External Compression Supersonic Inlet Research Articles

Buzz flow diversity in a supersonic inlet ingesting strong shear layers

To further explore the shear-layer-induced buzz diversity recently discovered in an overspeed supersonic inlet, an external-compression supersonic inlet is specially designed and carefully studied through wind-tunnel test at its design Mach number of 2.0. It is observed that the flow instability begins with the little buzz mode featuring a stable terminal shock and a lightly unsteady shock-induced separation bubble. Then a combination of another two stronger buzz styles replaces the little buzz. Of them the medium buzz is characterized by a locally destabilized terminal shock and the varying cowl-side reverse flow, whereas the big buzz is distinguished by intense separation unsteadiness. Interestingly, the medium buzz cannot stay long and vanishes eventually. Instead of an ordinary buzz case in the design Mach number situation, current buzz flows are found pretty similar to the overspeed inlet case superficially and fundamentally under the influence of two strong shear layers, albeit with a clear difference in the duration of the medium buzz. Accordingly, the possibility of buzz flow diversity and the related shear-layer effect proposed recently get further supported. Besides, it is indicated that the overspeed operation is not the only situation allowing for the diversified inlet buzz.

Side Gust Effects on the Performance of a Supersonic Inlet with Bleed

A computational investigation on the effects of side gust on the performance of a supersonic inlet with a bleed system at Mach 1.8 is reported. This is an extension to a recent study (Halwas, H. K., and Aggarwal, S. K., “Effect of Side Gust on the Performance of External Compression Supersonic Inlet,” Journal of Aircraft, Vol. 56, No. 2, March 2019, pp. 569–582) that analyzed a similar inlet without bleed. The supersonic section is designed using the Taylor–Maccoll method, and the rest of the inlet is designed using a methodology from the literature. The bleed is positioned at the throat inner wall, and its shape is a combination of a ram scoop and a flush slot. ICEM CFD software is used to generate a three-dimensional (3-D) structured mesh. ANSYS Fluent 18.2 code is used to perform 3-D simulations. Grid sensitivity tests are performed using local grid refinements and by varying the total grid points from 3 to 10 million. The effects of side gust on the shock system and flow are analyzed by examining the Mach number and pressure distributions. The inlet performance is characterized in terms of the total pressure loss, flow distortion, and mass flow ratio parameters for a gust speed of and angles of 0, 30, and 60 deg. The results indicate that the inlet performance is significantly affected by the side gust, although the bleed system mostly removes the low-momentum flow.

Design of Inward-Turning External Compression Supersonic Inlet for Supersonic Transport Aircraft

Inward-turning external compression intake is one of the hybrid intakes that employs both external and internal compression intakes principle. This intake is commonly developed for hypersonic flight due to its efficiency and utilizing fewer shockwaves that generate heat. Since this intake employ less shockwaves, this design can be applied for low supersonic (Mach 1.4 - 2.5) intakes to reduce noise generated from the shockwaves while maintaining the efficiency. Other than developing the design method, a tool is written in MATLAB language to generate the intake shape automatically based on the desired design requirement. To investigate the intake design tool code and the performance of the generated intake shape, some CFD simulation were performed. The intake design tool code can be validated by comparing the shockwave location and the air properties in every intake's stations. The performance parameters that being observed are the intake efficiency, flow distortion level at the engine face, and the noise level generated by the shockwaves. The design tool written in MATLAB is working as intended. Two dimensional axisymmetric CFD simulations validation has been done and the design meets the minimum requirement. As for the 3D inlet geometry, with a little modification on diffuser and equipping vent to release the buildup pressure, the inlet has been successfully met the military standard on inlet performance (MIL-E-5007D). This design method also has feature to fit every possible throat cross sectional shapes and has been proven to work as designed.Keywords— Inward-turning, Supersonic, Engine Intakes, Low- noise, Design Method

Effect of Side Gust on Performance of External Compression Supersonic Inlet

This Paper reports a computational investigation on the effects of side gust on the performance of a double-cone external compression supersonic inlet at Mach 1.8. The supersonic inlet geometry comprises three zones: a supersonic zone, a transit zone, and a subsonic zone. The supersonic zone is designed numerically using the Taylor–Maccoll method, while the transit and subsonic zones are designed using a methodology from literature. A three-dimensional structured mesh is generated by using the ICEM computational fluid dynamics software. Detailed three-dimensional simulations are performed using the ANSYS Fluent 18.2 code. The grid independency tests are performed by using local grid refinements and varying the number of grid points from 2.5 to 9.5 million. Further, the Fluent code is validated by comparing predictions with those using the SUPIN and Wind-US codes. The effects of side gust on the flowfield and oblique and normal shocks are analyzed by examining the Mach number and pressure distributions in the three zones. The inlet performance is characterized in terms of total pressure loss, flow distortion, and mass flow ratio parameters for a side gust of and gust angles of 30, 60, and 90 deg. Results indicate that the inlet performance is significantly affected by the side gust, especially at small gust angles.

Throttling Process and Buzz Mechanism of a Supersonic Inlet at Overspeed Mode

A rectangular external-compression supersonic inlet with a shock-on-lip Mach number of 2.0 is carefully studied at a freestream Mach number of 2.5. In the throttling process before the onset of buzz, the interaction between the ramp shock and the cowl-induced bow shock appears as regular reflection, Mach reflection, and λ-shaped pattern successively. It is interestingly found that the currently observed Mach reflection, which should have been broken down shortly after the emergence as reported previously, keeps fairly stable until the throttling ratio changes. In the remaining buzz stage, three types of instabilities of similar frequencies but of different oscillating behaviors and amplitudes are observed. The term medium buzz is adopted to describe the special buzz phenomenon with moderate oscillations. It is shown that both the little buzz and big buzz are characterized by the separation oscillation, whereas the medium buzz features the oscillation of the cowl-induced bow shock. Further analysis indicates that all three buzz phenomena probably share a common origin (i.e., the ramp-side flow separation). Also, it is demonstrated that the transition of buzz flows in the throttling process is attributed to the changes of disturbance strength and dominant role between two feedback loops.

Vortex Generators in a Two-Dimensional External-Compression Supersonic Inlet

Vortex generators within a two-dimensional external-compression supersonic inlet for Mach 1.6 were investigated to determine their ability to increase total-pressure recovery, reduce total-pressure distortion, and improve the boundary layer. The vortex generators studied included vanes and ramps. The geometric factors of the vortex generators studied included height, length, spacing, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated through the computational solution of the steady-state Reynolds-averaged Navier–Stokes equations on multiblock structured grids. The vortex generators were simulated by either gridding the geometry of the vortex generators or modeling the vortices generated by the vortex generators. The inlet performance was characterized by the inlet total-pressure recovery, total-pressure distortion, and incompressible shape factor of the boundary layer at the engine face. The results suggested that ramps were not effective, and only vanes positioned downstream of the terminal shock were effective in reducing distortion and improving the boundary layer. In the comparison study of the Reynolds-averaged Navier–Stokes simulations to the experiment, microvanes were used to reduce boundary-layer separation completely. Of the geometric factors studied, the height of the vortex generator had the greatest effect.

Buzz Flows in an External-Compression Inlet with Partially Isentropic Compression

A rectangular external-compression supersonic inlet with partially isentropic compression is experimentally studied at a freestream Mach number of 2.0. During the test, the inlet operates at the little buzz state after going through the supercritical state and the steady subcritical state. A striking feature of the little buzz state is that both the most upstream and downstream positions of the unsteady separation point of the forebody boundary layer take on a linear variation trend with the increase of throttling ratio, yielding an intensifying flow oscillation. Then, the mixed buzz phenomenon that the little and big buzz coexist appears. As the throttling ratio further rises, the occurrence probability of little buzz is reduced, and finally the big buzz state takes place. In this state, two unique transient flow patterns with the inflow and supersonic backflow coexisting at the duct entrance can be observed. Different from the prevailing viewpoint, the present little buzz belongs to the Dailey instability and shares a common origin with the big buzz in essence. Furthermore, a theoretical analysis based on Fisher’s criterion indicates that a single shear layer from a -shaped terminal shock with an upstream Mach number below 1.60 is unable to trigger the Ferri instability of a supersonic inlet.

Simulations of a low-boom, axisymmetric, external compression inlet

AbstractComputational simulations have been used to study a new low-boom, axisymmetric, external compression supersonic inlet with an on-design Mach number of 1·7. The inlet incorporates a relaxed compression surface with a near-zero cowl angle to help reduce external oblique shock waves due to spillage and cowl geometry. To reduce mechanical complexity the inlet is designed with zero-bleed. To understand the impact on performance and shock overpressure caused by the inlet itself, several throat and diffuser designs were simulated. The computations utilised a Reynolds-averaged Navier-Stokes code. Inflow properties were held consistent with the operational characteristics of the NASA GRC 8′ × 6′ Supersonic Wind Tunnel (SWT) for experimental testing of the inlet of Mach 1·67 at a Reynolds number of 5·4 × 106. Stagnation pressure recovery performance for the baseline condition exceeded 94% at design mass flow rates, and reduced only slightly with increases in Mach number (consistent with theoretical predictions) and extension of cowl position. The simulations also showed that the relaxed compression surface combined with the near-zero cowl angle helps to significantly reduce external oblique shocks This is partly due to the reduced inlet spillage in combination with a reduced overall turning angle placed on the free-stream flow relative to the cowl shape.

Normal shock boundary layer control with various vortex generator geometries

Various vortex generators which include ramp, split-ramp and a new hybrid concept “ramped-vane” are investigated under normal shock conditions with a diffuser at Mach number of 1.3. The dimensions of the computational domain were designed using Reynolds Average Navier–Stokes studies to be representative of the flow in an external-compression supersonic inlet. Using this flow geometry, various vortex generator concepts were studied with Implicit Large Eddy Simulation. In general, the ramped-vane provided increased vorticity compared to the other devices and reduced the separation length downstream of the device centerline. In addition, the size, edge gap and streamwise position respect to the shock were studied for the ramped-vane and it was found that a height of about half the boundary thickness and a large trailing edge gap yielded a fully attached flow downstream of the device. This ramped-vane also provided the largest reduction in the turbulent kinetic energy and pressure fluctuations. Additional benefits include negligible drag while the reductions in boundary layer displacement thickness and shape factor were seen compared to other devices.

카울 위치변화에 의한 초음속 공기흡입구의 버즈억제

An experimental study was conducted at a Mach number of 2.0 to investigate the buzz suppression method on an axisymmetric, external compression supersonic inlet. The inlet model has a fixed geometry with no internal contraction. The inlet configuration was altered by changing the cowling. Results show that source of buzz has been related to the existence in the flow field of velocity discontinuity across a vortex sheet which originates from a shock intersection point. With external compression inlet, buzz can be suppressed by positioning the oblique shock slightly inside or outside of the cowl.