ALBERT

Description: The results of the NACA tests to determine the tensile strength of some structural sheet materials heated to failure at temperature rates from 0.2 deg. F to 100 deg F per second under constant load conditions are reviewed . Yield and rupture stresses obtained under rapid-heating conditions are compared with the results of conventional elevated-temperature tensile tests. The relation between rapid-heating tests, short-time creep tests, and conventional creep tests is discussed . The application of a phenomenological theory for calculating rapid-heating curves is shown. Methods are given for predicting yield and rupture stresses and temperatures from master curves and temperature-rate parameters

Description: The addition of fuselage volume, concentrated on top of the forward portion of the fuselage, for the purpose of delaying the drag-rise Mach number of subsonic airplanes at lifting conditions is investigated. The additions have been designed on the basis of the area rule and other important considerations to provide greater practicability of application compared with shapings previously investigated. The addition delayed the drag-rise Mach number by an increment of approximately 0.03 for a configuration having a wing with moderate thickness and 35 deg of sweepback at a lift coefficient of 0.3. A lesser delay was obtained for a configuration with a thicker wing. The additions increase the nonlinearities of the variations of pitching moment with lift,

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.

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: A flight investigation was conducted to determine the effects of an inlet modification and rocket-rack extension on the longitudinal trim and low-lift drag of the Douglas F5D-1 airplane. The investigation was conducted with a 0.125-scale rocket-boosted model which was flight tested at the Langley Pilotless Aircraft Research Station at Wallops Island, Va. Results indicate that the combined effects of the modified inlet and fully extended rocket racks on the trim lift coefficient and trim angle of attack were small between Mach numbers of 0.94 and 1.57. Between Mach numbers of 1.10 and 1.57 there was an average increase in drag coefficient of about o,005 for the model with modified inlet and extended rocket racks. The change in drag coefficient due to the inlet modification alone is small between Mach numbers of 1.59 and 1.64

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: The basic aerodynamic characteristics of a 0.04956-scale model of the Convair TF-102A airplane with controls undeflected have been determined at Mach numbers from 0.60 to 1.135 for angles of attack up to approximately 22 deg in the Langley 8-foot transonic tunnel. In addition, comparisons have been made with data obtained from a previous investigation of a 0.04956-scale model of the Convair F-102A airplane. The results indicated the TF-102A airplane was longitudinally stable for all conditions tested. An increase in lift-curve slope from 0.045 to 0.059 and an 11-percent rearward shift in aerodynamic-center location occurred with increases in Mach number from 0.60 to approximately 1.05. The zero-lift drag coefficient for the TF-102A airplane increased 145 percent between the Mach numbers of 0.85 and 1.075; the maximum lift-drag ratio decreased from 9.5 at a Mach number of 0.60 to 5.0 at Mach numbers above 1.025. There was little difference in the lift and pitching-moment characteristics and drag due to life between the TF-102A and F-102A configurations. However, as compared with the F-102A airplane, the zero-lift drag-rise Mach number for the TF-102A was reduced by at least 0.06, the zero-lift peak wave drag was increased 50 percent, and the maximum lift-drag ratio was reduced as much as 20 percent.

Description: Four solid-propellant rocket engines of nominal 1000-pound-thrust were tested for starting characteristics at pressure altitudes ranging from 112,500 to 123,000 feet and at a temperature of -75 F. All engines ignited and operated successfully. Average chamber pressures ranged from 1060 to ll90 pounds per square inch absolute with action times from 1.51 to 1.64 seconds and ignition delays from 0.070 t o approximately 0.088 second. The chamber pressures and action times were near the specifications, but the ignition delay was almost twice the specified value of 0.040 second.

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.