SEARCH · Search NASA
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Thermodynamic flow analysis of feasibility of coolant gas addition to nozzle flow for improved particle formation efficiency in mixed flow colloid thrustor
Choked flow analogy for very low quality two-phase flows - homogeneous flow and hydraulic jump theoretical models for predicting axial pressure distribution
A study of the potential implementations of localized active flow control (AFC) technology onto future airplanes is presented. This collaborative investigation addresses key objectives of the NASA Advanced Air Transport Technology (AATT) Project, in terms of reduction in fuel consumption and lower emission. It specifically targets the goals set forth in a roadmap developed by the NASA/Boeing team. The roadmap is a result of a series of meetings held between the two parties over the years and it represents a shared vision for practical implementations, leading up to flight demonstrations of localized flow control. If successful, localized flow control may lead to important ramifications for next generation airplanes from both the economic and environmental perspectives. Under this contract localized AFC has been used to improve aerodynamic performance during high-lift operations using Computational Fluid Dynamics (CFD). Specifically, AFC has been applied at the aileron and at various location in the wing leading edge (LE) regions. The applications target reduced drag and enhanced lift over the range of practical angles of attack, including stall. These benefits translate to airplane performance improvements, such as longer range or larger payload. The CFD results are used to quantify potential aerodynamic benefits, as well as the input required for actuation. This helps identify the most promising candidates, which potentially provide material net airplane level enhancements using onboard fluidic sources. The airplane configuration selected for the CFD study is a representative of a future short/medium-range twin-engine airplane dubbed the Reference Aircraft. A slew of AFC applications has been explored and their aerodynamic performance enhancements were benchmarked against the baseline Reference Aircraft. Promising AFC candidates have been deemed practical and potentially suitable for both the aileron and the wing LE implementations. The findings on the Reference Aircraft are used to guide the development of the AFC-enhanced aileron for the CRM-HL. The wind-tunnel model of the CRM-HL will be used by NASA to validate the AFC concepts, complementing the CFD-based analysis and the integration study (final report document #3).
Supersonic two-dimensional and axisymmetric nonequilibrium flow theory and calculations for axisymmetric flow through nozzles
Three-dimensional compressor flow theory and real flow effect
Steady-state pressure drop characteristics of normal hydrogen gas flowing through electrically heated small diameter tube - laminar flow instability
Shadowgraph and flow visualization wind tunnel test of supersonic flow mechanisms during turbulent separation of cylinder-flare
Short duration simulation of rocket exhaust products, flow duration through nozzle and effect on recirculation flow fields
Laminar to turbulent flow transition concepts studied in relation to oscillatory flows found in circulatory system
Methane and natural gas flow through critical flow nozzles, calculating real gas effects on mass flow rate
Comparisons of the results of testing a single-stage axial-flow compressor with a solid-wall casing and with grooved casings are presented. The depth, location, and number of circumferential grooves in the casing over the casing over the rotor tip were varied. The near-stall weight flow was lower than that with the solid-wall casing for all but one grooved configuration indicating an improvement in the stall margin. The greatest reduction in the near-stall weight flow was noted for the configuration with five grooves located over the blade midchord region.
Flow visualization studies were conducted to evaluate techniques for injecting simulated-fuel and simulated-propellant gases into a spherical cavity for application to open-cycle gaseous-core nuclear rockets. Preliminary studies were conducted with six simulated-fuel injectors and eight simulated-propellant injection configurations. Additional tests were conducted with the best configuration to determine the effect of weight-flow ratio, gas density ratio, injector location, and flow distribution on the simulated-fuel containment characteristics.
Composite photographs of the surface flow and shadowgraphs of the shock wave pattern are presented that depict the hypersonic flow field of a typical delta wing space shuttle orbiter. Results from a wind tunnel test in air are given in side, oblique, and projected plan views for angles of attack from 0 to 60 deg. The tests were conducted at a Mach number of 7.4 and for Reynolds numbers based on body length of 6,000,000 and 9,000,000. The interrelationship is shown for the intersecting bow and wing leading edge waves with the surface flow for angles of attack for which: (1) the leading edge wave is attached, and (2) the leading edge wave is detached.
Wall static pressure measurements and performance parameters are presented for axisymmetric supersonic nozzles with relatively steep convergent sections and comparatively small radius-of-curvature throats. The nozzle walls were essentially adiabatic. These results are compared with those obtained in other nozzles tested previously to appraise the influence of contraction shape on performance. Both the flow coefficient and the thrust were less than the corresponding values for one-dimensional, isentropic, plane flow for both the axial and radial inflow nozzles considered, but the specific impulse, the most important performance parameter, was found to be relatively unchanged. The thrust decrement for the axial inflow nozzles was established primarily by the shape of the contraction section, and could be estimated reasonably well from a conical sink flow consideration. The radial inflow nozzle has a potential advantage from a cooling point of view if used in a rocket engine.
A wind tunnel study of the basic nature of plume induced flow separation is reported with emphasis on the unsteady aspects of the flow. Testing was conducted in a 6 inch by 6 inch blow-down supersonic wind tunnel. A cone-cylinder model with a pluming jet was used as the test model. Tests were conducted with a systematic variation in Mach number and plume pressure. Results of the tests are presented in the form of root-mean-squared surface pressure levels, power spectral densities, photographs of the flow field from which shock angles and separation lengths were taken, and time-averaged surface pressure profiles.
A second-order theory for supersonic flow past slender bodies is presented. Through the introduction of characteristic coordinates as independent variables and the expansion procedure proposed by Lin and Oswatitsch, a uniformly valid solution is obtained for the whole flow field in the axisymmetric case and for far field in the general three-dimensional case. For distances far from the body the theory is an extension of Whitham's first-order solution and for the domain close to the body it is a modification of Van Dyke's second-order solution in the axisymmetric case. From the theory useful formulas relating flow deflections to the Whitham F-function are derived, which permits one to determine the sonic boom strength from wind tunnel measurements fairly close to the body.
Development of a method of estimating deviation angles by analytical procedures was begun. Solutions for inviscid, irrotational flow in the blade-to-blade plane were obtained with a finite-difference calculation method. Deviation angles for a plane cascade with a rounded trailing edge were estimated by using the inviscid-flow solutions and three trailing-edge hypotheses. The estimated deviation angles were compared with existing experimental data over a range of incidence angles at inlet flow angles of 30 deg and 60 deg. The results indicate that deviation angles can be estimated accurately (within 1 deg) by using one of the three trailing-edge hypotheses, but only when pressure losses are low. A new trailing-edge hypotheses is presented which is suitable (for the cascade considered) for both low- and high-loss operating points.
Measurements have been made of the mean flow properties and turbulent fluctuations in adiabatic turbulent boundary layer flows subjected to distributed adverse pressure gradients. In the freestream region upstream of the adverse pressure gradient the Mach number was 3.86, the unit Reynolds number 5.3 million per foot. The boundary layer developed on the wall of an axisymmetric nozzle and straight test section. In order to avoid the effects of streamwise surface curvature the adverse pressure gradients at the test section wall were induced by contoured centerbodies mounted on the wind tunnel centerline. The flow under study simulated that which might be found in an axially symmetric engine inlet of a supersonic aircraft.