Inlet Cowl Research Articles

Design of an Internal Osculating Waverider Intake

The design of an inward-turning high-speed three-dimensional streamline-traced intake is presented from osculating axisymmetric theory. To satisfy the osculating intake design criteria, a stitched Busemann diffuser and internal conical flow-A solution are used as the basic isentropic compressive streamline. This new contour provides efficient compression, high flow uniformity, and straight leading-edge shocks of equal strength. Additionally, a novel method for constructing the inlet cowl is presented. The combined process leads to a new method of high-speed intake design. A generic shape-transitioned intake is constructed and named the Indiana inlet for the Indiana universities that contributed to the project. Computational fluid dynamic results are assessed to validate the design method for the two-dimensional parent flowfields and the full three-dimensional design.

Assessing the Performance of Hypersonic Inlets by Applying a Heat Source with the Throttling Effect

Utilization of a heat source to regulate the shock wave–boundary layer interaction (SWBLI) of hypersonic inlets during throttling was computationally investigated. A plug was installed at the intake isolator’s exit, which caused throttling. The location of the heat source was established by analysing the interaction of the shockwave from the compression ramp and the contact spot of the shockwave with that of the inlet cowl. Shockwave interaction inside the isolator was investigated using steady and transient cases. The present computational work was validated using previous experimental work. The flow distortion (FD) and total pressure recovery (TPR) of the inflows were also studied. We found that varying the size and power of the heat source influenced the shockwaves that originated around it and affected the SWBLI within the isolator. This influenced most of the performance measures. As a result, the TPR increased and the FD decreased when the heat source was applied. Thus, the use of a heat source for flow control was found to influence the performance of hypersonic intakes.

Supersonic flow in the rectangular duct of an air inlet with the separation-induced interaction of the boundary layer with shock waves

Results of a numerical study of the flow structure in the rectangular duct of a model air inlet with the interaction of the shock wave generated by the inlet cowl with the turbulent boundary layer on the inner duct surfaces are reported. A configuration involving a convergent entrance section and a constant-cross-section throat was analyzed. The numerical simulation of the three-dimensional flow with free-stream Mach number M = 4 was performed on the basis of the Reynolds-averaged Navier-Stokes equations and the k-ω SST model of turbulence. As a result of the calculations, the limit streamlines and the distributions of static pressure on the duct surfaces, as well as the distributions of total pressure and the fields of Mach number in duct cross sections were obtained. The structure of the flow is analyzed and the possibility of occurrence of longitudinal vortex structures in the corner region of the intake-duct throat is shown.

Three-dimensional shock interactions and vortices on a V-shaped blunt leading edge

The three-dimensional flow on a plate with a V-shaped blunt leading edge (VsBLEP) is investigated numerically and experimentally at a freestream Mach number 6. A complex saddle-shaped shock front is observed on this VsBLEP under the interactions between the detached shock (DS) induced by the swept blunt leading edge and the bow shock (BS) induced by the crotch. It is demonstrated that a new type of spatial transition exists on this saddle-shaped shock front, which involves the transition of shock interactions (i.e., DS and BS) from the same family upstream of the crotch to opposite families downstream of the crotch. Moreover, this transition is quantitatively identified according to the shock-induced spanwise velocity along the inflection line between DS and BS, which is of great importance because it affects the crossflow significantly. The inward crossflow induced by the swept blunt leading edge is enhanced in the region where the DS and BS are from the same family, and the shear layers generated in this region converge gradually to the spanwise symmetry plane, which results in the formation of a streamwise counter-rotating vortex pair (CVP). In the region where the DS and BS turn to opposite families, the inward crossflow is eliminated, and a five-shock structure is identified downstream of the crotch. The CVP remains close to the spanwise symmetry plane as it trails downstream, showing a far-reaching influence on the flowfield. This study indicates that the V-shaped blunt leading edge affects the downstream flow significantly and therefore should be examined carefully in practical applications, such as in the design of an inlet cowl lip.

Load characterization on the joints of the A320 engine inlet cowl acoustic panel

Aircraft operate in environments in which the components are subject to large temperature and pressure variations. In aircraft structures such as the engine nacelles, composed by several components and different materials, the presence of wear and corrosion becomes noticeable due to their operation in such environments. Corrective actions must be employed to the components which present this kind of problems. The acoustic panels of the inlet cowl of the Airbus A320/A321, present corrosion problems on the aluminium doublers of the joints. In order to develop a corrective action to the joint of the acoustic panels, the analysis of the mechanical behaviour and forces acting on the joint must be carried out. In this work, a methodology involving Computational Fluid Dynamics, Finite Element Method and Computer Aided Design tools is developed in order to evaluate the mechanical behaviour of the acoustic panels joint. The assessment of the aerodynamic loads acting on the inlet cowl was performed using STAR CCM+ software. The structural analysis of the joint members of the acoustic panels was made with the use of FEM tools in ANSYS Workbench software. The critical steps involved in the analysis are explained. Obtained results are discussed and conclusions are presented.

Numerical investigation on the assistant restarting method of variable geometry for high Mach number inlet

To compromise the compression efficiency and the starting properties, the inner contraction ratio (ICR) of a general high Mach number inlet is usually designed in the range of dual solution area. When going into an unstarted status, the high Mach inlet needs an assistant method to restart. This work explores a variable geometry method to restart the inlet. The rotating cowl is adopted to a typical Mach 4 cruising inlet, and the unsteady computation method with a dynamic Chimera grid technique is applied to simulate the rotating process of the inlet cowl. The change characteristics of the restarted performance at different rotating angle amplitude of the inlet cowl are investigated systematically. The numerical results reveal that the unstarted status of this typical inlet induced by the effect of high backpressure failed to restart if the inlet cowl rotating angle amplitude is under a small critical value, which is called lower critical angle. The inlet could restart if the cowl rotating angle amplitude is a little larger than the lower critical angle, and the flow may rapidly go to a steady condition after the inlet cowl returns to the design position. However, the performance of the restarted inlet is still worse than the design condition, because of the existence of an stable separation bubble on its shoulder, even if the inlet cowl stops rotating. The separation bubble becomes shrunk with an increasing the cowl rotating angle amplitude. When the inlet cowl rotating angle amplitude reaches a large critical value which is called upper critical angle, the separation bubble disappears, and all the separation is swallowed by the mean flow. Therefore the design performance of the inlet can be recovered, which means that the flow mass capture coefficient, total pressure recovery coefficient, drag and the outlet Mach number go back to the design level. It also shows that within the range of the lower and upper critical angles, the larger the rotating angle amplitude is, the more rapidly the separation bubble reaches stable state.

Influence of Hypersonic Inlet Cowl Lip on Flowfield Structure and Thermal Load

An asymmetrical hypersonic inlet cowl lip was numerically studied to explore a new approach of thermal protection. Comparative analyses of the flowfield structure and wall heat transfer with a symmetrical cowl lip were conducted. Results indicate that the peak heat transfer rate to the wall of the asymmetrical cowl lip decreases greatly when an impinging shock intersects a bow shock. Without the impinging shock, the location of the peak heat transfer rate occurs on the wall with a small radius, which does not coincide with the location of the air stagnation point; and the wall peak heat transfer rate of the asymmetrical cowl lip is higher than those of the symmetrical cowl lip, unless the flow attack angle to the wall with a large radius exceeds 12 deg.

A TIPSO algorithm assessment for aerothermodynamic optimization of hypersonic compression systems

A two-step improved particle swarm optimization (TIPSO) algorithm was recently proposed and used for the optimization of flexible satellite-control parameters. It was found to be much more stable and much less complex than other evolutionary algorithms. In this article the efficacy of the TIPSO algorithm is investigated for multidisciplinary optimization of aerothermodynamic parameters of performance, cowl deflection angle and shock–boundary layer separation on a cone-derived, wedge-integrated hypersonic (waverider) compression system. This algorithm uses an aggregate objective function. Optimization results from the TIPSO algorithm are compared with those obtained from a hybrid genetic algorithm, particle swarm optimization using an inertial weight approach and a multi-objective genetic algorithm. Since optimality of the forebody configuration is the basic requirement, the optimization variables selected are the isolator exit Mach number, static pressure ratio across the forebody–inlet configuration, cycle temperature ratio and cowl (inlet) deflection angle, which is the maximum deflection angle corresponding to the optimum oblique shock angle generated by the inlet cowl. Simulation results show that for the given case study only the TIPSO algorithm can locate the globally best aerothermodynamic–geometric parameters, compared with all the other optimization methods.

A New Support Structure in Waverider Force Measurement

In this paper a new support structure was given to solve the problem met in the force measurement with engine jet. The force measurements of waverider were undertaken in the hypersonic wind tunnel. The test condition was at Ma=6, angle of attack α=-6°-6°, at which we researched the effects on the vehicle aerodynamics of inlet cowl opening and closing, support system, engine jet, and pressure ratios. To decrease the effects of strut on the jet flow-field, we took sharp belly strut to support the model in the wind tunnel. The belly strut could support the waverider model, force measurement balance, and it could make the inlet flow set up, and provide the high pressure jet. The test results showed that the belly sharp strut had little effects on the flowfield and could inject the inlet flow, and could provide very high quality jet.

Surface-Flow Visualization and Pressure-Sensitive Paint Measurements in the Large-Scale Low-Boom Inlet

The local flowfield effects of vortex generators on a supersonic axisymmetric-relaxed external compression inlet have been investigated with surface-flow visualization and pressure-sensitive paint measurements. This inlet entails a mechanically simple design for providing a high-recovery/low-distortion flowfield and low sonic-boom signature. A set of vortex generators was incorporated into the design both upstream of the terminating normal shock and downstream inside the subsonic inlet diffuser for reducing separation zones near the shock/boundary-layer interaction and reducing radial distortion at the engine face, respectively. The inlet centerbody and downstream diffuser vortex-generator regions were imaged during wind-tunnel testing internally through the inlet cowl with a camera housing attached to the cowl. Surface-flow visualization revealed separated regions along the inlet centerbody for large mass-flow rates without vortex generators. Upstream vortex generators did reduce separation in the subsonic diffuser, and a unique perspective of the flowfield produced by the downstreamvortex generators was obtained. In addition, pressure distributions on the inlet centerbody, vortex generators, and struts were measured.

Influence of the Primary Jets and Fuel Injection on the Aerodynamics of a Prototype Annular Gas Turbine Combustor Sector

Transverse dilution jets are widely used in combustion systems. The current research provides a detailed study of the primary jets of a realistic annular combustion chamber sector. The combustor sector comprises an aerodynamic diffuser, inlet cowl, combustion dome, primary dilution jets, secondary dilution jets, and cooling strips to provide convective cooling to the liner. The chamber contracts toward the end to fit the turbine nozzle ring. 2D PIV is employed at an atmospheric pressure drop of 4% (isothermal) to delineate the flow field characteristics. The laser is introduced to the sector through the exit flange. The interaction between the primary jets and the swirling flow as well as the sensitivity of the primary jets to perturbations is discussed. The perturbation study includes: effect of partially blocking the jets, one at a time, the effect of blocking the convective cooling holes, placed underneath the primary jets and shooting perpendicular to it. In addition, the effect of reducing the size of the primary jets as well as off-centering the primary jets is explained. Moreover, PIV is employed to study the flow field with and without fuel injection at four different fuel flow rates. The results show that the flow field is very sensitive to perturbations. The cooling air interacts with the primary jet and influences the flow field although the momentum ratio has a 100:1 order of magnitude. The results also show that the big primary jets dictate the flow field in the primary zone as well as the secondary zone. However, relatively smaller jets mainly influence the primary combustion zone because most of the jet is recirculated back to the CRZ. Also, the jet penetration is reduced with 25% and 11.5% corresponding to a 77% and 62% reduction of the jet’s area, respectively. The study indicates the presence of a critical jet diameter beyond which the dilution jets have minimum impact on the secondary region. The jet off-centering shows significant effect on the flow field though it is in the order of 0.4 mm. The fuel injection is also shown to influence the flow field as well as the primary jets angle. High fuel flow rate is shown to have very strong impact on the flow field and thus results in a strong distortion of both the primary and secondary zones. The results provide useful methods to be used in the flow field structure control. Most of the effects shown are attributed to the difference in jet opposition. Hence, the results are applicable to reacting flow.

Development and evaluation of TiAl sheet structures for hypersonic applications

A cooperative program between the National Aeronautics and Space Administration (NASA), the Austrian Space Agency (ASA), Pratt & Whitney, Engineering Evaluation and Design, and Plansee AG was undertaken to determine the feasibility of achieving significant weight reduction of hypersonic propulsion system structures through the utilization of TiAl. A trade study defined the weight reduction potential of TiAl technologies as 25–35% compared to the baseline Ni-base superalloy for a stiffener structure in an inlet, combustor, and nozzle section of a hypersonic scramjet engine. A scramjet engine inlet cowl flap was designed, along with a representative subelement, using design practices unique to TiAl. A subelement was fabricated and tested to assess fabricability and structural performance and validate the design system. The TiAl alloy selected was PLANSEE's third generation alloy Gamma Met PX, 1 1 Gamma Met PX is a trademark of PLANSEE AG, Austria. a high temperature, high-strength γ-TiAl alloy with high Nb content. Characterization of Gamma Met PX sheet, including tensile, creep, and fatigue testing was performed. Additionally, design-specific coupons were fabricated and tested in order to improve subelement test predictions. Based on the sheet characterization and results of the coupon tests, the subelement failure location and failure load were accurately predicted.

Acoustically treated structurally reinforced sound absorbing panel

An acoustically treated structurally reinforced sound absorbing panel having particular application for use at the inlet cowl of a jet engine to reduce engine noise leaked into the environment. The sound absorbing panel includes a sound receiving chamber having a honeycomb shaped core manufactured from corrugated metallic ribbons. The sound receiving chamber is sandwiched between a metallic backskin at one side thereof, and a metallic face sheet at the opposite side. The face sheet is perforated so that incoming sound waves which represent the engine noise enter the sound receiving chamber within which they are channeled towards and reflected between the metallic backskin and face sheet. An acoustic attenuation medium having a plurality of short metal fibers that are held together by a sintering process is diffusion bonded to the perforated face sheet so that the reflected sound waves which are channeled through the sound receiving chamber are transmitted through and attenuated by the acoustic attenuation medium.

Prediction and Measurement of the Total Pressure Loss in an Engine Representative Diffuser System

An experimental and computational investigation of the flow in an engine representative diffuser system has been performed. Many parametric changes to the system were considered including inlet conditions, prediffuser geometry, cowl geometry and prediffuser strut position. The focus of this paper is an overview of the results from the twelve configurations that were both predicted and experimentally tested. It was shown that the CFD predictions contained a high degree of numerical error and that to reduce this to a low level would currently be impractical for a parametric study of this type. However, by carefully maintaining approximately the same degree of numerical error in the predictions, it was shown that a valid parametric study could be performed to the extent that the same conclusions about the parametric changes can be drawn equally from the predictions or measurements with one small exception. It was recognized that the level of agreement between the predicted and measured losses was to a degree fortuitous and that the k–ε turbulence model performs poorly for this type of flow.

Aspiration Efficiency and Inlet Wall Deposition in the Fiber Sampling Cassette

The 25-mm diameter sampling cassette with a 50-mm long conductive inlet (cowl) is widely used for sampling fibers. Comprehensive testing of the sampling accuracy of this device has not been carried out. Several researchers have found significant fiber deposits in the inlet rather than on the collection filter. The sampler approaches the dimensions of a thin-walled tubular inlet, which has been studied extensively both theoretically and experimentally. A semiempirical model was used to calculate aspiration efficiency and inlet deposition for compact particles. The model was based on particle aerodynamic diameter and should apply with sufficient accuracy to fibers. This semiempirical model includes effects due to particle inertia, gravitational settling, and vena contracta deposition. The results of the model were compared to experimental measurements from a wind tunnel and from previous field studies. The laboratory measurements agree well with the model, but field measurements of fibers show higher inlet ...

Shock interference prediction using direct simulation Monte Carlo

The shock interaction produced when an incident shock impinges on an inlet cowl lip is investigated for a cowl lip radius of 0.1 in. and flight conditions of Mach 15 and 35-km altitude. This problem is of interest in the design of hypersonic flight vehicles because the interference heating at the cowl lip surface from such a shock interaction is expected to be very high and the design of adequate thermal protection in this region will be challenging. Noncontinuum effects may be significant at this combination of altitude and small dimensions. If this is so, a computational solution based on the Navier-Stokes equations could significantly overpredict the surface heating on the cowl lip. Therefore, a kinetic theory solution technique, direct simulation Monte Carlo (DSMC), is used to predict the flowfield and surface heating. The details of modeling this problem and some preliminary results are presented. The same qualitative flow patterns are seen with DSMC and a continuum solution of the shock interference case, but the peak heating predicted with DSMC is somewhat lower.

Euler flow analysis of turbine powered simulator and fanjet engine

An Euler flow analysis of turbine powered simulator (TPS) and fanjet engine was made utilizing MacCormack's finite volume scheme. TPS simulation computations for an axisymmetric geometry were compared with experimental data obtained in the National Aerospace Laboratory transonic wind tunnel. The pressure distributions on the inlet cowl and corejet cowl surfaces showed excellent agreement. The effect of differences in engine conditions of corejet temperature and mass flux between the TPS flow and a real fanjet engine was analyzed with the computations. The results showed no significant differences in the external flowfield surrounding the exhaust jet plume. A three-dimensional analysis showed the effects of angle of attack on the inlet flowfield and exhaust jet plume.

A wind tunnel inlet flow simulation

Method The inlet flow is simulated by solving the unsteady compressible Euler equations in three dimensions using MacCormack's explicit predictor-corrector scheme1'2 in Cartesian coordinates with a finite volume formulation. The Euler code used here is a modified version3 of that used in Ref. 4. The code was used to simulate the flow induced in the wind tunnel by a fan downstream of the test section (sink-type flow). The three-dimensional computational grid for the inlet, as shown in Fig. 1, is divided into two zones. The grid is generated algebraically and plane by plane using an interpolation scheme. The grid size for zone 1 is 1 1 x 32 x 30 and for zone 2 is 39 x 19 x 17. The nozzle has a 5:1 contraction ratio, defined from where the inlet cowl ends (the x = 0 cross-flow plane) to the test section. A planar grid, enlarged near the cowl and representing a horizontal cut (i.e., the z plane), is shown in Fig. 2. The computations are carried out corresponding to quiescent atmospheric conditions at the upstream boundary. Initially, stagnation conditions are imposed everywhere. The pressure drop that generates the flow in the inlet and yields a given test section velocity is introduced at the downstream boundary gradually in time. The flow can be imagined to exit at the downstream boundary at somewhat lower than the upstream stagnation pressure. This pressure is given by Pe =Ps ~

Effect of inlet additive drag on aircraft performance

The percentage of the theoretical additive drag (DADD? the force associated with spilling excess air around a jet engine inlet) that may be recovered as a change in the pressure drag of the inlet cowl is discussed. It is shown that, for off-design operation, DADD can be a significant percentage of the total airplane drag. An additive drag correction factor KADD is defined such that the drag force felt by the vehicle equals (KADD X DADD). The effect of KADD on typical aircraft missions is presented. Miscellaneous test results are shown to substantiate the KADD values used. A model wind-tunnel program aimed specifically at getting systematic additive drag data is described, and first phase test results are presented.