameter ratios of 1.06 and 1.40, with several spacing ratios for each. The experimental results indicated that the pumping range and amount of cooling-air flow obtained with a 1.06 diameter ratio ejector were relatively small for cooling purposes but that the maximum possible thrust loss, which occurred with no secondary flow, was only 7 percent of convergent nozzle thrust. The 1.40 diameter ratio ejector produced a large cooling air flow and showed a possible thrust loss of 29.5 percent with no cooling air flow. Thrust gains were attained with ejectors of both diameter ratios at secondary pressure ratios greater than 1.0. The limiting primary pressure ratio below which an ejector can operate at a specific secondary pressure ratio (cut-off point) may be estimated for various flight conditions from data contained herein.

Beschreibung: The stator-blade angles in the twelfth to fifteenth stages of a 16-stage high-pressure-ratio axial-flow compressor were decreased 3 deg The over-all performance of this compressor is compared with the performance of the same compressor with standard blade angles. The matching characteristics of the modified compressor and a two-stage turbine were also obtained and compared with those of the compressor with the original blade angles and the same turbine.

Beschreibung: A theoretical method for evaluating the stability characteristics and the amplitude and the frequency of pulsation of ram-jet engines without heat addition is presented herein. Experimental verification of the theoretical results are included where data were available. Theory and experiment show that the pulsation amplitude of a high mass-flow-ratio diffuser having no cone surface flow separation increases with decreasing mass flow. The theoretical trends for changes in amplitude, frequency, and mean-pressure recovery with changes in plenum-chamber volume were experimentally confirmed. For perforated convergent-divergent-type diffusers, a stability hysteresis loop was predicted on the pressure-recovery mass-flow-ratio curve. At a given mean mass-flow ratio, the higher.value of mean pressure recovery corresponded to oscillatory flow in the diffuser while the lower branch was stable. This hysteresis has been observed experimentally. The theory indicates that for a ram-jet engine of given diameter, the amplitude of pulsation of a supersonic diffuser is increased by decreasing the relative size of the plenum chamber with respect to the diffuser volume down to a critical value at which oscillations cease. In the region of these critical values, the stable mass-flow range of the diffuser may be increased either by decreasing the combustion chamber volume or by increasing the length of the diffuser.

Beschreibung: An investigation of the effect of inlet pressure, corrected engine speed, and turbine temperature level on turbine-inlet gas temperature distributions was conducted on a J40-WE-6, interim J40-WE-6, and prototype J40-WE-8 turbojet engine in the altitude wind tunnel at the NAC.4 Lewis laboratory. The engines were investigated over a range of simulated pressure altitudes from 15,000 to 55,000 feet, flight Mach numbers from 0.12 to 0.64, and corrected engine speeds from 7198 to 8026 rpm, The gas temperature distribution at the turbine of the three engines over the range of operating conditions investigated was considered satisfactory from the standpoint of desired temperature distribution with one exception - the distribution for the J40-WE-6 engine indicated a trend with decreasing engine-inlet pressure for the temperature to exceed the desired in the region of the blade hub. Installation of a compressor-outlet mixer vane assembly remedied this undesirable temperature distribution, The experimental data have shown that turbine-inlet temperature distributions are influenced in the expected manner by changes in compressor-outlet pressure or mass-flow distribution and by changes in combustor hole-area distribution. The similarity between turbine-inlet and turbine-outlet temperature distribution indicated only a small shift in temperature distribution imposed by the turbine rotors. The attainable jet thrusts of the three engines were influenced in different degrees and directions by changes in temperature distributions with change in engine-inlet pressure. Inability to match the desired temperature distribution resulted, for the J40-WE-6 engine, in an 11-percent thrust loss based on an average turbine-inlet temperature of 1500 F at an engine-inlet pressure of 500 pounds per square foot absolute. Departure from the desired temperature distribution in the Slade tip region results, for the prototype J40-WE-8 engine, in an attainable thrust increase of 3 to 4 percent as compared with that obtained if tip-region temperature limitations were observed.

Beschreibung: An altitude-test-chamber investigation was conducted to determine the operational and performance characteristics of a McDonnell afterburner with a fixed-area exhaust nozzle on a J34 engine. At rated engine speed, the altitude limit, as determined by combustion blow-out, occurred as a band of unstable operation of about 6000-foot altitude in width with minimum altitude limits from 31,000 feet at a simulated flight Mach number of 0.40 to about 45,500 feet at a simulated flight Mach number of 1.00. Considerable difficulty was experienced in attempting to establish or maintain balanced-cycle engine operation at altitudes above 36,000 feet. The fuel-air ratio for balanced-cycle operation and lean blowout of the afterburner, the augmented-thrust ratio, the total specific fuel consumption, and the afterburner combustion efficiency for balanced-cycle operation are summarized in a table. Satisfactory afterburner ignition was obtained over a range of flight Mach Numbers from 0.32 to 0.60 at altitudes from 10,000 to 30,000 and engine speeds from 10,000 to 12,500 rpm.

Beschreibung: A method for calculation of a counterrotating propeller which is similar to Walchner's method for calculation of the single propeller in the free air stream is developed and compared with measurements. Several dimensions which are important for the design are given end simple formulas for the gain in efficiency derived. Finally a survey of the behavior of the propeller for various operating conditions is presented.

Beschreibung: Results of measurements on a shrouded propeller are given. The propeller is designed for the high ratio of advance and high thrust loading. The effect of the shape of propeller and shroud upon the aerodynamic coefficients of the propulsion unit can be seen from the results. The highest efficiency measured is 0.71. The measurements permit the conclusion that the maximum efficiency can be essentially improved by shroud profiles of small chord and thickness. The largest static thrust factor of merit measured reaches according to Bendemann, a value of about zeta = 1.1. By the use of a nose split flap the static thrust for thin shroud profiles with small nose radius can be about doubled. In a separate section numerical investigations of the behavior of shrouded propellers for the ideal case and for the case with energy losses are carried out. The calculations are based on the assumption that the slipstream cross section depends solely on the shape of the shroud and not on the propeller loading. The reliability of this hypothesis is confirmed experimentally and by flow photographs for a shroud with small circulation. Calculation and test are also in good agreement concerning efficiency and static thrust factor of merit. The prospects of applicability for shrouded propellers and their essential advantages are discussed.

Beschreibung: The requirements on gas turbines for aircraft power units, namely, adequate efficiency, operation at high gas temperatures, low weight, and small dimensions, must be taken into consideration during the design of the blading. To secure good efficiency, it is necessary that the gas flow past the blades as smoothly as possible without separation. This is relatively easily obtainable in the accelerated flow of turbine blading, if the blade spacing is chosen small enough. A small blade spacing, however, is detrimental to the other requirements outlined above. Operation at high gas temperatures usually calls for blade cooling. This cooling is associated with a power input that lowers the turbine efficiency. Since the amount of heat that must be carried off for coding a blade can be influenced rather little, the gross power input for a turbine stage can be reduced by keeping the number of blades to a minimum, that is, with blades of high spacing ratio. But here also a limit is imposed, the exceeding of which is followed by separation of flow. Hence the requirement of finding blade forms on which the flow separates at rather high spacing ratios .

Beschreibung: An investigation was conducted to determine the performance characteristics of the rotor and inlet guide vanes used in the axial-flow supersonic compressor of the XJ55-FF-1 turbojet engine. Outlet stators used in the engine were omitted to facilitate study of the supersonic rotor. The extent of the deviation from design performance indicates that the design-shock configuration was not obtained. A maximum pressure ratio of 2.26 was obtained at an equivalent tip speed of 1614 feet per second and an adiabatic efficiency of 0.61. The maximum efficiency obtained was 0.79 at an equivalent tip speed of 801 feet per second and a pressure ratio of 1.29. The performance obtained was considerably below design performance. The effective aerodynamic forces encountered appeared to be large enough to cause considerable damage to the thin aluminum leading edges of the rotor blades.

Beschreibung: As part of the performance investigation of compressors for the J33 turbojet engine, the A-21 model and the A-23 model with a 17- and a 34-blade impeller were operated with water injection at their respective design equivalent speeds of 11,500 and 11,750 rpm. Inlet conditions of pressure of 14 inches of mercury absolute and of ambient temperature correspond to those of the investigation of these models without water injection. The water-air ratio by weight ranged from 0.05 to 0.06. By the use of water injection, the peak pressure ratio of the A-21 compressor and the A-23 compressor with a 34-blade impeller increased approximately 0.38, whereas that of the A-23 compressor with a 17-blade impeller increased only 0.14. The decrease in maximum efficiency for the three compressors ranged from 0.12 to 0.14. The highest increase in maximum equivalent weight flow of air plus weight flow of water was 10.90 pounds per second obtained with the A-21 compressor. The increase in air weight flow alone was approximately 5.70 pounds per second for the A-21 compressor end the A-23, 17-blade compressor, which exceeded the increase of 3.15 pounds per second for the A-23; 34-blade compressor.

Beschreibung: A single-stage modification of the turbine from a Mark 25 torpedo power plant was investigated to determine the performance with two nozzle designs in combination with special rotor blades having a 20 inlet angle. The performance is presented in terms of blade, rotor, and brake efficiency as a function of blade-jet speed ratio for pressure ratios of 8, 15 (design), and 20. The blade efficiency with the nozzle having circular pas- sages (K) was equal to or higher than that with the nozzle having rectangular passages (J) for all pressure ratios and speeds investigated. The maximum blade efficiency of 0.571 was obtained with nozzle K at a pressure ratio of 8 and a blade-jet speed ratio of 0.296. The difference in blade efficiency was negligible at a pressure ratio of 8 at the low speeds; the maxim difference was 0.040 at a pressure ratio of 20 and a blade-jet speed ratio of 0.260.

Beschreibung: The J33-A-27 compressor was operated at an inlet pressure of 14 inches of mercury absolute and ambient inlet temperature over a range of equivalent impeller speeds from 6100 to 11,800 rpm. At the design equivalent speed of 11,800 rpm, the J33-A-27 compressor had a peak pressure ratio of 4.40 at an equivalent weight flow of 105.7 pounds per second and a peak adiabatic temperature-rise efficiency of 0.745. The maximum equivalent weight flow at design speed was 113.5 pounds per second.

Beschreibung: An investigation was conducted to determine the effect of turbine-disk cooling with air on the efficiency and the power output of the radial-flow turbine from the Turbo Engineering Corporation TT13-18 turbosupercharger. The turbine was operated at a constant range of ratios of turbine-inlet total pressure to turbine-outlet static pressure of 1,5 and 2.0, turbine-inlet total pressure of 30 inches mercury absolute, turbine-inlet total temperature of 12000 to 20000 R, and rotor speeds of 6000 to 22,000 rpm, Over the normal operating range of the turbine, varying the corrected cooling-air weight flow from approximately 0,30 to 0.75 pound per second produced no measurable effect on the corrected turbine shaft horsepower or the turbine shaft adiabatic efficiency. Varying the turbine-inlet total temperature from 12000 to 20000 R caused no measurable change in the corrected cooling-air weight flow. Calculations indicated that the cooling-air pumping power in the disk passages was small and was within the limits of the accuracy of the power measurements. For high turbine power output, the power loss to the compressor for compressing the cooling air was approximately 3 percent of the total turbine shaft horsepower.

Beschreibung: An investigation was conducted to determine the performance characteristics of the axial-flow supersonic compressor of the XJ-55-FF-1 turbo Jet engine. The test unit consisted of a row of inlet guide vanes and a supersonic rotor; the stator vanes after the rotor were omitted. The maximum pressure ratio produced in the single stage was 2.28 at an equivalent tip speed or 1814 feet per second with an adiabatic efficiency of approximately 0.61, equivalent weight flow of 13.4 pounds per second. The maximum efficiency of 0.79 was obtained at an equivalent tip speed of 801 feet per second.

Beschreibung: An investigation was conducted to determine the performance characteristics of the axial-flow supersonic compressor of the XJ55-FF-1 turbojet engine. An analysis of the performance of the rotor was made based on detailed flow measurements behind the rotor. The compressor apparently did not obtain the design normal-shock configuration in this investigation. A large redistribution of mass occurred toward the root of the rotor over the entire speed range; this condition was so acute at design speed that the tip sections were completely inoperative. The passage pressure recovery at maximum pressure ratio at 1614 feet per second varied from a maximum of 0.81 near the root to 0.53 near the tip, which indicated very poor efficiency of the flow process through the rotor. The results, however, indicated that the desired supersonic operation may be obtained by decreasing the effective contraction ratio of the rotor blade passage.