ALBERT

Description: 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. The agreement obtained was within 3 percent at the midspan and tip regions of the blade. At the root region of the blade, the agreement was within 3 percent for coolant flows within the turbulent flow regime and within 10 percent for coolant flows in the laminar regime. The calculated and measured cooling-air pressure losses through the blade agreed within 5 percent. Calculated spanwise blade temperatures for assumed turboprop engine operating conditions of 2000 F turbine-inlet gas temperature and flight conditions of 300 knots at a 30,000-foot altitude agreed well with those obtained by the extrapolation of correlated experimental data of a static cascade investigation of these blades.

Description: Two short turbojet combustors designed for use with vaporized hydrocarbon fuels were tested in a one-quarter annular duct. The experimental combustors consisted of many small "swirl-can" combustor elements manifolded together. This design approach allowed the secondary mixing zone to be considerably reduced over that of conventional combustors. The over-all combustion lengths, for the two configurations were 13.5 and 11.0 inches, approximately one-half the length of the shortest conventional combustors. These short combustors did not provide combustion efficiencies as high as those for conventional combustors at low pressures. However, over the range of combustor-inlet total-pressures expected in aircraft capable of flight at Mach numbers of 2.5 and above, these short combustors gave very high efficiencies. A combustion efficiency of 97 percent was obtained at a combustor-inlet total-pressure of 25.0 inches of mercury absolute, reference velocity of 120 feet per second, and inlet-air total temperature of 1160 deg R. By proportioning the fuel flow between the manifold rows of can combustor elements, control of the combustor-outlet radial total-temperature profile was demonstrated. Combustor totalpressure loss varied from 0.75 percent of the inlet total pressure at isothermal conditions and a reference velocity of 75 feet per second to 5.5 percent at a total-temperature ratio of 1.8 and a reference velocity of 180 feet per second.

Description: This analysis investigates the application of gas turbine engines at a cruise Mach number of 4.

Description: A program was conducted in an altitude facility at the NACA Lewis laboratory to investigate the effects of rapid inlet pressure oscillations on the operation of a current turbo jet engine. These pressure oscillations were approximately sinusoidal in form and were generated to cover a frequency range of 2 to 75 cycles per second and an amplitude range of 10 to 70 percent of the free-stream total pressure. As the oscillation progressed through the compressor, the amplitude was attenuated considerably and a relatively large phase shift (lag) occurred. Engine stall limits obtained during pressure oscillations differed from quasi-steady-state stall limits as defined by over-all compressor pressure ratio.

Description: A lightweight turbine rotor assembly was devised, and components were evaluated in a full-scale jet engine. Thin sheet-metal airfoils were brazed to radial fingers that were an integral part of a number of thin disks composing the turbine rotor. Passages were provided between the disks and in the blades for air cooling. The computed weight of the assembly was 50 percent less than that of a similar turbine of normal construction used in a conventional turbojet engine. Two configurations of sheet-metal test blades simulating the manner of attachment were fabricated and tested in a turbojet engine at rated speed and temperature. After 8-1/2 hours of operation pieces broke loose from the tip sections of the better blades. Severe cracking produced by vibration was determined as the cause of failure. Several methods of overcoming the vibration problem are suggested.

Description: The performance and operational characteristics of two afterburner configurations for the Iroquois turbojet engine were evaluated in an altitude test chamber over a range of afterburner equivalence ratios at afterburner-inlet pressures from 733 to 3186 pounds per square foot absolute. These conditions correspond to an altitude range from 38,700 to 66,800 feet at a flight Mach number of 1.5. The only difference between the two afterburner configurations was in the pattern of afterburner fuel injection. At an afterburner-inlet pressure of approximately 3100 pounds per square foot absolute, corresponding to an altitude of 38,700 feet and a_ flight Mach number of 1.5, the combustion efficiency of both configurations reached peak values of 0.80 to 0.85 at equivalence ratios of 0.35 to 0.40. However, further reduction in the afterburner-inlet pressure severely affected combustion efficiency. For example, at an afterburner inlet pressure level of 700 to 1000 pounds per square foot absolute, the efficiency for both configurations was 0.20 to 0.40.

Description: The multistage turbine from the J73 turbojet engine has previously been investigated with standard and with reduced-chord rotor blading in order to determine the individual performance characteristics of each configuration over a range of over-all pressure ratio and speed. Because both turbine configurations exhibited peak efficiencies of over 90 percent, and because both units had relatively wide efficient operating ranges, it was considered of interest to determine the performance of the first stage of the turbine as a separate component. Accordingly, the standard-bladed multistage turbine was modified by removing the second-stage rotor disk and stator and altering the flow passage so that the first stage of the unit could be operated independently. The modified single-stage turbine was then operated over a range of stage pressure ratio and speed. The single-stage turbine operated at a peak brake internal efficiency of over 90 percent at an over-all stage pressure ratio of 1.4 and at 90 percent of design equivalent speed. Furthermore, the unit operated at high efficiencies over a relatively wide operating range. When the single-stage results were compared with the multistage results at the design operating point, it was found that the first stage produced approximately half the total multistage-turbine work output.

Description: Internal performance of an XJ79-GE-1 variable ejector was experimentally determined with the primary nozzle in two representative after-burning positions. Jet-thrust and air-handling data were obtained in quiescent air for 4 selected ejector configurations over a wide range of secondary to primary airflow ratios and primary-nozzle pressure ratios. The experimental ejector data are presented in both graphical and tabulated form.

Description: Nine divergent-shroud ejector configurations were investigated to determine the effect of shroud divergence angle on ejector internal performance. Unheated dry air was used for both the primary and secondary flows. The decrease in the design-point thrust coefficient with increasing flow divergence angle (angle measured from primary exit to shroud exit) followed very closely a simple relation involving the cosine of the angle. This indicates that design-point thrust performance for divergent-shroud ejectors can be predicted with reasonable accuracy within the range investigated. The decrease in design-point thrust coefficient due to increasing the flow divergence engle from 120deg to 30deg (half-singles) was approximately 6 percent. Ejector air-handling characteristics and the primary-nozzle flow coefficient were not significantly affected by change in shroud divergence angle.

Description: A two-spool turbojet engine was operated in the Lewis altitude wind tunnel to study the inception of compressor surge. In addition to the usual steady-state pressure and temperature measurements, the compressors were extensively instrumented with fast-response interstage pressure transducers. Thus it was possible to obtain maps for both compressors, pressure oscillations during rotating stall, effects of stall on efficiency, and stage-loading curves. In addition, with the transient measurements, it was possible to record interstage pressures and then compute stage performance during accelerations to the stall limit. Rotating stall was found to exist at low speeds in the outer spool. Although the stall arose from poor flow conditions at the inlet-stage blade tips, the low-energy air moved through the machine from the tip at the inlet to the outer spool to the hub at the inlet to the inner spool. This tip stall ultimately resulted in compressor surge in the mid-speed region, and necessitated inter-compressor air bleed. Interstage pressure measurements during acceleration to the compressor stall limit indicated that rotating stall was not a necessary condition for compressor surge and that, at the critical stall point, the circumferential interstage pressure distribution was uniform. The exit-stage group of the inner spool was first t o stall; then, the stages upstream stalled in succession until the inlet stage of the outer spool was stalled. With a sufficiently high fuel rate, the process repeated with a cycle time of about 0.1 second. It was possible to construct reproducible stage stall lines as a function of compressor speed from the stage stall points of several such compressor surges. This transient stall line was checked by computing the stall line from a steady-state stage-loading curve. Good agreement between the stage stall lines was obtained by these two methods.

Description: The performance characteristics of the 19B-8 and 19XB-1 turbojet engines and the windmilling-drag characteristics of the 19B-6 engine were determined in the Cleveland altitude wind tunnel. The investigations were conducted on the 19B-8 engine at simulated altitudes from 5000 to 25,000 feet with various free-stream ram-pressure ratios and on the 19XB--1 engine at simulated altitudes from 5000 to 30,000 feet with approximately static free-stream conditions. Data for these two engines are presented to show the effect of altitude, free-stream ram-pressure ratio, and tail-pipe-nozzle area on engine performance. A 21-percent reduction in tail-pipe-nozzle area of the 19B-8 engine increased the let thrust 43 percent the net thrust 72 percent, and the fuel consumption 64 percent. An increase in free-stream ram-pressure ratio raised the jet thrust and the air flow and lowered the net thrust throughout the entire range of engine speeds for the 19B-8 engine. At similar operating conditions, the corrected jet thrust and corrected air flow were approximately the same for both engines, and the corrected specific fuel consumption based on jet thrust was lower for the 19XB-1 engine than for the 19B-8 engine. The thrust and air-flow data obtained with both engines at various altitudes for a given free-stream rampressure ratio were generalized to standard sea-level atmospheric conditions. The performance parameters involving fuel consumption generalized only at high engine speeds at simulated altitudes as high as 15,000 feet. The windmilling drag of the 19B-8 engine increased rapidly as the airspeed was increased.