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Controlling and directing the boundary layer on the surfaces of a flight vehicle are two of the most demanding challenges in advanced aerodynamic designs. The design of highly integrated and submerged inlets with a large offset between the entrance and compressor face is particularly challenging because of the need for controlling or reducing the adverse effects of the boundary layer on propulsive efficiency. S-duct diffusers are used widely in flight vehicles when the compressor face needs to be hidden, and their performance is generally sensitive to the quality of ingested boundary layer from the fuselage. Passive or active flow control mechanisms are needed to prevent flow separations at the bends. In this paper, a new method is presented for optimal inlet/body integration based on a pair of ridges ahead of the inlet and its effects on the performance of a semicircular S-duct inlet integrated on a flat surface using CFD. In this design, the ridge changes an inefficient inlet concept to one with acceptable performance. The new method of integration is practicable for top-mounted inlet configurations where the use of diverters and other mechanisms results in higher amounts of drag, weight, and complexity.

Highly integrated inlet concepts have become an important issue in UAV design studies. The main challenge of inlet/body integration is to prevent the thickened boundary layer on the fuselage from entering into the inlet. Low kinetic energy boundary layer flow and its degradation of inlet performance are particularly problematic. Swirling, flow, distortion, and shock/boundary-layer interactions are common adverse phenomena, related to the nature of boundary layer in airbreathing propulsion systems [

Top-mounted inlets on different UAVs.

General Atomics, Avenger [

Northrop Grumman, Global Hawk [

Boeing, X-47 [

EADS, Barracuda [

Conceptual UAV with ridge/inlet configuration.

The basic structure of the ridge is shown in Figure

Symmetric ridge

Cross-section

When this geometry is subjected to high-speed flow, three different pressurized surfaces appear on it: a high-pressure surface (HPS) resulting from surface C1 (in Figure

The boundary layer (BL) streamlines passing over the ridge cannot escape from the vortex. The interaction between the vortex and the BL on the central surface draws the low energy parts of the flow into the low pressure on the inboard side of the ridge. The combination of vortex and steep pressure gradient creates a powerful trap for capturing and diverting the BL from a uniform straight path into a vortical flow pattern along a preset direction. Controlling the vortex based on the aerodynamic characteristics of the ridge is the unique property of this shape. The continual shear layer separation at the ridge constantly adds mass into the vortex structure and decreases the vortex strength. Therefore, the expanding LPS along the oblique line traps the vortex by increasing the height of the ridge along the oblique line. Figures ^{−5}. The capability of the ridge to completely capture the boundary layer has been demonstrated by accurate simulations from the low subsonic to low supersonic (

Transversal total pressure contours show the vortex structure and boundary layer diversion [

Dense group of streamlines released from upstream at

Information on advanced concepts based on this design can be found in [

The candidate inlet for integrating with the ridge configuration is a pitot inlet with a semicircular entrance. Such inlets are becoming common in different flight vehicles, although their integration is not straightforward. In some cases, the curvature of the inlet limits the use of diverters. This problem is more critical for top-mounted inlets, and the designs are limited to a serpentine configuration or aft inlet/diverter configuration. The most important factor involved in this design is aerodynamic performance sensitivity. Bended inlets generally suffer from internal aerodynamic problems such as secondary flow and flow separation after the first bend of the diffuser. Previous research has shown that the internal aerodynamic performance factors of these diffusers are very sensitive to the resulting flow pattern of the entrance geometry [

Figure

Highly integrated inlet with and without ridge on a flat surface, used in CFD simulations.

Highly integrated inlet

Highly integrated inlet with ridge concept

Sectioned view of ridge/inlet integration

The vortex formed by the rollup of the boundary layer must pass to the side of the inlet entrance. The distance between the ridge and entrance provides this passage. Although the diameter of the ridge vortex is smaller than that generated by a wing strake or canard, it must not enter the inlet. Analysis of flow over the ridge at high yaw angles at low subsonic speeds is required to estimate the required gap between the inlet entrance and the ridge. In this paper, CFD simulation results of inlet integration without a ridge have been used to produce a baseline for measurement of the comparative aerodynamic performance of the new concept.

In the previous study [

The numerical simulations have two primary objectives. The first is to investigate the boundary layer redirecting ability of the ridge with the existence of the inlet entrance on the central surface, especially at high subsonic speeds. Understanding the effects of inlet pressure gradient on the ridge vortex and the stability of the vortex core is the key issue for producing a baseline for aerodynamic testing. The second and most important objective is to investigate the effects of the ridge on the internal flow quality of the S-duct.

Analyzing the boundary layer interactions and related phenomena inside the inlet depends on the abilities of turbulence equations in the numerical solver. An SST turbulence model has been used widely in simulations related to viscous interactions. The model is as economical as the k-^{−5}.

The governing equations for the present problem are the three-dimensional Navier-Stokes equations. The compact conservation form for these equations are as follows [

Continuity equation:

Momentum equation:

Energy equation:

_{l} + _{t} is the total viscosity;

Turbulent kinetic energy

Turbulent dissipation rate

Specific dissipation rate

The turbulent viscosity,

The Fluent solver uses a control-volume-based technique to convert a general scalar transport equation to an algebraic equation that can be solved numerically. This control-volume technique consists of integrating the transport equation about each control volume, yielding a discrete equation that expresses the conservation law on a control-volume basis. More information about solver structure may be found in [

In these equations,

A schematic of the computational domain with the boundary conditions is shown in Figure

Schematic of computational domain and boundary conditions.

The density-based solver in ANSYS Fluent solves the governing equations of continuity, momentum, and energy simultaneously as a set, or vector, of equations. Governing equations for additional scalars will be solved sequentially (i.e., segregated from one another and from the coupled set). The selected algorithms for solving the coupled set of equations is the implicit formulation [

Example of acceptable residual accuracy for ridge/inlet configuration.

Number of iterations | Continuity | Energy | |||||
---|---|---|---|---|---|---|---|

20,000 | 8.251 × 10^{−4} |
3.384 × 10^{−6} |
1.079 × 10^{−5} |
3.917 × 10^{−4} |
4.112 × 10^{−5} |
2.328 × 10^{−4} |
4.062 × 10^{−5} |

For a reliable simulation, a fully structural grid has been created around the ridge and space around the inlet. The

Perspective view of surface grid for inlet without ridge profile.

Inlet without ridge

Inlet with ridge

Cross-section view of grid structure around the ridge.

Close-up view of the entrance.

Diffuser duct mesh structure. The plane of symmetry is omitted in this picture.

Perspective view of diffuser duct

Outlet mesh (AIP)

Grid dependency is evaluated using different mesh structures. The main cases for which results are considered in this paper contain more than 6.38 million cells for the ridge/inlet configuration. Lower density meshes (more than 4.0 million) can also reach the desired accuracy if the boundary layer block remains dense. The pressure distribution on a cross-section of the duct is selected for comparison. Figure

Residual condition at the accepted accuracy.

Main grids in CFD simulations.

Grid | Cell number (million) | |
---|---|---|

M1 (method 1) | 0.91 | 128,679.9 |

M2 (method 1) | 2.47 | 118,892.1 |

N1 (method 1) | 4.26 | 117,486.3 |

N2 (method 1) | 5.26 | 117,579.1 |

N3 | 5.81 | 117,581.8 |

N4 (results are presented in the current paper, method 1) | 6.38 | 117,589.6 |

N5 (different internal mesh structure) | 6.48 | 117,585.3 |

Table

Variation of total pressure for fine grids.

The effect of the grid domain on the resultant pressure distribution is shown in Figure ^{−6}.

Pressure distribution on a longitudinal cross-section at lower surface of diffuser.

Grid study

Effect of solver discretization method

For the dense grid domain (N4), changing the solver discretization algorithm from the first-order to second-order upwind (after 20,000 iterations) does not result in significant changes in the internal pressure distribution after reaching 10^{−6}. This is illustrated by comparing the pressure distribution on the centerline of the lower diffuser wall as shown in Figure

Effect of discretization scheme on the vortex structure.

First-order upwind algorithm

Second-order upwind algorithm

For the inlet configuration without upstream ridges, simulations show that the duct flow contains turbulence and three-dimensionalities even at low subsonic speeds (

The vortex structure is indicated by the total pressure contours in Figure

Perspective view of ridge/inlet with total pressure contours at different sections

Streamlines close to the flat surface, released from upstream,

As was expected, the inlet entrance is completely protected from the boundary layer developed from upstream. On some parts close to the vortex region on the central surface, the boundary layer is swept by the vortex. Oil patterns in Figure

Oil pattern in black and the vortex in white color, _{∞}

Flow quality at the inlet entrance is an important attribute for ridge/inlet integration. The passage between ridge and inlet should let the vortex pass away safely, and any vortex breakdown should happen after this passage. Figure _{∞}

Close-up view of total pressure contours before and at the passage, _{∞}

Top view of the ridge/inlet configuration, _{∞}

Downstream of the ridge, the vortices that it generates provide a small amount of additional lift, and the rear part of the ridge plays an important role in controlling the vortex structure. In contrast to [

Apart from the CFD flow visualization, the flow quality can be measured by the total pressure ratio (

Vertical total pressure ratio distribution.

_{∞}

_{∞}

Another interesting difference between the concepts is the captured mass flow rate. Mass flux calculations at the AIP show that the captured mass flow ratio in the transonic mode is much higher for the ridge/inlet configuration. The calculation results are summarized in Table

Duct flow parameters.

0.3 | 0.991 | 0.991 | 0.034 | 0.351 | ~0.0 |

0.5 | 0.981 | 0.985 | 0.040 | 0.239 | 5.1 |

0.8 | 0.979 | 0.989 | 0.110 | 0.262 | 12.7 |

Black streamlines at _{∞}

Flow pattern at _{∞}

When the high-speed layers of the flow pass over the vortices, they are guided downward into the central surface and inlet. This kind of inlet feeding is still under study in our research.

With the effective diversion of the boundary layer before the intake entrance, the secondary flow separation problem has been solved for subsonic to transonic speeds. The ridge captures the low kinetic energy layers of the flow close to the surface, so that the fresh flow spills into the entrance. The thinner boundary layer after the ridge is more stable and can resist the adverse pressure gradient inside the duct. This issue can be studied by the shape factor (

The oil pattern in Figure _{∞}_{∞}

Shape factor and flow separation coordinate.

0.3 | 3.501 | 2.341 | 0.156 | — |

0.5 | 3.068 | 2.481 | 0.177 | — |

0.8 | 3.804 | 2.663 | 0.203 | 0.139 |

Mach contours on the plane of symmetry at _{∞}

Inlet without ridge

Inlet with ridge

Mach contours on the plane of symmetry at _{∞}

Inlet without ridge

Inlet with ridge

Close-up view of tangential streamlines on the plane of symmetry at _{∞}

Inlet without ridge

Ridge/inlet

Mach contours at different sections at _{∞}

Inlet without ridge

Ridge/inlet configuration

Mach contours at different sections at _{∞}

Inlet without ridge

Ridge/inlet configuration

A comparison between the AIP total pressure and Mach number distributions in Figures

Total pressure and Mach contours at the AIP, _{∞}

Total pressure and Mach contours at the AIP, _{∞}

Tangential streamlines in Figures

The numerical solution was analyzed at the AIP with respect to overall performance factors such as total pressure recovery (

In these equations, the

The results in Table

AIP calculations.

Total pressure ratio variation with Mach number

Captured mass flow rate at the AIP

Although the inlet with a uniform large lip radius profile in this analysis is not designed for supersonic applications, in this section, the effect of ridge on the shock structure of the inlet is discussed based in the CFD results. The Mach number for the simulation is 1.50, and ambient pressure is 100,000.0 Pa.

By applying back pressure (^{5} Pa, a three-dimensional bifurcated shock appears in front of the inlet. For the pitot inlet, the detached normal shock is important because it can result in strong vortex/shock interactions. Figure

Perspective view of the detached normal shock (bifurcated shock) at the entrance.

Close-up view of the vortex passing through the shock wave.

The ridge can be used with a supersonic pitot inlet when the inlet contains very smooth curvature, sharper lips, and long length to let the separated flow reattach to the surface. When the normal shock is placed at the lip (critical mode), the vortex/shock interactions are minimized. Figure

Cone compression surface for ridge/inlet configuration.

The effects of the ridge concept on the performance of a highly integrated inlet are investigated in this paper. The ridge structure contains neither a bleed system nor movable parts, and it can be integrated with different shapes of inlets. CFD simulations show that ridge can redirect the boundary layer based on two aerodynamic phenomena: pressure gradient on its surfaces and vortices. The combination of pressure reduction and vortices creates an effective trap to redirect low energy boundary layer at different Mach numbers. According to these CFD simulations, the basic ridge/inlet configuration shows a practicable passive solution over a wide range of subsonic to low supersonic Mach number. In the current design, the central surface does not play a significant role in boundary layer diverting. Adding an extra compression geometry or bump surface and redesigning the entrance for supersonic conditions may improve the diverting ability and propulsion efficiency. Results of simulations show that the new concept provides a higher value of pressure recovery in comparison with other submerged or highly integrated inlet combinations. By effective diversion of the upstream boundary layer, ridge provides air with higher kinetic energy and thinner boundary layer for the inlet which prevents or reduces boundary layer separation at speeds up to the high subsonic region. The inlet has also been considered for supersonic study. According to these results, the vortex has a weak interaction with the bifurcated shock wave in front of the entrance. The shock/vortex interaction is located at the shock root which is attached to the boundary layer, and the vortex core leaves the ridge profile without entering into the inlet.

Boundary layer

_{θ}

Distortion coefficient based on

Shear stress transport

Aerodynamic interface plane

Total energy

Static pressure

_{b}:

Back pressure

_{t}:

Total pressure

_{r}:

Turbulent Prandtl number

_{t∞}:

Freestream total pressure

_{t-θ}:

Sectorial mean total pressure at AIP

_{AIP}:

Mean dynamic pressure at AIP

Total pressure recovery coefficient

_{i,j}:

Stress tensor

Mach number

_{f}:

Increment of mass-flow ratio coefficient

_{∞}

Freestream Mach number

Boundary layer shape factor

_{t}:

Total enthalpy

Normal distance from the surface

_{z}:

Ratio of the distance from plane of symmetry to the width of the inlet

Turbulence kinetic energy

_{A}:

Averaged turbulence kinetic energy at AIP

Viscosity coefficient

_{l}:

Laminar viscosity coefficient

_{t}:

Turbulent viscosity coefficient

Turbulence dissipation rate

_{l}:

Length ratio of separation point

Counter variables

_{1},

_{2},

_{3}:

Velocity components in x, y, and z directions

Density

_{k}

Production term of

_{k}

Dissipation term of

_{k}

Diffusion term of

_{ω}

Production term of

_{ω}

Dissipation term of

_{ω}

Diffusion term of

_{h}

Source term vector

Internal energy

_{s}:

Static enthalpy.

The data used to support the findings of this study are available from the corresponding author upon request.

The authors declare that they have no conflicts of interest.

Inlet team members in the College of Power and Energy of NUAA are gratefully acknowledged for their cooperation. This research is supported by National Basic Research Program of China (No. 2014CB239602) and Natural Science Foundation of China (No. 11372134).

The flow pattern of the ridge surface is evaluated by a series of wind tunnel tests in NUAA. The vortex structure and its stability are studied. No breakdown or vortex core distortion was visualized during the low subsonic tests. The stability of vortex in low subsonic flow was our main concern during ridge/inlet integration.