Steady Analysis of NACA Flush Inlet at High Subsonic and Supersonic Speeds

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2 , July 1-3, 2015, London, U.K. external flow affects the aerodynamics of airframe. Speed of air at engine face is required to be moderately subsonic, lower than the principle aircraft speed. Due to this fact, the basic shape of the duct becomes important. In order to reduce the speed of incoming air, the front part of the duct is in the form of a diffuser; increasing in area from the entry, to a position representing the engine face. The rear part of the duct is then convergent, simulating in essence the engine nozzle system [1].Principal stations in the flow are indicated in Figure 1. Mass Flow Rate and Pressure Recovery are the factors among many which govern the performance of inlet. Mass flow rate (kg/s) is the primary variable upon which all the others are dependent [4]: denoted by m, it is defined by m = ρ A V (1) In high speed flight an air inlet is a form of compressor; it accepts air initially at free stream Mach number and pressure, and converts it to lower Mach number and correspondingly higher static pressure, as required by the engine. For compressible flow conditions, this definition can further simplified (2) Equation 2 is the total pressure ratio. The efficiency of the inlet is defined by this equation is termed as pressure recovery of the inlet. This efficiency denotes the loss of total pressure from the free stream. Loss of total pressure can occur due to friction on the walls of the duct and on any external surface which is wetted by flow going into the duct, turbulent mixing associated with flow separation and due to shock waves [1]. Pressure recovery has a determining effect upon engine thrust [4], which may be defined as the resultant force in the direction of flight produced on the aerodynamic duct system by the internal flow. III. SUBMERGED INLET NACA duct or NACA scoop is a common form of lowdrag intake design. Originally it was developed by the National Advisory Committee for Aeronautics (precursor to NASA) in When properly implemented, it allows air to be drawn into an internal duct, often for cooling purposes, with a minimal disturbance to the flow. The design was originally called a "submerged inlet," since it consists of a shallow ramp with curved walls recessed into the exposed surface of a streamlined body, such as an aircraft. Submerged inlet experiments showed poor pressure recovery due to slow-moving boundary layer entering the intake. This design is believed to work because the combination of the gentle ramp angle and the curvature profile of the walls create counter-rotating vortices which deflect the boundary layer away from the intake and draws in the faster moving air, while avoiding the form drag and flow separation. This type of flush inlet generally cannot achieve the larger pressure recovery and flow quality(low distortion factor) of an external design, and so is rarely used for the jet aircraft intake application. It is, however, common for aircraft cooling applications in the form of submerged scoops. A. Working Principle of Flushed Inlet The working principle of submerged inlet is based on the formation of two counter-rotating vortices along the sharp leading edges of the surfaces leading up to the entry plane [1]. These vortices sweep the ramp boundary layer sideways and carry a proportion of it past the ends of the entry and out into the external stream. Thus the ramp arrangement works as a form of boundary layer diverter, reducing the effective position ratio of the inlet. Results can sometimes be enhanced by the addition of small ridges along the divergent sidewalls, increasing their effective height, and also by increasing the sidewall divergence to allow for slots at the ends of the entry, easing the passage of the boundary layer [4]. The other function of these vortices is that they drag the external flow into the inlet by their rotation. Other factors affecting success of the NACA submerged inlet depends on practical considerations of layout, such as space for a comparatively longer ramp at a sufficiently small angle (about 7degrees ) and on having some indication of the extent to which the boundary layer diverted from the inlet adds to the drag of the aircraft. Submerged inlet are generally considered unsuitable for supersonic speeds, because velocity, or Mach number, on the initial part of the ramp is higher than that of the free stream, however in the context of small auxiliary inlets, they can also be used at supersonic speeds [1]. Submerged inlets offer few distinctive advantages like low drag, small overall dimensions and small radar cross-section. These inlets also have some drawbacks like high attitude sensitivity (particularly at negative incidence), limited pressure recoveries, especially if the upstream Mach number exceeds unity and large internal volume requirement (as the inlets are long). B. Basic Parameters of Flush Inlets Submerged inlet has three basic parameters [6] which are Ramp Planform, Throat Aspect Ratio (Width to depth Ratio) and Ramp Angle. Different types of ramp plans were evaluated including Rectangular, straight divergent and curved divergent. It has been observed from previous investigations [5] that curved divergent shape give higher pressure recovery. Higher pressure recoveries are obtained for diverging ramps over a wide range of inlet velocity ratios [6]. These experimental results [5] suggest that pressure recovery becomes higher as the throat gets wider for the same entrance area. Due to this extension of throat width, the divergent angle of ramp walls is also increased and the height of throat becomes low. The vortex generated along ramp walls separates early and becomes stronger due to increased divergent angle. As the roll-up vortex moves downstream in the duct, the downwash of the vortex results in thinning the boundary layer on the ramp floor [19]. Thin boundary layer shall ultimately aid in achieving high pressure recovery.

4 , July 1-3, 2015, London, U.K. V. GRID INDEPENDENCE STUDY Fig.6. Meshing of complete geometry (4.09 million cells) The flow through and around the NACA flush inlet is modeled as follows: a) The computational domain was modeled as threedimensional (Cartesian coordinates) with a plane of symmetry (X-Z plane), which significantly reduced the computational effort. b) Flow was modeled as turbulent and steady for some computations c) Air was modeled as a compressible fluid obeying the ideal gas law. d) Temperature variations / effects are inherent in compressible analysis; therefore the energy equation was solved as well. Since heat transfer is not of prime interest in the present analysis therefore adiabatic conditions on flow domain boundaries (default) were accepted. Viscous dissipation was ignored since it is not expected to influence the computed results. e) Reynolds-averaged Navier-Stokes (RANS) system of equations for turbulent flow was solved. f) The two-equation Standard K-ε,, RKE, SST KW and Spalart-Allmaras (SA) turbulence models were used to determine the sensitivity of computed results with the turbulence model. No-slip velocity boundary conditions specified at the bottom wall that is flush inlet and the duct. Zero-shear (slip) boundary conditions specified at the top wall and the side wall in order to avoid the unnecessary formation of viscous boundary layers. At the inlet of the domain the total pressure, total temperature, static pressure and velocity were specified corresponding to the desired Mach number under standard sea-level conditions as part of the pressure-inlet boundary condition in FLUENT. Symmetry boundary condition (i.e. zero normal temperature gradient, zero-shear and zero normal velocity) was specified on the symmetry plane (X-Z). At the outlet of domain, static pressure and total temperature were specified as part of the pressure-outlet boundary condition. The convergence criteria for the scaled residuals for 3D steady analysis was set at 1x10-4. Numerically computed results change with the type and fineness of the mesh / grid used for computations. Because each result is an approximate solution of non linear partial differential equations, truncation errors exist in the flow domain solution. The truncation errors depend on the grid resolution hence it is important to decide that the grid size is sufficient to solve the flow domain accurately and small enough to save Computational time. To be able to decide the sufficient number of grid points, a grid sensitivity analysis for this specific problem is necessary. The purpose of the study is to determine the effects of truncation errors involved in the numerical analysis of submerged inlet. In order to determine the dependence of the computational results on grid, three different grids were computed at Mach 0.8, one with 0.36 million, 2.16 million and 4.09 million mesh size. SA turbulence model was used as turbulence model with default values of its constants. Mass flow rate and pressure recovery were selected as prime parameters for convergence. Following table shows that fine mesh constituting 4.09 million size gives best result with experimental values as calculated by Seddon [1]. Thus further analysis is done on fine mesh geometry. TABLE1 GRID INDEPENDENCE TABLE AT MACH 0.8 Grid Cell Size MFR PR Y+ REMARKS Coarse Not selected Medium Not selected Fine Selected VI. RESULTS Primary performance parameter for the present NACA flush inlet is the mass flow rate of air passing through the inlet because present configuration is representative of auxiliary cooling application. Since present configuration also represents an air intake system therefore secondary performance parameter is the total pressure recovery of the inlet. The computed mass flow rate of air passing through the inlet under different free stream Mach numbers is shown for various turbulence models. The mass flow rate data is made non-dimensional in terms of mass flow ratio which is the inlet mass flow rate divided by free stream mass flow rate through inlet capture area. Total pressure recovery is also shown for Mach numbers of 0.8, 1.2, 1.5 and 1.8 with different turbulence models. Total pressure recovery has been calculated at the NACA flush inlet exit plane.

5 , July 1-3, 2015, London, U.K Mass Flow Rate Variation (SA Turbulence Model) Mass Flow Rate (Kg/Hr) Fig.9. Pressure contour on symmetry plane flush inlet (Mach 0.8) Mach Number Fig.7. Mass Flow Rate at Mach 0.8,1.2,1.5 and 2.0 Mass flow rate of air passing through the inlet / duct decreases with increasing free stream Mach number. This is caused by the increased backpressure in the duct due to the stronger interaction of duct exit airflow with the free stream as the Mach number increases. Basic flow field in and around the NACA flush inlet at subsonic free stream Mach numbers consists of a part of the free stream airflow being forced into the intake by virtue of two strong counter-rotating vortices that form on the sidewalls. The airflow inside the inlet then passes through the duct having three (03) 90 o bends and about 2.27 m average length (from inlet throat to duct exit) and exits back to the free stream at a downstream location. Pressure Recovery (SA Turbulence Model) Pressure Recovery Mach Number Fig.8. Total pressure recovery at Mach 0.8,1.2,1.5 and 2.0 Fig.10. counter rotating vortices in flush inlet (Mach 0.8) The formation of the counter-rotating vortices at the flush inlet sidewalls influences the flow quite a distance above the flush surface and forces it inside the inlet. However the strength and location of these vortices do not solely determine the amount of airflow going inside the inlet. This is due to the uniqueness of the present configuration i.e. the airflow inside the duct being exited again to the free stream. Region of local flow acceleration at the forward half of the flush inlet is also evident and is probably caused by the close proximity of the sidewall vortices. Generally the total pressure recovery of air passing through the inlet / duct decreases with increasing free stream Mach number. This is caused by the increased total pressure loss associated with flow across a stronger shockwave in the flush inlet as the free stream Mach number (supersonic) increases. At subsonic free stream conditions (M=0.8) the computed total pressure recovery at the inlet exit plane is around 0.88 which is the optimum value indicated by past research on the flush inlet. Fig.11. Pressure contour on symmetry plane flush inlet (Mach 1.2)

6 , July 1-3, 2015, London, U.K. Region of local flow acceleration at the forward half of the flush inlet (before the shockwave) is evident and is caused by the supersonic expansion in the divergent channel formed by inlet geometry. After initial acceleration, the airflow undergoes rapid change through a series of shockwaves that are identifiable by rapid pressure increase and Mach number reduction on the symmetry plane. These shockwaves are located significantly ahead of throat. The location of the front shockwave is primarily determined by the amount of airflow entering the inlet. At design point this shockwave should form just ahead of the throat with highest airflow into the inlet. VII. CONCLUSIONS It is concluded that NACA flush inlet can be used in cooling application rather than primary source of air inlet due decreased mass flow rate, mass flow ratio and decreased total pressure recovery. REFERNCES [1] Seddon, J. and E.L. Goldsmith, Intake aerodynamics. 2nd ed. AIAA education series. 1999, Reston, VA: American Institute of Aeronautics and Astronautics / lackwell Science Ltd. [2] Raymer, D.P., Aircraft Design : A Conceptual Approach. 3rd ed. AIAA Education Series. 1999: AIAA [3] Masud, J., Computational Analysis Of NACA Flush Inlet for Secondary Air Source Applications, in 3rd International Bhurban Conference on Applied Sciences and Technology. 2004: Bhurban, Pakistan. [4] Seddon, J. and E.L. Goldsmith, Practical intake aerodynamic design. 1993, Oxford ; Boston: Blackwell Scientific Publications. [5] Mossman, E.A. and L.M. Randall, An Experimental Investigation of the Design Variables for NACA Submerged Duct Entrances [6] Frick, C.W., et al., An Experimental Investigation of NACA Submerged Duct Enterances. 1945, Ames Aeronautical Laboratory:Moffett Field, Calif. [7] ESDU, Drag and pressure recovery characteristics of auxiliary air inlets at subsonic speeds. 2004, ESDU International: London, UK. [8] Pope, S. B Turbulent Flows. Cambridge university press. ISBN [9] Sacks, A.H. and J.R. Spreiter, Theoretical Investigation of Submerged Inlets at Low Speeds [10] Frank, J.L., Pressure Distribution and Ram Recovery characteristics of NACA submerged inlets at high subsonic speed. 1950, NACA. [11] Frank, J.L. and R.A. Taylor, Comparison of Drag, Pressure Recovery, and Surface Pressure of a scoop type inlet and a NACA Submerged Inlet at Transonic Speeds. 1951, NACA.

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