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Results for “AIR INLETS - CENTRAL, SUPERSONIC”

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.

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At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1997

Bleed Hole Flow Phenomena Studied

Boundary-layer bleed is an invaluable tool for controlling the airflow in supersonic aircraft engine inlets. Incoming air is decelerated to subsonic speeds prior to entering the compressor via a series of oblique shocks. The low momentum flow in the boundary layer interacts with these shocks, growing in thickness and, under some conditions, leading to flow separation. To remedy this, bleed holes are strategically located to remove mass from the boundary layer, reducing its thickness and helping to maintain uniform flow to the compressor. The bleed requirements for any inlet design are unique and must be validated by extensive wind tunnel testing to optimize performance and efficiency. To accelerate this process and reduce cost, researchers at the NASA Lewis Research Center initiated an experimental program to study the flow phenomena associated with bleed holes. Knowledge of these flow properties will be incorporated into computational fluid dynamics (CFD) models that will aid engine inlet designers in optimizing bleed configurations before any hardware is fabricated. This ongoing investigation is currently examining two hole geometries, 90 and 20 (both with 5-mm diameters), and various flow features.

Source record↗
NASA NTRS1951

Preliminary Investigation of a Translating Cowl Technique for Improving Take-off Performance of a Sharp-lip Supersonic Diffuser

A preliminary investigation was conducted in quiescent air on a translating cowl technique for improving the take-off performance of a sharp-lip supersonic diffuser. The technique consists of cutting the cowling in a plane normal to its axis and then translating the forepart of the cowling in the forward direction. The leading edge of the fixed portion of the cowling is rounded. Appreciable improved inlet performance was obtained with a cowling translation corresponding to a gap of only 1/4 inlet radius.

ENGINES, TURBOJET - PERFORMANCE↗
NASA NTRS1988

Modeling of supersonic reacting flow fields

A detailed understanding of the scramjet combustor flow field is critical to the achievement of a successful design. Even though the combustor flow field is quite complex, it can be realistically viewed as a collection of spatially developing and reacting supersonic mixing layers that are initially discrete, but that ultimately merge into larger more complex zones. These mixing layers begin downstream of a set of fuel injectors that introduce gaseous hydrogen in both a parallel and transverse direction into a supersonic air stream entering from the engine inlet. The behavior of the initial portion of the combustor flow, in the mixing layers near the fuel injectors, appears to be most critical, since this is where the mechanism for efficient high speed mixing must be established to achieve the required degree of combustion downstream. Because of the structure of the flow field in this initial portion of the combustor, a single supersonic, spatially developing and reacting mixing layer serves as an excellent physical model for the overall flow field. Even though this reacting mixing layer flow is geometrically simple, it can still be made to retain all of the fluid mechanical and chemical complexities present in the actual combustor flow field.

Drummond, J. P.↗
NASA NTRS1952

Pressure Recovery, Drag, and Subcritical Stability Characteristics of Three Conical Supersonic Diffusers at Stream Mach Numbers from 1.7 to 2.0

A study of a 20 degree and a 25 degree half-angle high mass-flow ratio conical supersonic inlet was made on a 16-inch ram jet in the 8- by 6-foot supersonic tunnel. A greater range of stable subcritical operation was obtained with the low mass-flow ratio inlets; a greater range was obtained with the 25 degree than with the 20 degree half-angle low mass-flow ratio inlet. The high mass-flow ratio inlet had the least drag.

DIFFUSERS, SUPERSONIC↗
NASA NTRS1952

Performance of supersonic scoop inlets

A brief summary is made of the performance of a series of variable-geometry type scoop inlets investigated at the Lewis Laboratory on a model of the X-3 airplane fuselage. Data are presented for a range of inlet mass-flow ratios over a Mach number range from 0 to 2 and from 0 degrees to 12 degrees angle of attack. Rounded-lip inlets are found to give satisfactory performance to a Mach number of 1.5. Use of sharp rather than blunt lip configurations is shown to provide considerable gains in available thrust at Mach number 2.0 but penalize take-off performance to the extent that auxiliary air intakes may be required. No one inlet type was found to be superior over the entire range of operating variables. Thus, the choice of a specific inlet design will be influenced by structural and mechanical as well as aerodynamic considerations.

DIFFUSERS, SUPERSONIC↗
NASA NTRS1957

Inlets, exits, and cooling problems. 1 - Inlets

As design flight speeds are pushed progressively higher, the supersonic inlet becomes an increasingly important component of air-breathing propulsion systems. Currently, the turbojet engine is being considered for application at Mach numbers up to approximately 4 and the ramjet engine for application in the hypersonic region, or Mach numbers of 5 and above. Herein the inlet situation is surveyed and the merits of the various inlet-design philosophies are assessed on the basis of recent experimental data obtained at Mach numbers up to 5.

James F. Connors↗
NASA NTRS1977

Direct-connect tests of hydrogen-fueled supersonic combustors

Direct-connect tests of hydrogen-fueled supersonic combustors were performed using arc-heated air at combustor inlet Mach numbers of 2.9 to 3.2. Various axisymmetric combustor geometries of 5.89 and 6.96 cm (inner diameter) inlet were investigated; the fuel was injected from the wall either from a ring of equally spaced holes normal to the air stream, or from a circumferential slot oriented 45 deg downstream. The hole-type injectors consistently gave better results. The effects of various parameters are examined, and the performance comparison procedure is described. A theoretical model of the supersonic combustion process which includes a precombustion shock-compression is used to explain the character of the observed pressure distributions and to assess the effects of the measured heat transfer rates, deduced wall shear, and combustor geometry on performance.

Waltrup, P. J.↗
NASA NTRS1998

Inlet Performance Analysis Code Developed

The design characteristics of an inlet very much depend on whether the inlet is to be flown at subsonic, supersonic, or hypersonic speed. Whichever the case, the primary function of an inlet is to deliver free-stream air to the engine face at the highest stagnation pressure possible and with the lowest possible variation in both stagnation pressure and temperature. At high speeds, this is achieved by a system of oblique and/or normal shock waves, and possibly some isentropic compression. For both subsonic and supersonic flight, current design practice indicates that the inlet should deliver the air to the engine face at approximately Mach 0.45. As a result, even for flight in the high subsonic regime, the inlet must retard (or diffuse) the air substantially. Second, the design of an inlet is influenced largely by the compromise between high performance and low weight. This compromise involves tradeoffs between the mission requirements, flight trajectory, airframe aerodynamics, engine performance, and weight-all of which, in turn, influence each other. Therefore, to study the effects of some of these influential factors, the Propulsion System Analysis Office of the NASA Lewis Research Center developed the Inlet Performance Analysis Code (IPAC). This code uses oblique shock and Prandtl-Meyer expansion theory to predict inlet performance. It can be used to predict performance for a given inlet geometric design such as pitot, axisymmetric, and two-dimensional. IPAC also can be used to design preliminary inlet systems and to make subsequent performance analyses. It computes the total pressure, the recovery, the airflow, and the drag coefficients. The pressure recovery includes losses associated with normal and oblique shocks, internal and external friction, the sharp lip, and diffuser components. Flow rate includes captured, engine, spillage, bleed, and bypass flows. The aerodynamic drag calculation includes drags associated with spillage, cowl lip suction, wave, bleed, and bypass.

Jules, Kenol↗
NASA NTRS1984

Flow and heat transfer measurements along a cooled supersonic diffuser

An experimental investigtion was conducted to ascertain the mean flowfield, including shock wave structure, separated flow regions, turbulent boundary-layer growth, static pressure variations, wall heat transfer, and shear stresses in a second-throat, axisymmetric, supersonic diffuser with wall cooling. The diffuser inlet Mach number of the heated air flow was 3.76, the stagnation pressure was 6.8 atm, the ratio of wall to total gas temperature was 0.44, and the diffuser discharged to the atmosphere. The complex flowfield involved deceleration and acceleration regions, supersonic and embedded subsonic regions, and strong viscous regions with relatively large radial and axial variations. The heat transfer and wall static pressure distributions were remarkably similar, and heat transfer rates were high locally at oblique shock/turbulent boundary-layer interactions, in the pseudoshock region, and in the separation region in the diffuser outlet section.

Back, L. H.↗
NASA NTRS2010

Shock Positioning Controls Designs for a Supersonic Inlet

Under the NASA Fundamental Aeronautics Program, the Supersonics Project is working to overcome the obstacles to supersonic commercial flight. The supersonic inlet design that is utilized to efficiently compress the incoming air and deliver it to the engine has many design challenges. Among those challenges is the shock positioning of internal compression inlets, which requires active control in order to maintain performance and to prevent inlet unstarts due to upstream (freestream) and downstream (engine) disturbances. In this paper a novel feedback control technique is presented, which emphasizes disturbance attenuation among other control performance criteria, while it ties the speed of the actuation system(s) to the design of the controller. In this design, the desired performance specifications for the overall control system are used to design the closed loop gain of the feedback controller and then, knowing the transfer function of the plant, the controller is calculated to achieve this performance. The innovation is that this design procedure is methodical and allows maximization of the performance of the designed control system with respect to actuator rates, while the stability of the calculated controller is guaranteed.

Kopasakis, George↗
NASA NTRSDate not supplied

Auxiliary Inlet Design Study for Mach 1.4

Auxiliary inlet configurations were developed for inlets designed for a NASA Mach 1.4 Supersonic Technology Concept Aeroplane (STCA). The inlets included axisymmetric pitot and axisymmetric spike inlets. The auxiliary inlets were full circumferential slots within the cowl. Computational fluid dynamics (CFD) simulations were performed to obtain inlet performance metrics of total pressure recovery and radial distortion at the engine face for take-off and approach conditions. Methods of design of experiments were used to explore the statistical significance of the design factors for the auxiliary inlets which included the axial location, length of the opening, and auxiliary inlet angle. The results demonstrate the effect of the design factors and show the potential for improved inlet performance with the use of auxiliary inlets.

supersonic inlets↗