Axisymmetric Inlet Research Articles

Numerical investigations of shock/boundary-layer interaction and control for Mach 2.5 flow in an axisymmetric inlet

The present work investigates the application of vortex generators for separation control in axisymmetric isolator flows. Reynolds-Averaged Navier-Stokes computations were performed to simulate a Mach 2.5 flow past a half-isolator geometry with a 20∘ axisymmteric compression ramp based on the experiments of Funderburk and Narayanaswamy at North Carolina State University. Single vortex generator and a vortex generator array placed upstream of the shock-induced separation region were investigated. Near wall streamlines, surface pressure contours and density contours on crosswise planes were compared with experiments for flow control using a single vortex generator. Results indicate that the present computations are able to capture the wake and shock structures, and also predict reduction in the streamwise extent of flow separation downstream of the device trailing edge. Finally, the effects of the shape/orientation (forward facing/backward facing) of a single vortex generator, and the design of an array of three vortex generators (with devices of similar and mixed orientations) on the flow separation were also investigated. Contours of near-surface streamwise velocity showed that device orientation had a strong effect on separation control, which is attributed to the differences in the primary streamwise vortices shed in the two cases. Further, the overall reduction in the footprint of separated flow was determined to be most with the use of an array of vortex generators with mixed orientations.

Design of Reverse Bleed Slot for Curved Axisymmetric Inlet Based on Kantrowitz Criterion and Flow Field Characteristics

Design of Reverse Bleed Slot for Curved Axisymmetric Inlet Based on Kantrowitz Criterion and Flow Field Characteristics

Flow field reconstruction in inlet of scramjet at Mach 10 based on physical information neural network

This paper proposed the physical information residual spatial pyramid pooling (PIResSpp) convolutional neural network that is highly robust and introduces a residual neural network architecture that can satisfactorily fit high-dimensional functions by using jumping connections to reduce the risk of overfitting. Key features of the flow field were extracted by using pooling kernels of different sizes and were then stitched together to fuse its local and global features. The axisymmetric inlet of the scramjet generated by the Bezier curve was established through highly precise numerical simulations, and datasets of flow fields under different geometric configurations were constructed according to the parametric design. The PIResSpp model was trained on a sample dataset, and mapping relationships were established between the parameters of incoming flow/those of the geometry of the inlet, and the velocity, pressure, and density fields in it. Finally, the results of reconstruction of the flow field at the inlet with different design parameters were tested and compared with the outcomes of various deep learning models. The results show that the average peak signal-to-noise ratio of the flow field reconstructed by the proposed model was 36.427, with a correlation coefficient higher than 97%.

Flow Coefficient and Starting Performance Prediction of Variable Geometry Curved Axisymmetric Inlet

With the development of combined cycle engines, it is urgent to estimate more quickly and accurately the flow capture capacity and starting performance of variable geometry inlets over a wide Mach number range. Based on the flow field and parameter fitting, two prediction methods for the curved axisymmetric inlet with lip translation scheme have been proposed. The method based on the flow field of the reference inlet is more efficient than the parameters-based prediction method, as it can accurately predict the lip translation distance and the corresponding flow coefficient over the entire working range of the inlet without additional numerical calculations. Moreover, the starting Mach number is accurately predicted by the fitting method based on the throat Mach number of the reference inlet, with a relative error of only 0.95% compared to the numerical simulation. The flow coefficient-based method is simple and accurate for predicting lip translation distances with a known starting Mach number, with a relative error of only 1.65% compared to numerical simulations. The prediction approaches can overcome the drawbacks of the standard iterative algorithms and significantly enhance computational accuracy and efficiency.

Fan Stability Enhancement With Partial Casing Treatments

Abstract Casing treatments are used to extend the stability margin of fans or compressors. These typically have an axisymmetric arrangement, but for a fan exposed to non-axisymmetric distortion, a casing treatment applied to a fraction of the annulus is of interest to target a particular distortion feature. In this paper, a partial casing treatment has been designed with axial-skewed slots, and an experimental investigation has been carried out to determine the effects of the treatment extent and position. Three inlet conditions were tested for each extent of casing treatment: clean, radial tip-low distortion, and half-annulus tip-low distortion. For axisymmetric inlet conditions, the stall margin and efficiency penalty increased approximately linearly with casing treatment extent. However, for half-annulus tip-low distortion, there was a strong dependence between the position of the casing treatment relative to the distorted sector and the stall margin improvement. If the casing treatment is positioned where the rotor enters the distorted sector, unsteady disturbances in the rotor tip flow are minimized. However, if the rotor meets the casing treatment later in the distorted sector, the disturbances are able to grow, leading to earlier stall inception. Overall, this shows that partial casing treatments can target a distorted flow to give a disproportionately positive impact compared to a full annulus casing treatment. For example, it was found that a 72 deg casing treatment extent could be positioned to achieve over 50% of the full annulus casing treatment stall margin improvement with only 20% of the efficiency penalty.

Design and analysis of a supersonic axisymmetric inlet based on controllable bleed slots

Aiming at the problem of starting and inlet-engine matching of axisymmetric variable geometry inlet with complex structure, a novel design concept of axisymmetric inlet is proposed based on controllable bleed slots switch. The laws of the typical geometric parameters of multi-slots effecting on inlet start, supercritical and subcritical performances are numerically studied. Accordingly, a three-dimensional aerodynamic design scheme of the inlet operating at Ma1.5-4 and control laws of the bleed slots are gained, which is slightly altered inlet surface and structure is simplified. The aims of minimum starting Mach number to 2 and inlet-engine matching inlet are achieved by controlling switch of bleeding slots, and inlet operates well at high total pressure recovery and low exit Mach number. The results show that: at Ma=2, the inlet starts with 30% bleeding and 16.5% total pressure recovery coefficient improving; At Ma=3, mass flow of inlet-engine matches with 23% subsonic bleeding but cost of total pressure loss about 2.6%.

Directivity of sound propagation from an commercial supersonic engine inlet

The effects of mean flow variations on sound propagation from an axisymmetric commercial supersonic engine inlet were studied using numerical methods. A finite element model of the inlet was constructed in Ansys Fluent and used to solve for flow fields given by different initial conditions. Results from this model were fed into the aeroacoustic solver, Actran, and used to calculate far field radiated noise as well as the directivity of that noise. The acoustic source of this noise was a plane wave of a known strength placed at the fan face. In addition to assessing the effects of mean flow on the radiated noise transfer functions, the duct modes of the model were compared across different flow regimes. Relationships between mean flow parameters and the directivity of duct modes are developed. The results of this study will be used in further studies to gain a deeper understanding of how the underlying physics which govern the system create favorable or unfavorable directivity patterns.

Parameterization and optimization design of a two-dimensional axisymmetric hypersonic inlet

Optimization method, as a promising way to improve inlet aerodynamic performance, has received increasing attention. The present research is undertaken to design a two-dimensional axisymmetric hypersonic inlet using parametric optimization. The inlet configuration is parameterized and optimized in consideration of total pressure recovery and starting performance. A Pareto front is obtained by solving the multi-objective optimization problem. Then, the flow structures of the optimized inlets are analyzed and the starting performances are evaluated. Results show that the total pressure loss mainly occurs in the internal contraction section, especially near the inlet entrance, and therefore the total pressure recovery coefficient can be greatly improved by decreasing external compression. As a result, the guidance for designing high-performance inlets is concluded. Besides, it is found that as the internal contraction ratio increases, the inlet starting ability becomes worse, which attributes to the larger separation bubble at the inlet entrance. Finally, the total pressure recovery coefficient and the starting Mach number of the optimized inlets are obtained, which can be a reference for engineering design.

Design and aerodynamic performance analysis of a variable geometry axisymmetric inlet for TBCC

Abstract Design and aerodynamic performance analysis of a variable geometry axisymmetric inlet was carried out for tandem scheme turbine-based combined cycle (TBCC) propulsion system. The operation Mach number of the inlet was between 0 and 4. The design point was chosen as Mach number 4.0 in this paper. The determination of external and internal compression and the design method of annular to circle diffuser were illustrated. The inlet was unstart under Ma 3.0 without adjustment. Then, a variable scheme was designed to ensure self-start of the inlet and match the requirement of mass flow rate during the whole flight envelope. And four supports were used to fix the spike. According to the 3D numerical simulations, the total pressure recovery was 0.52 at Ma 4.0 at critical condition and the mass flow rate was consistent with the requirement at different flight Mach number.

Investigation of shock dynamics in an axisymmetric inlet/isolator with attached boundary layers

Abstract

Effects of vortex generators on unsteady unstarted flows of an axisymmetric inlet with nose bluntness

Shock tunnel tests and CFD simulations are conducted to study effects of vortex generators (VGs) on unsteady unstarted flows of an axisymmetric hypersonic inlet with nose blunt radii of 0.8 mm and 3.2 mm. The inlet exit is throttled with a preseted block to generate unstarted flows at a freestream Mach number of 5.9. Unstarted flows of no VGs, 0.5 mm and 1 mm-thick VGs are recorded through schlieren imaging and pressure measurements. Experimental results show that the unstarted flows are characterized with prominent shock oscillations and pressure fluctuations. VGs reduce scales of the external separations, enlarge angles of the separated shock waves, and increase time-averaged magnitudes of surface pressures. The oscillatory period can be shortened from 4 ms to 3.13 ms by 1 mm-thick VGs for the R0.8 mm configuration. The maximum increment of the time-averaged pressure magnitudes can reach 16.2 times the freestream static pressure (p∞) for the R0.8 mm configuration with 1 mm-thick VGs. In addition, VGs generally increase pressure standard deviations of the measuring points upstream of the throat section within 4.0 p∞ and decrease those of the downstream measuring points up to 8.6p∞, despite the 0.5 mm-thick VGs condition of the R3.2 mm configuration. Flow mechanisms dominating the effects of VGs on the unsteady unstarted flows are analyzed with the aid of the CFD simulations. Feasibility of VGs on unstarted flow control is also evaluated.

2D Hypersonic Scramjet Inlet Geometry of Mach 7 using CFD

In the era of space transportations there is a huge demand on space technology to improve on cost reduction and take the heavy loads into space. Thus the load carrying capacities will be increase with this air breathing engines. This work gives a report on the design, analysis of optimal 2D scramjet engine inlet operating at Mach 7 without use of movable geometry. A computational study for scramjet inlet with different ramp angles is carried out. Several cases are considered to compress the air by rounding leading edge without moving the whole cowl up and down, by fixing the cowl lip and assuming axisymmetric inlet with rounded edge. The numerical tests are conducted to obtain maximum total pressure recovery, drag force and outlet Mach number for given flight condition. Two dimensional effects are studied with Navier-strokes approach to compute the pressure and Mach number at a different location. Oblique shock waves, expansion waves and shock wave interactions are primarily focused. Computational Fluid Dynamics (CFD) solver is used; steady flow simulations are carried out for inlet geometries with one, two, three and four ramps. Geometrical shape is redesigned based on oblique shock wave analysis. The corrected model is tested on Fluent with boundary layer considerations that the theoretical analysis is not able to cover. Lastly, a conclusion summarizing the design process is drawn and the optimal model is recommended for the Mach 7 inlet with different ramps with contraction ratio 10. It had been observed that two ramp scramjet inlet model has been preferred to use which has optimum pressure recovery and lower drag.

Validation of a novel 3D flow model for the optimization of construct perfusion in radial-flow packed-bed bioreactors (rPBBs) for long-bone tissue engineering

Osteogenic cell culture in three-dimensional (3D) hollow cylindrical porous scaffolds in radial-flow packed-bed bioreactors (rPBBs) may overcome the transport limitations of static and axial perfusion bioreactors in the engineering of long-bone substitutes. Flow models of rPBBs help optimize radial flux distribution of medium and tissue maturation in vitro. Only a 2D model is available for steady flow transport in rPBBs with axisymmetric inlet and outlet accounting for the fluid dynamics of void spaces, assessed against literature information. Here, a novel 3D model is proposed for steady flow transport in the three compartments of rPBBs with a more practical lateral outlet. A 3D model of transient tracer transport was developed based on the flow model to predict bioreactor residence time distribution (RTD). Model-predicted flow patterns were validated in terms of RTD against tracer experiments performed with bioreactor prototypes equipped with commercial scaffolds for bone tissue engineering. Bioreactors were challenged with a step change in entering tracer concentration in an optimized set-up under conditions promoting uniform radial flux distribution and typical shunt flows. Model-predicted RTDs agreed well with those experimentally determined. In conclusion, tracer experiments validate the use of the 3D flow model for optimizing construct perfusion in rPBBs to engineer long-bone substitutes.

On the question of starting conditions for frontal axisymmetric inlets tested in hot-shot wind tunnels

The work presents the results of an analysis of starting conditions for some frontal axisymmetric inlets of internal compression tested at freestream Mach numbers М = 3−8.4 in the hot-shot wind tunnels based at Khristianovich Institute of Theoretical and Applied Mechanics (ITAM). The results of these inlets test are compared with the data of numerical computations of inviscid, laminar, and turbulent flows carried out by the pseudo-unsteady method. There were determined the inlet throat areas limiting either with regard to the inlet starting or with regard to providing the maximally possible degree of geometric compression of the inlet-captured supersonic airstream at its deceleration in the already started inlet. Reshaping of computed flow patterns in the inlets depending on the variation of the minimal cross section of the inlet internal duct is analyzed.

A Study on Blended Inlet Body Design for a High Supersonic Unmanned Aerial Vehicle

The design process of blended inlet body (BIB) for the preliminary design of a near-space high supersonic unmanned aerial vehicle (HSUAV) is presented. The mass flow rate and cowl area of inlet at a design point are obtained according to the cruise condition of the HSUAV. A mixed-compression axisymmetric supersonic inlet section with a fixed geometry reasonably matching the high supersonic cruise state is created by using the inviscid theory of aerodynamics. The inlet section is optimized and used as a baseline section for the BIB design. Three BIB concepts for the HSUAV are proposed, and their internal aerodynamic characteristics of inlet are evaluated using Euler computational fluid dynamics (Euler CFD) solver. The preferred concept is identified, in which the straight leading edge of the baseline HSUAV configuration is modified into the convex leading edge to accommodate the inlet and meet the requirements of the cowl area to capture the sufficient air flow. The total recovery of inlet for the preferred BIB concept and the aerodynamic characteristics of the modified HSUAV configuration are verified using Navier-Stokes computational fluid dynamics (NS CFD) solver. The validation indicates that the preferred BIB concept can meet both the requirements of the inlet and aerodynamic performance of the HSUAV.

Acoustically Induced Shock Oscillations in a Low-Boom Inlet

Experimental unsteady centerbody surface pressures measured in a low-boom inlet have been analyzed and compared with an unsteady computational flow approach. The experimental dataset was gathered at the supersonic wind tunnel at the NASA John H. Glenn Research Center in 2010. The axisymmetric external compression inlet considered herein featured a relaxed isentropic centerbody compression spike followed by a short subsonic diffuser to the aerodynamic interface plane. The axisymmetric inlet computational domain comprised a 10 deg sector, starting with the freestream inflow region, and included both the internal flowpath up to the mass flow plug and the external flow past the sharp-edged cowl for external flow. The selected inlet test conditions were based on a 1.67 freestream Mach number at a zero angle of attack with a near-design spillage rate of approximately 4%. Both experiments and simulations revealed temporal shifts between pressure peaks at different streamwise locations, indicating upstream-running compression waves that moved at acoustic speeds. These waves became amplified near the geometric throat and produced streamwise oscillations of the external normal shock. The simulations also revealed a second smaller shock on the centerbody at the geometric throat, for which the complex dynamics suggested an opportunity for further study.

Near-On-Design Unsteadiness in a Supersonic Low-Boom Inlet

The data collected in the supersonic wind tunnel at the NASA John H. Glenn Research Center of a low-boom supersonic inlet have been analyzed and compared to computational results generated using a detached eddy simulation methodology based on the Nichols–Nelson model. The objective of this study is to better understand the capability of this methodology and the source of significant external shock oscillations. The external-compression axisymmetric inlet includes a relaxed-compression spike followed first by a short subsonic diffuser to the aerodynamic interface plane that represents the inflow face of an engine and then by a long low-speed diffuser that terminates in the mass flow plug. The flow conditions for the experiments and simulations were based on a Mach 1.67 freestream coupled with 4.5% spillage (higher than the 1.5% on-design spillage condition). A 10 deg flow sector was modeled with a three-dimensional structured grid and solved with the WIND-US code. The time-averaged computational flowfield was not sensitive to including the low-speed diffuser domain, but this extra length was important for unsteady characteristics (especially pressure waves). Computations and experiments showed these waves originated downstream and propagated upstream, initially as weak acoustic waves, but led to strong pressure oscillations due to shock motion, as well as significant complexity in the transonic region near the throat.

A hybrid CFD/characteristics method for fast characterization of hypersonic blunt forebody/inlet flow

A hybrid CFD/characteristic method (CCM) was proposed for fast design and evaluation of hypersonic inlet flow with nose bluntness, which targets the combined advantages of CFD and method of characteristics. Both the accuracy and efficiency of the developed CCM were verified reliably, and it was well demonstrated for the external surfaces design of a hypersonic forebody/inlet with nose bluntness. With the help of CCM method, effects of nose bluntness on forebody shock shapes and the flowfield qualities which dominate inlet performance were examined and analyzed on the two-dimensional and axisymmetric configurations. The results showed that blunt effects of a wedge forebody are more substantial than that of related cone cases. For a conical forebody with a properly blunted nose, a recovery of the shock front back to that of corresponding sharp nose is exhibited, accompanied with a gradually fading out of entropy layer effects. Consequently a simplification is thought to be reasonable for an axisymmetric inlet with a proper compression angle, and a blunt nose of limited radius can be idealized as a sharp nose, as the spillage and flow variations at the entrance are negligible, even though the nose scale increases to 10% cowl lip radius. Whereas for two-dimensional inlets, the blunt effects are substantial since not only the inlet capturing/starting capabilities, but also the flow uniformities are obviously degraded.

Gas-dynamic design of an axisymmetric tunnel air inlet with isentropic compression

A method of designing a supersonic axisymmetric tunnel air inlet based on the problem of an inverted flow in an annular nozzle with isentropic expansion is considered. The nozzle contour is constructed by the method of characteristics. Parameters of one inlet for viscous and inviscid gas flows are calculated.

Numerical modeling of the conditions for realization of flow regimes in supersonic axisymmetric conical inlets of internal compression

The results of the numerical investigation of flow regimes in the axisymmetric inlets of internal compression at a supersonic flow around them are presented in the work. The main attention is paid to the determination of the ranges of the duct geometric convergence, in which a supersonic inflow in the inlet realizes. The investigation has been carried out at high supersonic freestream velocities corresponding to the Mach numbers М = 2−8 by the example of the frontal conical (funnel-shaped) inlets with the angles of the internal cone wall inclination δw = 7.5−15° under the variation of the relative area of the throat cross section. The flow structure alteration was studied and the critical relative areas of the inlet throat were determined, at which either there is no starting of the inlet in the process of flow steadying at the initial subsonic flow in it or a flow breakdown occurs in the process of flow steadying at an initial supersonic inflow. Numerical computations of the axisymmetric flow were done on the basis of the solution of the Navier–Stokes equations and the k-ω SST turbulence model.