The operation of a propulsion system in terms of horizontal takeoff/landing and full-speed range serves as one of the main difficulties for hypersonic travelling. In the present work, a three-dimensional inward-turning inlet with tri-ducts for combined cycle engines is designed for the operation of three different modes controlled by a single rotational flap on the compression side, which efficiently simplifies the inlet structure and the flap control mechanism. At high flight speed between Mach 4 and 6, the pure scramjet mode is switched on, whereas both the ejector and the scramjet paths are open for a moderate Mach number between 2 and 4 with a larger throat area guaranteeing the inlet startability. In the low flight speed range with Mach number below 2, the additional turbojet path will be turned on to supply air for the turbine engine, whereas the other two paths remain open for spillage. Numerical simulations under different operation modes have proven the feasibility and good performance of the designed inlet, e.g., a nearly full mass flow ratio and a total pressure recovery around 0.5 can be achieved at the cruise speed. Meanwhile, the inlet works properly at low flight speeds which overcomes the typical starting problem of similar inlet designs. In the near future, wind tunnel experiments will be carried out to validate our inlet design and its performance.

For centuries, people have dedicated to achieve higher and higher flight speeds. Imagining a flight trip from Shanghai to New York within only two hours is incredibly promoting. Recently, hypersonic travelling is getting intensive attention from both commercial and scientific communities. The difficulty however lies mostly on the propulsion system when considering its requirement of both horizontal takeoff/landing and full-speed range operation. In this case, only combined cycle engines can satisfy the rigorous expectation since any single engine faces deficiency,

The design of the three-dimensional inward-turning inlet needs to meet the mass requirement within the full speed range of a combined cycle engine. Regarding the full mass flow ratio feature of the inlet at the cruise speed, the inflow capturing area can be established directly. The scramjet flow path can be designed with a preferred intake and exit shape based on the consideration of inlet startability as well as efficient compression. This is an iterative process in consideration of the performance of the inlet, especially its startability which needs to be verified carefully based on viscous simulation, since currently an accurate analytical tool is still lacking. But very soon, an efficient in-house code to evaluate the inlet startability will be addressed by our group based on the Kantrowitz theory and the isentropic limit [

For hypersonic inward-turning inlet design, a basic flowfield with high performance needs to be identified as the first step. A number of our previous researches have accumulated different design methods as well as optimization for the basic flowfield [

(a) Mach number distribution of the basic ICFD flowfield. (b) Geometry of the tri-duct inward-turning inlet.

Based on the above basic flowfield, the three-dimensional scramjet flow path can be conducted with the technique of streamline tracing and osculating theory [

By rotating the flap around the hinge on the compression surface to the position 2, the ejector flow path will be fully open. The cross-section of the entrance of the ejector flow path keeps a rectangular shape restricted by the flap sweeping. The entrance area, which determines the rotation angle of the flap, is computed based on the mass requirement of the ejector engine. The ejector flow path of the inlet is designed as a straight tube from the flap end to the exit plane, i.e., the exit has the same rectangular shape without contraction or expansion. Further optimization of the ejector flow path can be modified by a one-dimensional theoretical analysis tool which is under construction by our group currently.

Similar to the ejector flow path, the entrance cross-section of the turbojet flow path must be restricted as a rectangular shape due to requirement of further flap rotation. The exit shape, however, has to be circular to fit the geometrical requirement of the turbine engine. Therefore, a variable cross-section passage from rectangular to circular shape is designed for the turbojet flow path with a weak expansion ratio of 1.25. Since the turbine engine has been selected in advance (

In the present work, the targeted aircraft is a drone where the takeoff weight is of unit ton and the lift/drag ratio at the cruise speed is about 4.5. The combined cycle engine is designed for the aircraft from horizontal takeoff to hypersonic cruise with three basic working modes: turbojet mode from takeoff to Mach 2, ejector mode from Mach 2 to Mach 4, and scramjet mode from Mach 4 to cruise Mach 6,

Working modes of the inlet.

In order to prove the inlet design methodology, simulations were conducted based on the commercial software Ansys Fluent 14.0©. Density-based steady Reynolds-averaged Navier-Stokes (RANS) equations were solved, in which the two equation ^{st}-order upwind scheme is used to discretize the equations spatially to get a stable initial flowfield, based on which the 2^{nd}-order upwind scheme is further applied for a refined resolution. The fluid is treated as ideal gas, with a constant specific heat 1006.43 J/(kgK) and a constant viscosity 1.789e-5 kg/(ms). For all simulations in the present work, the Courant number is chosen as 1, and the convergence is judged by a residual decrease of at least three orders as well as a stable mass flow rate.

To check the reliability of the numerical methods mentioned above, a validation case based on experiment results from [

(a) Configuration of the hypersonic inlet model GK01 extracted from [

Only half of the inlet model is used for simulation in consideration of its symmetric configuration. Structured grids based on different inlet geometries under various working modes are built with Ansys ICEM 14.0©. The near-wall

Structured grid for the tri-duct inlet.

Summary of the flight conditions for the different working modes.

Case | Density (kg/m^{3}) |
Pressure (Pa) | Re (×10^{6}) |
Grid (×10^{6}) | |
---|---|---|---|---|---|

Single duct | 6 | 0.025 | 1616.17 | 2.52 | 3.6 |

4.5 | 0.046 | 2971.69 | 3.06 | ||

Dual ducts | 4 | 0.064 | 4047.42 | 4.24 | 6.8 |

3 | 0.112 | 6994.78 | 5.54 | ||

Tri-ducts | 2 | 0.246 | 15327.56 | 8.11 | 8.6 |

1.2 | 0.651 | 46415.57 | 13.8 |

The inlet is mounted ventrally on the hypersonic vehicle in the present work, i.e., the cowl lip is in the bottom part, as indicated in Figures _{0}) of around 2.4. The isobaric surfaces above the inlet leading edge and that under the cowl lip are due to the wall thickness which leads to additional external shocks. The initial shock hits the cowl lip and reflects inside the inlet with a pressure ratio

Flow contour of the inlet at _{0}, whereas the three isobaric surfaces, respectively,

Flow contour of the inlet at _{0}, whereas the three isobaric surfaces, respectively,

Flow contour of the inlet at _{0}, whereas the two isobaric surfaces, respectively,

Flow contour of the inlet at _{0}, whereas the two isobaric surfaces, respectively,

Flow contour of the inlet at

Flow contour of the inlet at _{0}, whereas the two isobaric surfaces, respectively,

Flow contour of the inlet at _{0}. The symmetry slice is contoured by the Mach number, accompanied with streamlines showing mass flow distribution into different paths.

Mass-averaged performance of the single-duct inlet under both design and off-design conditions.

Mass flow ratio | Total pressure recovery | Exit Mach number | Exit pressure ratio | |
---|---|---|---|---|

6 | 1 | 0.61 | 2.96 | 20.83 |

4.5 | 0.88 | 0.66 | 2.09 | 18.75 |

To guarantee the reliability of the inlet system, the single-duct inlet must work properly at low Mach numbers. Figure

When the operating Mach number decreases below 4, the flap rotates outwards and consequently, the ejector duct is switched on. Figure

By further decreasing the inflow Mach number to 3, the strength of the initial shock is even decreased to

Mass-averaged performance of the dual-duct inlet at

Mass flow ratio | Total pressure recovery | Exit Mach number | Exit pressure ratio | ||
---|---|---|---|---|---|

4 | E | 0.51 | 0.63 | 2.60 | 3.89 |

S | 0.32 | 0.76 | 2.91 | 3.19 | |

3 | E | 0.40 | 0.65 | 1.97 | 2.74 |

S | 0.34 | 0.84 | 2.06 | 3.59 |

Compared to the previous single-duct inlet, the throat area of the dual-duct inlet is twice larger, which guarantees a better startability. As mentioned in Section

For inflow Mach number below 2, the turbojet duct is switched on by further rotating the flap to a larger angle, which also indicates a stronger Prandtl-Meyer expansion around the hinge. The leakage between the initial shock and the inlet cowl lip is further enlarged, spilling around 40% of the captured mass flow. The strength of the initial shock is decreased to a pressure ratio of only 1.25. Due to low air speed inside the inlet, shock reflections are weaker than those in the previous analysis. But obvious shock reflections with pressure ratio of 2.5 can still be observed in Figure

The tri-duct inlet is also simulated at the inflow Mach number 1.2, with a flow contour shown in Figure

Mass-averaged performance of the tri-duct inlet at

Mass flow ratio | Total pressure recovery | Exit Mach number | Exit pressure ratio | ||
---|---|---|---|---|---|

2 | T | 0.06 | 0.48 | 0.56 | 2.93 |

E | 0.32 | 0.84 | 1.52 | 1.68 | |

S | 0.21 | 0.89 | 1.60 | 1.61 | |

1.2 | T | 0.06 | 0.76 | 0.53 | 1.47 |

E | 0.25 | 0.90 | 0.84 | 1.38 | |

S | 0.15 | 0.89 | 1.16 | 0.92 |

The exit Mach number distribution of the turbojet flow path at both

To meet the requirement of three combined engines under a wide speed range and meanwhile keep robust startability and performance, we had to warily design the inlet. The mass flow ratio of the entire inlet and the mass distribution among different ducts are summarized for different modes in Figure

(a) Mass flow ratio of the inlet (with total and each single duct) at different Mach numbers. (b) Total pressure recovery of the inlet at different Mach numbers.

The mass-averaged total pressure recovery of the three ducts is also summarized in Figure

A three-dimensional tri-duct hypersonic inward-turning inlet has been designed for combined cycle engines in the present work, with only a single rotational flap on the compression side, which significantly simplifies the inlet structure and the flap control mechanism in a quite efficient way. The performance of the tri-duct inlet is analyzed based on three different working modes. Under the scramjet mode, only the high-speed duct is open, i.e., it works as a pure scramjet inlet which obtains a nearly unity mass flow ratio and a total pressure ratio of 0.61 at the cruise speed. At Mach 4.5, this single-duct inlet also exhibits nice startability and performance. For the dual-duct mode, however, the biggest challenge lies on the inlet start in a wide speed range. The throat area is therefore warily determined, which is limited nearly identical to the exit area in order to guarantee the startability. For the tri-duct mode, over 40% of the captured mass flow has been spilled out of the inlet at Mach 2 due to the weak initial shock with a pressure ratio of only 1.25. The mass distribution of the air is about 1 : 4 : 2.5 into the three flow paths with respect to turbojet, ejector, and scramjet duct. We also addressed that the mass distribution among different flow paths of similar combined cycle inlets must be carefully determined to meet the requirement of the entire propulsion system, including thrust balance, startability, and performance. In the future, more efforts will be put on the mass distribution analysis as well as validation experiments of similar multiduct inlet.

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.

The authors thank the National Natural Foundation of China (No. 51606161), Fundamental Research Funds for the Central Universities (No. 20720170055), Soft Science Foundation of Fujian Province (No. 2017R0099), and Jiangsu Province Key Laboratory of Aerospace Power System for the financial support of this project.