Design and flow calculation of mixed compression supersonic air intake

Dissertant

Partu, Fawwaz Farid

University

University of Technology

Faculty

-

Department

Department of Electromechanical Engineering

University Country

Iraq

Degree

Master

Degree Date

2007

English Abstract

A theoretical investigation on a supersonic air intake was made.

The supersonic air intake considered is of variable geometry 2-D mixed compression intake and this investigation the air intake is divided into.

External! and internal parts, each part ts analyzed and designed optimally.

The design conditions chosen were at (Mo= 1, 6), altitude (H) = 9000m, and mass flow rate (m') = 60kg / see, (for the aircraft C39A).

The external part consists of two wedges which begin with the vertex of the first wedge up to inlet of intake ; where the internal pan was begin.

The flow is assumed compressible, invjscid, and one dimensional .The flow properties (P, T, p, P0, Tol and po) were calculated using the analytical method hazed on correlations across oblique shock waves.

This part was analyzed for several Mach numbers (1.6-25) and for wedge angles (12-3).

this part was optimized to get the maximum total pressure recovery.

The internal part of intake begins downstream of the external part of the intake.

This part of intake has axe symmetric convergence-divergence form.

The first part of supersonic diffuser consists of a convergent passage with a sharp inward facing ramp with wall deflection angle (3).

The number of oblique shock waves is induced by the sharp cowl lips at the diffuser entrance and their reflections that may continue along the convergent passage depending on the free-stream Mach number and diffuser wall turning angle.

The location of the normal shock wave is fixed at diffuser throat at its "on-design" and "off-design" operations.

The convergent part was analyzed for different convergent lengths to maintain the oblique shock and their reflection was terminated before the throat position to avoid flow distortion and to get maximum total pressure recovery.

The second part of diffuser consists of divergent passage with divergence angle (0).

The flow axisymmetric.

The flow equations (continuity, momentum, energy and state) in Cartesian coordinate system, was solved by using numerical solution based on a finite difference technique, where the set of equations governing the problem is to be solved iteratively by using Newton-Raphson method.

For the external part, the analytical results show that, the flow over wedge is initially supersonic just behind the shock and remains supersonic but at Mach number lower than free-stream Mach number.

The total pressure recovery increase with increasing of the first wedge angle for the range (2.5-4.5) for constant second wedge angle until it reaches maximum value then it decreases fur values above (4.5).

For the internal part, the analytical results for convergent diffuser shows that, the flow is supersonic and remains supersonic but at Mach number lower than the entering Mach number.

The total pressure recovery decreases along the convergent length, because of the loss in total pressure across the re fleet ion of oblique shock waves.

The numerical results for divergent diffuser show that Mach number decreases with distance, the static pressure increases with distance and the total pressure recovery remains constant with distance.

The total pressure recovery value depends upon free-stream Mach number and position of the normal shock, since in this study the position of normal shock is fixed at diffuser throat.

A computer program in (MATLAB) language was built and developed to solve die equations of the flow over the angle of wedge and along the whole diffuser.

Main Subjects

Mechanical Engineering

Topics

• Atmospheric pressure
• Air flow

• APA
• MLA
• AMA

Language

English

Data Type

Arab Theses

Record ID

BIM-305496