The supersonic wind tunnel is an indispensable facility for basic education in any course that covers compressible fl ows and one of the main pillars of any aerodynamic laboratory. The introduction of a supersonic wind tunnel at the aerodynamics laboratory of the Aerospace Engineering Department at Cairo University had often been postponed and was hindered by a lack of funds for the purchase of foreign equipment and expertise. Thoughts therefore turned to building such facility instead of buying it, substituting high-tech and complex foreign equipment for locally produced equipment, and 'thinking out of the box' to make the most use of available resources, even when this led to some unconventional applications. An extensive scheme for the design, fabrication, and realization of a multi-Mach number (M = 1.5, 2, and 2.5) supersonic wind tunnel for laboratory experiments is proposed in this paper. The proposed scheme is simple, detailed and multi-level; it starts by utilizing one-dimensional isentropic fl ow theory for the conceptual design phase and makes full use of computational fl uid dynamics at the detailed design phase. This ensured that we had a working design before we embarked on the manufacture of any components, which would have been costly to modify had there been any design error. A parametric study has been carried out for a number of design parameters, using numerical simulations. After the design and fabrication, a number of successful standard textbook experiments, for Mach number 2, were carried out as validation for the proposed scheme. The results showed good agreement with the theoretical predictions.Keywords compressible fl ow; wind tunnel testing; supersonic fl ow; computational fl uid dynamics; isentropic fl ow; method of characteristics; shock waves Subscripts 1 condition upstream of normal shock wave 2 condition downstream of normal shock wave e nozzle exit condition o stagnation condition Superscripts * critical condition at which M = 1 perturbation component normalized by critical speed of sound ∼ full velocity components = a + u or a + v

The Organization of the Petroleum Exporting Countries (OPEC) oil crisis of the mid 1970s led to a revival in interest in the propeller as a possible fuel-efficient propulsion for aircraft operating at subsonic cruise speeds. A propeller aerodynamics is complex and should be analyzed carefully to ensure maximum propellers efficiency. Detailed knowledge of flow patterns and aerodynamics loads is necessary for blade material and manufacturing process. In this study, an isolated propeller blade is chosen as the base of analysis, the geometry of the propeller: twist and chord variation with radius, are taken from real case module. The boundary conditions of the computational domain are set corresponding to that exist in the propeller manuals. A three dimensional unstructured grid was generated and adopted using commercial grid generator GAMBIT software. The governing equations are solved using FLUENT6.3.26 a commercial CFD code, which uses a control volume approach on a grid over the computational domain. Results identified that the propeller efficiency, power coefficient are increases to reach maximum values and then decreases with increase Mach number. The thrust coefficient decreases with increase Mach number.

Propeller driven-engines operate efficiently at low speeds, and ground maneuvers, but its performance is affected by operating in unsuitable environment. Actually, it is susceptible to encounter many physical problems such as erosion, corrosion, foreign object damage, and icing. These problems not only cause changes in air path boundaries but also yield changes in the aerodynamic performance of the engine components due to the change of the propeller profile shape and increase in the overall surface roughness. This work aims to study the effect of the particle initial velocity on the propeller erosion phenomena and the subsequent deterioration for the blades profile. Particle trajectory, erosion rate, frequency and the critical erosion area on the blade are the main issues under investigation. The domain selected for computational study is a periodic sector through the propeller bounding and the boundary conditions are set corresponding to that exist in the propeller manuals. A three dimensional unstructured grid was generated and adopted using commercial turbomachinery grid generator GAMBIT software. The governing equations are solved using FLUENT6.3.26 a commercial CFD code, which uses a control volume approach on a grid over the computational domain. A Lagrangian-formulated particle equation of motion is added to predict particle velocity and trajectories once the air flow field is obtained.

This paper presents the numerical prediction for an inviscid , rotational and compressible flow in a two-dimensional curved duct with inlet total pressure distortion. The numerical method is a fully implicit finite difference technique solving the unsteady form of the continuity, Euler equations and the steady form of the energy equation. The results are presented for different cases of inlet total pressure distortion with the same average total pressure for all cases. These results are compared with those obtained from uniform inlet total pressure distribution to assess the deviation resulting from the uniform total pressure assumption.

Flutter constraint, applicable to aircraft conceptual design, is constructed using response surface methodology. It is presented by the critical flutter speed, as a function of wing torsion stiffness, root chord, sweep, mass ratio, taper ratio, aspect ratio, center of gravity location and radius of gyration. The constraint to is a quadratic response surface polynomial. The D-optimal design is used to find the best combinations of design points required to determine the function coefficients. 'The Regier number criterion is used to calculate the critical flutter speed at these design points. Analysis of variance is used to remove the unreliable terms from the function. To match the Regier number criterion, two constraint functions suitable for subsonic aircraft with traditional wing are constructed. The first one is applicable to aircraft with low sweepback wing while the second one is applicable to aircraft with moderate sweepback wing. As a case study, the constraint function is applied within the conceptual design of a subsonic aircraft leading to a considerable weight saving.