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The characteristics of a turbine pump for use with the nuclear engine for rocket vehicles are discussed. It was determined that the pump will be a two stage centrifugal pump with both stages having backswept impellers and an inducer upstream of the first stage impeller. The test program provided demonstration of the ability of the selected design to meet the imposed requirements.
Test bed No. 2 consists of 10 combustors welded in banks of 5 to 2 symmetrical tubular nozzle assemblies, an upper stationary thrust frame, a lower thrust frame which can be hinged, a power package, a triaxial combustion wave ignition system, a pneumatic control system, pneumatically actuated propellant valves, a purge and drain system, and an electrical control system. The power package consists of the Mark 29-F fuel turbopump, the Mark 29-0 oxidizer turbopump, a gas generator assembly, and propellant ducting. The system, designated as a linear aerospike system, was designed to demonstrate the feasibility of the concept and to explore technology related to thrust vector control, thrust vector optimization, improved sequencing and control, and advanced ignition systems. The propellants are liquid oxygen/liquid hydrogen. The system was designed to operate at 1200-psia chamber pressure at an engine mixture ratio of 5.5. With 10 combustors, the sea level thrust is 95,000 pounds.
Development in the design of leading heat pipes for the space shuttle are reported. The analysis, design, and integration of the heat pipes into the module structure are described along with the recommended tests. Results indicate the design goals were meet.
Several unique techniques and related devices are described which are in use at the Lewis Research Center for off-design testing of fan and compressor sections in full-scale jet engines. The devices presented permit a wide range of experimental conditions and minimize downtime for hardware changes. The techniques involve use of such devices as inlet pressure distortion jets, a hydrogen burner for inlet temperature distortions, fan back pressure jets to simulate a variable area nozzle, and either an inflow-outflow bleed system or a fuel spurt system to alter compressor discharge pressure.
The present work discusses in general terms the various kinds of ground facilities, in particular, wind tunnels, which support aerodynamic testing. Since not all flight parameters can be simulated simultaneously, an important problem consists in matching parameters. It is pointed out that there is a lack of wind tunnels for a complete Reynolds-number simulation. Using a computer to simulate flow fields can result in considerable reduction of wind-tunnel hours required to develop a given flight vehicle.
The feasibility of quiet, suction laminarized, high Reynolds number (Re) supersonic wind tunnel nozzles was studied. According to nozzle wall boundary layer development and stability studies, relatively weak area suction can prevent amplified nozzle wall TS (Tollmien-Schlichting) boundary layer oscillations. Stronger suction is needed in and shortly upstream of the supersonic concave curvature nozzle area to avoid transition due to amplified TG (Taylor-Goertler) vortices. To control TG instability, moderately rapid and slow expansion nozzles require smaller total suction rates than rapid expansion nozzles, at the cost of larger nozzle length Re and increased TS disturbances. Test section mean flow irregularities can be minimized with suction through longitudinal or highly swept slots (swept behind local Mach cone) as well as finely perforated surfaces. Longitudinal slot suction is optimized when the suction-induced crossflow velocity increases linearly with surface distance from the slot attachment line toward the slot (through suitable slot geometry). Suction in supersonic blowdown tunnels may be operated by one or several individual vacuum spheres.
Several unique techniques and related devices are in use at the Lewis Research Center for off-design testing of fan and compressor sections in full-scale jet engines. The devices presented not only permit a wide range of experimental conditions but also minimize downtime for hardware changes. The techniques involve use of such devices as inlet pressure distortion jets, a hydrogen burner for inlet temperature distortions, fan back pressure jets to simulate a variable area nozzle, and either an inflow-outflow bleed system or a fuel spurt system to alter compressor discharge pressure.
The severe HELIOS mission environment asked for an appropriate material selection for optical surfaces based on a good approach to predict degradation and contamination effects by analysis and an adequate vertification by test. Analysis results are compared with flight data of the HELIOS-A mission.
Test procedures necessary for the calibration and performance verification of the noncoherent C-band transponders after spacecraft hardware integration, but prior to spacecraft/launch vehicle integration are presented.
Results are summarized of acoustic tests on two advanced concepts of upper-surface-blowing propulsive lift devices.
The results of a static test of a Barber S-2 freight truck conducted to measure the stiffness and friction parameters of the modes of deformation which are being used in various mathematical models in the railroad industry were presented. Some difficulty was first experienced with the truck hardware since it was in an essentially new condition with many high spots causing interference. No difficulty was experienced once the interference was removed. The characteristics of the Barber S-2 are very similar to the ASF ride control truck. The major difference between the two trucks is the amount of friction between the bolster and side frames in both the vertical and lateral directions. The Barber S-2 has approximately twice the friction the ASF ride control truck has in the fully loaded condition. This does not necessarily imply that all trucks will have this same ratio of friction.
A one-component, fringe type laser velocimeter in the backscatter mode was used to measure flow velocities from 732 to 792 m/sec about a shock from an inclined flat plate in a Mach 5 tunnel. This test was chosen to determine system characteristics in applying the laser velocimeter to measure velocities in hypersonic flow fields. The seeding particles, less than 20 microns in diameter, were injected into the tunnel supply pipe, approximately 1 meter upstream of the nozzle throat, through a rake along the diameter. The LV signals were processed by two methods, a scanning spectrum analyzer and a high-speed counter. The measured velocity distributions obtained were analyzed to determine the errors due to particle lag, transit time broadening and high-speed counter quantizing error.
Concurrent tests were performed on a 1/16 and a 1/20 scale model (wing spans of 2.64 and 2.11 ft. respectively) of a modern low wing monoplane in the NACA 15 foot free-spinning wind tunnel. Results are presented in the form of charts that afford a direct comparison between the spins of the two models for a number of different conditions. Qualitatively, the same characteristic effects of control disposition, mass distribution, and dimensional modifications were indicated by both models. Quantitatively, the number of turns for recover and the steady spin parameters, with the exception of the inclination of the wing to the horizontal, were usually in good agreement.
During the abort sequence, the ET and orbiter separate under aerodynamic loading, with propellant remaining in the ET. The separation event included a seven second decelerating coast period during which the residual propellant accelerates relative to the ET/orbiter. At separation, ET clearance was primarily provided by aerodynamics acting on the ET to move it away. The motion of the propellant, primarily LOX, significantly influenced the resulting ET motion and could cause the ET to recontact the orbiter. A test program was conducted involving thirty-two drops with 1/60th scale models of the ET LOX tank. The objective was to acquire data on the nature of low g propellant reorientation, in the ET LOX tank, and to measure the forces exerted on the tank by the moving propellant.
The three engine cluster was fired at 70 percent power level for a nominal 15 seconds to evaluate the integrated performance of the main propulsion system. Engine ignition occurred at approximately 1403 with the planned mainstage duration achieved for all three engines. Operation of all systems was as expected with the exception of the recirculation pumps. The pumps were started while the propellant loading was in fast fill, but they cavitated and lost head at the termination of fast fill. The pumps were subsequently restarted after pressurizing the tank and draining back propellant to get good quality. Post test inspection of the engines revealed some discoloration on the inside of the thrust chamber and distorted drain lines for engine #2.
The methods by which aerodynamic coefficients are determined and discussed. These include: calculations, wind tunnel experiments and experiments in flight for various prototypes of the Alpha Jet. A comparison of obtained results shows good correlation between expectations and in-flight test results.
The engineering test program for the lander and the orbiter are presented. The engineering program was developed to achieve confidence that the design was adequate to survive the expected mission environments and to accomplish the mission objective.
Data for a NACA 0012-64 airfoil at Mach numbers of 0.3, 0.5, and 0.7 with the test section flexible walls set in the "straight" position are presented because they may prove useful in the development of wind tunnel correction techniques.