Turbine-based Combined Cycle Engine Research Articles

Thrust-matching and optimization design of turbine-based combined cycle engine with trajectory optimization

Abstract Based on the trajectory optimization method of the Gauss pseudospectral, an aircraft/engine matching method is established for the turbine-based combined cycle (TBCC) engine. For a horizontal-takeoff hypersonic aircraft designed at Mach 5, a thrust-matching analysis of the TBCC engine is performed, and a rocket is integrated for further optimization design. The results show that the aircraft for boost missions should adopt the TBCC thrust with a takeoff thrust-to-weight ratio of 99.8 % to reduce the acceleration time and fuel consumption. In contrast, due to the low thrust-to-weight ratio of the TBCC engine, a high-thrust TBCC increases the inert weight in the cruise phase. Therefore, the aircraft designed for cruise missions should adopt the takeoff thrust-to-weight ratio of 92.0 %. Introducing a rocket whose maximum thrust is 10 % of the takeoff weight could assist the aircraft in overcoming the problem of the “thrust pinch” during the transonic and mode transition. With the assistance of rockets, the optimal takeoff thrust-to-weight ratio is 65.3 % for cruise aircraft, and the cruise range is increased by 18 %. While for the boost aircraft, adopting an optimal TBCC of 86.8 % takeoff thrust-to-weight ratio, the introduced rocket could reduce the fuel consumption and the TBCC engine weight by 4 %.

Research on the flow and heat transfer characteristics of RP-3 and structural optimization in parallel bending cooling channels

The regenerative cooling channel is curved due to the unique structural form of the combined nozzle of the turbine-based combined cycle engine for hypersonic flight. Research on the flow and heat transfer characteristics of fuel in the parallel bending cooling channels is insufficient, which causes difficulties in properly designing the thermal protection structure of the combined nozzle. This paper numerically investigated of the flow and heat transfer characteristics of hydrocarbon fuel RP-3 under non-uniform heating conditions in parallel bending cooling channels. The basic heat flux is 0.50 MW/m2, and the heat flux ratio of the upper and lower channels is 1.2, 1.5, and 2.4 respectively. The heat transfer and flow distribution mechanisms of RP-3 in parallel bending channels are revealed. The bending structure is beneficial to reduce thermal stratification and prevent heat transfer deterioration caused by buoyancy. In the pseudo-critical region of the channel, a “thermal insulation zone” is formed due to the low thermal conductivity of RP-3, which hinders the transfer of heat to the mainstream and causes heat transfer deterioration. When the fluid temperature is below 455 K, viscosity is the primary factor affecting the flow resistance difference between parallel channels. As the temperature further increases, the density difference and velocity difference between the parallel channels dominate in the flow resistance differences. Additionally, a connected hole structure in the middle of the bending section has been applied to optimize the flow deviation of parallel bending channels, which promotes the secondary distribution of flow. Finally, the influence mechanism of the connected hole on flow distribution has been studied under different heat flux ratios and connected hole widths.

Multi-stage rotors assembly of turbine-based combined cycle engine based on augmented reality

To improve the assembly accuracy and efficiency of turbine-based combined cycle engines, we propose a multi-stage rotors assembly method based on augmented reality (AR) for turbine-based combined cycle engines. Combining the high fidelity interaction and edge computing capabilities of AR device, the real-time assembly guidance of multi-stage rotors of turbine-based combined cycle engine is realized. A new multi-stage rotors assembly mathematical model is proposed by introducing clearance errors to achieve prediction of coaxiality and perpendicularity in multi-stage rotors assembly. A point cloud registration method with equal-area constraints is proposed based on the requirement of AR virtual-real registration, combined with the triangulated graph. The experiment selected the multi-stage rotors of a simulated combination cycle engine for assembly. The experimental results showed that our AR system has an average delay of 11 ms and an average frame rate of 45fps. The accuracy of virtual-real registration is superior to current mainstream registration methods. The coaxiality prediction error and the perpendicularity prediction error for multi-stage rotors assembly is 0.5 μm and 0.4 μm, respectively. This system provides an efficient and accurate assembly plan for the rotors assembly of turbine-based combined cycle engines, meeting the actual assembly requirements on site.

Modeling and performance analysis of a pre-cooling and power generation system based on the supercritical CO2 Brayton cycle on turbine-based combined cycle engines

Turbine-based combined cycle (TBCC) engines are a promising system of hypersonic vehicles, the main problem being the thrust gap during propulsion. To solve this problem, we proposed a pre-cooling system of turbofan based on the supercritical carbon dioxide (sCO2) Brayton cycle for TBCC engines fueled by hydrocarbon. A pre-cooler was set in the intake of the turbofan engine to achieve coupling between the Brayton cycle and the turbofan engine. The turbofan and Brayton cycle models were established, and several weight and limited cold source parameters were used to evaluate the system's performance. The effects of the inlet air temperature drop and pressure drop caused by the pre-cooler on the engine performance were obtained. The relationship between thermal performance and system weight of the sCO2 Brayton cycle under different pre-cooling target temperatures and recuperated heat loads was also analyzed. The results showed that it is feasible to use sCO2 to pre-cool the inlet air and generate power. When the pre-cooling target was set to Ma 2.2, the simple cycle could output 347 kW power with a power-to-weight ratio of 0.39 kW/kg. This study provides a new scheme for the pre-cooling and power generation technology of hydrocarbon-fueled TBCC engines.

Analysis and Suppression of Thrust Trap for Turbo-Ramjet Mode Transition with the Integrated Optimal Control Method

An aircraft/engine integrated optimal control method is proposed for turbine-based combined cycle (TBCC) engines based on the Gauss pseudospectral method. The optimal flight trajectory and TBCC control law are obtained for a TBCC-powered aircraft, and the “thrust trap” that occurs during turbo-ramjet mode transition is further analyzed and suppressed. Results show that the aircraft goes through the mode transition phase using a “climb-dive” trajectory, which is a strategy of applying gravity-assist and temporarily reducing the drag. Furthermore, the TBCC engine adjusts at the quickest rate to minimize thrust loss. With the coupling of the trajectory and TBCC control law, the minimum thrust during the mode transition is only 23% of the thrust before the mode transition, suggesting the “thrust trap” phenomenon. By decreasing the mode transition time from 60 s to 15 s, the minimum thrust can only increase to 30%, and the “thrust trap” phenomenon cannot be effectively suppressed. When the operating speed range of the turbine engine increases from Ma2.5 to 2.9, the minimum thrust will reach 80%, and the “thrust trap” tends to level off.

The Performance of a Thermal Protection System for the Accessories of a TBCC Engine

A theoretical model for describing the heat transfer characteristics of a turbine-based combined cycle (TBCC) engine cabin was established in Matlab/Simulink to quickly predict the thermal protection performance for engine accessories. The model’s effectiveness was verified by comparing the numerical results with the experimental data. The effects of different heat insulation layer thicknesses and fuel temperatures on the thermal protection performance are discussed; based on these effects, the heat insulation layer of 5 mm and fuel of 353 K were chosen to design the thermal protection cases. Nineteen different thermal protection cases were proposed and evaluated by using the model. Two representative accessories were chosen for the evaluation of the thermal protection performance of these cases. For accessory 1 with an internal heat source of 1000 W and internal fuel access, the thermal protection effect of adding a heat insulation layer and ventilation was the best, which decreased the accessory temperature by 43 K. For accessory 2 without an internal heat source, the thermal protection effect of adding a heat insulation layer to the casing and fuel cooling was the most ideal, which decreased the accessory temperature by 190 K. In addition, a comprehensive assessment was made to compare the performances of thermal protection cases.

Analysis and manipulation of the separation zone in an overexpanded combined exhaust nozzle

The combined exhaust nozzle is an important component of the over-under turbine-based combined cycle engine. During the subsonic and transonic flight Mach number, the ramjet duct suffers severe overexpansion, resulting in complex separation phenomena and poor performance. In the present paper, the fluidic injection is applied to improve the severely overexpanded nozzle's performance based on the idea of separation-zone manipulation. The separation patterns in a ramjet duct are numerically investigated, and the position of the separation point is found to be critical to performance. Separation that originates near the throat produces a longer pressure plateau and raises the nozzle's thrust, while cowl-side separation leads to a better thrust angle compared to ramp-side separation. Based on this phenomenon, the fluidic injection is introduced to manipulate the separation zone to occur at proper locations. The results show that the fluidic injection with a secondary pressure ratio of 2.0 can successfully induce a favorable separation pattern over a large range of working conditions. The average thrust coefficient is increased by 7.3%, and the thrust angle is reduced by 34.8%. The separation-zone manipulation concept is also verified using a 3D nozzle model which shows a similar benefit.

Optimal Control of TBCC Engines in Mode Transition

This paper mainly studies the optimal control problem of turbine-based combined cycle (TBCC) engines in the mode-transition stage. Based on the TBCC scheme proposed by Xiamen University, an aerothermodynamic model is established as a verification model for the validity of control laws. To reduce the complexity of control design, a control-oriented linear parameter-varying (LPV) model with Mach number as a scheduling variable is established under a given flight path. The design of mode-transition points and distribution of air-flow-rate among paths during the mode-transition process are transformed into linear quadratic (LQ) optimal control problems for an LPV system under the initial and terminal as well as process constraints. By optimizing the opening of the splitters of the inlet and the fuel flow in each channel, the optimal mode-transition points are found to achieve coordinated control and complete the high-precision thrust tracking during the mode-transition process.

Flight/Propulsion Integrated Control of Over-Under TBCC Engine Based on GA-LQR Method

Turbine-based combined cycle (TBCC) engines are one of the ideal powers for reusable air-breathing supersonic aircraft, but the flight/propulsion integrated control and mode transition restricts its use. This paper takes the Mach 4 over-under TBCC engine as the research object. The inlet is established by the quasi-one-dimensional calculation theory, which can reflect the shock wave position. An iterative method is proposed, which points out that the flow rate in the mode transition depends on the flow capacity. By connecting the input and output that affect each other, the simulation of the coupling characteristics of the aircraft and engine are realized. A GA-LQR-based controller design method is proposed and verified through the aircraft’s climb and mode transition conditions. The simulation shows that the integrated control system can ensure the stability of the aircraft and the safe operation of the engine in the above two situations. During the mode transition process, the aircraft altitude and Mach number fluctuate less than 1%, and the normal shock wave of inlet is in a safe position.

Effect of Water Injection on Turbine Inlet under Different Flight Conditions

Numerical simulations were conducted to research the pre-cooling effects of water injection on the turbine inlet of a turbine-based combined cycle (TBCC) engine under different flight conditions. Then, the performance of the water injection pre-compressor cooling (WIPCC) engine was calculated by mathematical modelling with different water to air ratios (WAR). It was the first time that the mass injection field of the turbine inlet of a TBCC engine was simulated, and it was also the first time that the performance of a subcomponent turbine engine of a TBCC was assessed. The calculation results showed the relationship of the inlet temperature with respect to WAR, inlet length and flight Mach number. The strategy for inlet length and water mass flow was proposed in order to meet the requirements of pre-cooling. When the length of the turbine inlet was 10 times the diameter of the inlet exit, the air could be cooled by 167.5 K with WAR = 0.09. The highest evaporation ratio reached at 93%. Finally, the calculation results revealed the performance of the water-pre-cooled turbine engine, of which the flight envelope was expanded to Ma3.0 from Ma2.3 by pre-cooling. The engine thrust as well as the specific impulse were significantly improved. The thrust reached at 0.9 times the characteristic thrust meeting the TBCC mode transition requirement of thrust and working speed spectrum.

Performance analysis and dynamic optimization of integrated cooling and power generation system based on supercritical CO2 cycle for turbine-based combined cycle engine

Turbine-based combined cycle (TBCC) engines are the most promising propulsion systems for the horizontal takeoff and landing of hypersonic vehicles. However, the high air-inlet temperature, engine-wall temperature, and electrical-energy demand hinder the application of the system. In this study, a novel integrated cooling and power generation system based on a supercritical CO2 recuperative Brayton cycle for hypersonic vehicles was proposed to meet the cooling requirements of each operating stage of a TBCC engine and to provide continuous power. Several optimization parameters able to balance the characteristics of the limited cold source, system weight, and power output of the system, such as the power-to-weight ratio and heat sink saving ratio, were proposed. With the help of optimization algorithms, the global performance curves and loads of each component were obtained during flights in the range of Mach 3.2∼6. The recuperator heat duty and compressor inlet temperature were key design parameters that affected system performance. The compressor inlet temperature is known to alter the heat sink utilization by changing the pinch point location in the gas cooler. The recuperator heat duty impacted both the system weight and power generation. The dynamic performance of the system indicated that the minimum heat sink saving ratio was approximately 10% during the entire flight, implying that the volume of fuel carried by the hypersonic vehicle could be reduced to some extent, while sufficient electricity was generated. The proposed system is a novel solution for thermal protection of TBCC engines and aircraft power generation technology under limited cold-source conditions.

Mode transition analysis of a turbine-based combined-cycle considering ammonia injection pre-compressor cooling and variable-geometry ram-combustor

Mode transition (MT) is a significant process in the turbine-based combined-cycle (TBCC) involving the co-operation of high-speed and low-speed channels. To improve the thrust pinch caused by the deviation of the turbine and ramjet from the design point, a novel TBCC engine with ammonia mass injection pre-compressor cooling (Ammonia MIPCC) and variable-geometry ram-combustor (VGRC) is used to extend the operation envelope of the MT. CFD simulations and shock loss calculations are used to construct combined dual-channel inlet model. Through simulation, we find that the Ammonia MIPCC can reduce the pre-compressor temperature, improve the compressor efficiency, and increase the turbine's specific thrust. Meanwhile, increasing the ramjet combustor expansion ratio (CER) helps to alleviate the inlet unstart and extend the low-speed operating boundary of the ramjet. Furthermore, the extension of the operating envelope for TBCC under the effects of ammonia injection and CER is investigated. Based on specific thrust and specific impulse, the optimal starting Mach number field for MT process is studied, and the corresponding TBCC thrust is given. Overall, the results show that the optimal starting Mach number field for the MT of the proposed TBCC is 2.85–3.2, and two-stage sub-engines have high thrust performance in this range.

Parametric Research and Aerodynamic Characteristic of a Two-Stage Transonic Compressor for a Turbine Based Combined Cycle Engine

This paper researches the parametric optimization of a two-stage transonic compressor having a large air bypass at partial rotating speed according to flow analysis for a turbine-based combined cycle engine (TBCC). To obtain adequate thrust, the inlet transonic compressor of the turbofan part of the TBCC is required to have a wider frequently used corrected rotating speed range and a larger mass-flow rate at low rotating speed, which is different from a typical transonic compressor. The one-dimensional blade design parameters and flow path of the baseline two-stage transonic compressor are introduced. With the widely used CFD software Numeca, the three-dimensional flow fields of the baseline transonic compressor and effects of the flow path between Stage 1 and Stage 2 on the inlet mass flow rate are analyzed for indicating the further improvement direction. For design speed (NC = 1.0), to improve the efficiency at the design point, parametric research is carried out on Rotor 2 to optimize the shock structure and strength, resulting in enhanced efficiency at the design point due to reduced shock loss of Rotor 2. For partial speed (NC = 0.8 and 0.7), since the flow field analysis indicates that the flow blockage in S1 limits the entire mass flow rate, the parametric redesign of stator S1 aims at obtaining an increased blade throat width to enhance the flow capacity of S1. Simulation confirms the increase in the mass-flow rate and efficiency at partial speed due to the reduction in flow blockage and related viscous losses. Aerodynamic analysis at representative operation points indicates that the modifications of R2 and S1 lead to obvious aerodynamic improvement at all rotating speeds (NC = 1.0 to 0.7), while maintaining sufficient stall margin.

A free radical relay combustion approach to scramjet ignition at a low Mach number

Turbine-based combined cycle (TBCC) engines, which consist of turbine and scramjet engines, are a key power system for the acceleration, powered horizontal landing, and reuse of hypersonic vehicles. Low Mach number ignition of kerosene-based scramjet fuel is expected to enable the practical application of TBCC engines. However, ignition under these conditions is extremely difficult to achieve. In this study, free radical relay combustion (FRRC), motivated by multi-catalyst relay catalysis, was proposed to solve the problem of scramjet ignition at low Mach numbers. The ignition temperature of modified kerosene was significantly reduced from 620 to 150 °C, and the hydroxyl radicals (OH·) signal value of the modified kerosene at 210 °C was found to be 538 times higher than that of pure kerosene at 680 °C. The FRRC-based ignition of the modified kerosene presented a flat droplet evaporation. This significantly reduced the ignition delay time of the fuel to 3 ms at 220 °C, which was comparable to or better than that of hydrogen. In addition, the FRRC strategy described in this study achieved direct-connect supersonic combustor ignition at an inflow of Ma 1.5. This study provides a reference for the design of scramjet fuel for low Mach number ignition.

Modeling and analysis of windmilling operation during mode transition of a turbine-based-combined cycle engine

Mode transition control is a critical issue of Turbine-Based-Combined Cycle (TBCC) engines when the primary thrust providers change from gas turbine engines to ramjets/scramjets. During mode transition, windmilling operation is inevitable when the fuel mass flow rate of the turbojet engine is decreasing to zero gradually. Existed research on mode transition control of TBCC engines usually neglect this process or simplify this process to a rapid thrust loss due to lacking a windmilling model. However, the absence of the windmilling model might cause critical problems unconsidered at the control system design stage and increase the risk of mission abortion. This paper provides a method of modeling the windmilling operation of a TBCC engine. Sub-idle compressor component maps are obtained by the interpolation method based on a zero-speed line to consider the characteristics that the total pressure ratio is below one. Friction torque is added in the torque balance relationship to achieve the free windmilling operation state when both the compressor and the turbine provide positive torques. The inlet start/unstart status coupled with the operating point of the turbojet engine is also considered. Results show that this model could reflect the characteristics that the total pressure ratio of the compressor is below one, and it provides a positive torque instead of a negative one when the turbojet engine operates at some sub-idle states. Moreover, inlet unstart would happen at some point when the turbojet engine is shutting down to the free windmilling state, under which the turbojet engine would bring in the flight drag instead of a positive thrust. In the descent mode transition when the turbojet engine is accelerating to be restarted, an inappropriate rotating speed of the inlet splitter could result in a fail inlet restart.

Efficiency Analysis and Integrated Design of Rocket-Augmented Turbine-Based Combined Cycle Engines with Trajectory Optimization

An integrated analysis method for a rocket-augmented turbine-based combined cycle (TBCC) engine is proposed based on the trajectory optimization method of the Gauss pseudospectral. The efficiency and energy of the vehicles with and without the rocket are analyzed. Introducing an appropriate rocket to assist the TBCC-powered vehicle will reduce the total energy consumption of drag, and increase the vehicle efficiency in the transonic and the mode transition. It results in an increase in the total efficiency despite a reduction in engine efficiency. Therefore, introducing a rocket as the auxiliary power is not only a practical solution to enable flight over a wide-speed range when the TBCC is incapable but also probably an economical scheme when the the TBCC meets the requirements of thrust. When the vehicle drag is low, the rocket works for a short time and its optimal relative thrust is small. Thus, the TBCC combined with a booster rocket will be a more simple and suitable scheme. When the vehicle drag is high, the operating time of the rocket is long and the optimal relative thrust is large. The specific impulse has a significant impact on the flight time and the total fuel consumption. Accordingly, the combination form for the rocket-based combined cycle (RBCC) engines and the turbine will be more appropriate to obtain higher economic performance.

An improved model to calculate radiative heat transfer in hot combustion gases

Radiative heat transfer plays an important role in the chemical reactions in the combustor. The widely used WSGG model proposed by Smith is established for normal pressure, which shows inevitable computational errors when dealing with radiative heat transfer problems at reduced or elevated pressures. In this paper, an improved global model is established to calculate the radiant energy exchanges between combustion gases and combustor chamber walls. Compared with the Smith model, the new model shows better performance in a wide range of pressure regions. The model accuracy is examined by computing the emissivity, radiative heat flux as well as the radiative source of H2O–CO2 gas mixtures at different pressure values. Finally, the radiative heat transfer inside a 3D TBCC(turbine-based combined cycle) engine exhaust system where strong gradients of pressure and temperature exist, is also addressed. The computational results show that the developed model provides approximate results at much less computational costs than the high-precision MSMGFSK-c8 model, which makes it competitive in complicated combustion systems.

Analysis on the Low Speed Performance of an Inward-Turning Multiduct Inlet for Turbine-Based Combined Cycle Engines

The multiduct inlet for turbine-based combined cycle engines receives a lot of attention on its aerodynamic performance. Aside of the most studied mode of transition processes, another significant severe issue regarding the aerodynamic performance of the turbine duct (T-duct) at ground states has rarely been investigated which indeed directly determines the operability and reality of similar engine systems; this issue will be addressed in the present work. Our numerical and experimental studies of an inward-turning tetraduct inlet indicate that the performance of the T-duct is seldom affected by the angle of attack, which however is of crucial importance for takeoff/landing of flight vehicles. The two T-ducts exhibit weak asymmetrical aerodynamic performance during experiment due to nonsynchronization as well as mechanical oscillation of the two turbine engines. With increasing inflow speed, the surface pressure and the total pressure recovery increase accordingly. At Ma∞=0.24, the total pressure recovery achieves 0.96 at the exit of the turbine duct which is acceptable for the engine to generate sufficient thrust for horizontal takeoff. A further quantitative comparison between simulation and experiment reveals a maximum deviation of only 3% in terms of both surface pressure and total pressure recovery.

Control-oriented low-speed dynamic modeling and trade-off analysis of air-breathing aerospace vehicles

We present a control-oriented low-speed dynamic modeling and trade-off study framework for a conceptual air-breathing horizontal take-off and horizontal landing (HTHL) aerospace vehicle, which is powered by a turbine-based combined cycle engine. First, the 3D class/shape transformation method is modified to enhance the continuity property between different blocks, combined with the power function. Then, the panel method based on potential theory is employed to calculate the pressure distribution over discretized panel surfaces, resulting in the aerospace vehicle’s aerodynamic coefficients. To overcome the intractability of the physics-based model, stepwise regression analysis is adopted and simplified polynomials of aerodynamic coefficients are evaluated. Finally, stability and control analysis is conducted, aiming to find the proper center-of-gravity locations under different constraints. The proposed framework is verified through a conceptual aerospace vehicle simulation, with emphasis on horizontal take-off rotation and landing nose hold-off capabilities. Simulation results indicate that the proposed framework is capable of rapid control-oriented dynamic modeling and iterative design of HTHL aerospace vehicles.

Modeling and analysis for integrated airframe/propulsion control of vehicles during mode transition of over-under Turbine-Based-Combined-Cycle engines

Mode transition control is a critical issue of Turbine-Based-Combined Cycle (TBCC) engines when the primary thrust provider changes from gas turbine engines to ramjets/scramjets. Improper control laws might incur unexpected propulsion performance and even lead to mission abortion. Existed control laws are designed from the perspective of the combined engine itself without aerodynamic/propulsive couplings considered. However, like scramjet-powered aircraft, those with TBCC engines propelled also have intricate couplings between the engine and the airframe, especially during TBCC mode transition. In this paper, a nonlinear longitudinal TBCC-powered dynamic vehicle model is derived from first principles with extra aerodynamic effects and adjustable surfaces involved when a typical TBCC engine is integrated into the airframe. Then, a control-oriented model with six control inputs is obtained by approximately expressing aerodynamic coefficients through curve-fitting methods. Trim results indicate that the total thrust required to maintain a steady flight increases at different steady stages during mode transition of the TBCC engine. One feasible control law is designed with dynamics of the integrated system considered for a typical mode transition process to reveal a way that control inputs should be regulated to maintain a steady flight during mode transition of the TBCC engine.