Turbine-based Combined Cycle Research Articles

Multi-fidelity simulation of inlet mode transition with smooth thrust of turbine based combined cycle

A multi-fidelity simulation method of external and internal flow has been developed using commercial software to achieve a smooth thrust of the turbine-based combined-cycle (TBCC) propulsion system. This platform enables the investigation of the flow field and performance of the TBCC propulsion system at different mode transition schemes. The integrated multi-fidelity simulation includes the inlet, turbojet engine, ramjet engine, and nozzle, providing insights into the operation process of the TBCC propulsion system during the transition from turbojet mode to ramjet mode. Three mode transition schemes are proposed: critical mode transition, constant aerodynamic interface plane (AIP) Mach number mode transition, and linear mode transition. From the perspective of TBCC inlet, the performance of the critical mode transition exhibits the best performance among these mode transition schemes, while the linear mode transition performs the worst. However, from the viewpoint of the TBCC propulsion system and the hypersonic vehicle, the performance of constant AIP Mach number mode transition is the best. The non-dimensional thrust increases almost linear from 1.0 to 1.2, enabling hypersonic vehicle to accelerate steadily during the transition from turbojet mode to ramjet mode.

Analysis of the impact of key similarity criteria numbers of TBCC inlet during the mode transition

Abstract A simplified configuration was developed to facilitate the mode transition process within an over-under Turbine-Based Combined Cycle (TBCC) inlet. Leveraging dynamic mesh technology, an unsteady numerical simulation of the mode transition was conducted, emphasising the flow characteristics of the mode transition and the impact of key similarity criteria numbers. The findings indicate that at an incoming Mach number of 2.0, the mode transition is paired with a continuous alteration in the capture mass flow of the high-speed duct. This continual change instigates the inlet unstarting, with subsequent flow characteristics being contingent on the historical effect, exhibiting a degree of hysteresis characteristics. When the scale effect is considered, it is observed that a larger model scale results in higher Reynolds (Re) and Strouhal (St) numbers. This directly contributes to a notable delay in the unstart moment, a decrease in the unstart interval, and an enlargement of the hysteresis loop. An examination of control variables reveals that the Re number marginally influences mode transition characteristics, while the St number’s effect constitutes approximately 90% of the scale effect. This conclusively demonstrates that the St number is the predominant similarity criterion number in the mode transition process.

Enhancing Ignition Reliability with Tail-Groove Strut Designs in Cavity-Based Combustors

As the flight envelopes of turbine-based combined cycle (TBCC) engines expand, ensuring reliable ignition and flame stability under varying conditions becomes increasingly critical. Previous work has shown a significantly changed ignition performance and flow pattern in cavity-based combustors when strut structure parameters were altered, indicating a strong correlation between the ignition process, flame structure, and the strut configuration. This suggests that further investigation is required to determine the optimal strut design. Therefore, this study examines the impact of various strut configurations through numerical simulations, validated by high-speed imaging. Findings show that the tail-groove strut designs improve the flame propagation performance compared to the normal struts, with a critical depth beyond which further increases do not enhance performance. Changes in strut length have a lesser impact than depth. Flow analysis indicates that tail-groove struts create additional recirculation zones that enhance fuel atomization and flame stability. These results suggest that optimizing strut configurations is vital for achieving reliable ignition and flame stability, advancing the development of efficient engines across a wide range of operational conditions.

Energy and exergy analyses on the high-pressure bleeding-air variable cycle designed for a wide-speed-range airbreathing propulsion system

The performance of the traditional turbojet decreases sharply with the increase of the Mach number in the supersonic flight, especially for the turbine-based combined cycle (TBCC) engine in the transition region of the turbojet and scramjet engine. This paper proposes a new high-pressure bleeding-air variable cycle engine (HB-VCE) concept that aims to ease the sharp contradiction between thrust output and fuel consumption of TBCC in the transition region. Through the energy and exergy analyses, the HB-VCE shows greater potential in energy utilization under high Mach flight conditions. The performance of HB-VCE significantly improves as the bleed ratio increases, with a more pronounced effect at higher Mach numbers. At Mach 1.5, the HB-VCE increases the specific thrust (ST) by 2.84 % and reduces the specific fuel consumption (SFC) by 2.73 %. At Mach 2.4, it increases the ST by 6.60 % and reduces the SFC by 6.52 %. Additionally, the controlling law of the critical bleeding ratio constrained by the turbine outlet temperature is proposed. Finally, the effect of the turbine inlet temperature, overall pressure ratio, and Mach number on the critical bleeding ratio is analyzed. This paper provides a promising HB-VCE concept with greater energy utilization for the TBCC design in the transition region.

Investigation of the impact of splitter rotation speed on mode transition characteristics of an over-under turbine-based combined cycle inlet

Investigation of the impact of splitter rotation speed on mode transition characteristics of an over-under turbine-based combined cycle inlet

Simple model of turbine-based combined cycle propulsion system and smooth mode transition

Abstract Turbine-based combined cycle (TBCC) propulsion system is becoming one of the most promising propulsion systems for two-stage-to-orbit reusable launch vehicle. Mathematic model of this combined cycle engine is helpful for the basic understanding of performance analysis of this propulsion system. We developed mathematic model of TBCC propulsion system based on C++ platform in this paper. Firstly, turbojet engine was built on component level and ramjet engine was calculated through stream thrust function. Then, performance of turbine-based combined cycle propulsion system along a specific flight trajectory was investigated. According to the thrust of this combined cycle engine, mode transition point was suggested at Ma 2.5, which may achieve smooth mode transition from turbine mode to ramjet mode. Finally, mode transition based on smooth mass flow and smooth thrust criteria were studied. The thrust gap arises during smooth mass flow mode transition, particularly when the turbojet engine afterburner is powered off at the start of the mode transition, and it meets 34 % of the total thrust. Smooth thrust mode transition may be achieved by delaying the turbojet engine afterburner power off and injecting additional fuel into ramjet burner.

Unsteady flow characteristics in an over-under TBCC inlet during mode transition under unthrottled and throttled conditions

The study presents an experimental exploration into the mode transition of an over-under TBCC (Turbine-Based Combined Cycle) inlet, with a specific emphasis on the flow characteristics at off-design transition Mach number. A systematic investigation was undertaken into the mode transition characteristics in both unthrottled and throttled conditions within a high-speed duct, employing high speed Schlieren and dynamic pressure acquisition systems. The results show that the high-speed duct faced flow oscillations primarily dictated by the separation bubble near the duct entrance during the downward rotation of splitter, leading to the duct’s unstart under the unthrottled condition. During the splitter’s reverse rotation, a notable hysteresis of unstart/restart of the high-speed duct was observed. Conversely, hysteresis vanishes when the initial flowfield nears the critical state owing to downstream throttling. Moreover, the oscillatory diversity, a distinctive characteristic of the high-speed duct, was firstly observed during the mode transition induced by throttling. The flow evolution was divided into four stages: an initial instability stage characterized by low-frequency oscillations below 255 Hz induced by shock train self-excitation oscillation and high-frequency oscillations around 1367 Hz caused by the movement of separation bubble. This stage is succeeded by the “big buzz” phase, comprised of pressure accumulation/release within the overflow-free duct and shock motion outside the duct to retain dynamic flow balance. The dominant frequency escalated with the increase of the internal contraction ratio in the range of 280 Hz to 400 Hz. This was followed by a high-frequency oscillation stage around 453 Hz dominated by a large internal contraction ratio with low pulsating energy, accompanied by a continuous supersonic overflow. Lastly, as the splitter gradually intersected the boundary layer of the first-stage compression surface, the capture area and the turbulence intensity of the incoming flow underwent a sudden shift, leading to a more diverse flow oscillation within the duct, manifested as various forms of mixed buzz.

Influence of a double-mixer configuration on the mixing and combustion characteristics of a multibypass combustor for a turbine-based combined cycle engine

The changes in the flow field and combustion characteristics induced by the mode transition process of a turbine-based combined cycle (TBCC) engine may lead to failure of the mode transition. In this paper, the cold mixing characteristics, combustion efficiency, and cold and thermal flow losses of a TBCC multibypass combustor with a double-mixer configuration are investigated through numerical simulations, and the correlations between the thermal mixing efficiency and combustion efficiency is established. The results show that the thermal mixing efficiency of double-lobed mixer combustors is better at low ram inlet Mach numbers, but the ring-plus-lobed mixer combustor has better mixing and combustor characteristics at relatively high Mach numbers. However, the effect of the turbo core inlet Mach number on the combustion efficiency is not regular. The spacing between the double mixers that maximizes the mixing and combustion characteristics of the combustor is 75 mm. With increasing of expansion angle of the lobed mixer, the thermal mixing efficiency and combustion efficiency increase by 31.9 % and 23.3 %, respectively. In addition, the positive correlation coefficient between combustion efficiency and thermal mixing efficiency caused by changes in the inlet Mach number is greater than that caused by changes in structural parameters.

Conceptual design methodology and performance evaluation of turbine-based combined cycle inward-turning inlet with twin-design points

Turbine-based combined cycle (TBCC) engines attract intensive attention due to adequate high working efficiency within the full speed range of air-breathing vehicles. Challenges of TBCC engines lie mostly in common-used components in particular the inlet system, which needs to satisfy the mass flow distribution of different sub-engines and to reconcile the performance requirement in the full speed range. In the present work, a new conceptual design methodology of the three-dimensional inward-turning TBCC inlet with twin-design points is proposed according to the basic flow field with double incident shock waves. The shock structures and flow field parameters at the high-speed design point Ma=4 and mass flow distribution at the low-speed design point Ma=3 both can be designed by the basic flow field. On that basis, a typical inlet model is designed and the corresponding aerodynamic performance is evaluated at two design points. Simulation results prove that the shock structure, the mass flow distribution, and the averaged throat parameters of the TBCC inlet are all consistent with the basic flow field with sufficient accuracy at twin-design points, where the maximum relative error is only 6.3%. Meanwhile, the aerodynamic performance of the inlet shows that the inlet has stable and efficient operation characteristics, which fully affirms the feasibility of the design method.

Experimental and numerical study on flow mixing and combustion characteristics of a novel multibypass integrated combustor

The integrated design of multibypass augmented/ramjet combustors can reduce the weight of turbine-based combined cycle (TBCC) engines and improve the thrust-to-weight ratio, but low-resistance mixing and efficient stable combustion of multiple airflows over short distances are necessary prerequisites. In this study, a novel structure for a TBCC multibypass integrated augmented/ramjet combustor is proposed. The influence of the inlet aerodynamic parameters on the flow field, mixing efficiency, flow loss, and combustion performance of the combustor under different working modes was obtained via experimental and numerical methods. The experimental results show that the outlet mixing efficiency is greater than 86% in the double-bypass mode (DB-mode). While the triple-bypass mode (TB-mode) has a larger decrease, the total pressure loss is slightly reduced (by approximately 0.5%). The opening of the ram duct has a significant impact on the flow field, resulting in different rules for the influence of the inlet temperature on the outlet mixing efficiency: in the DB-mode, the mixing efficiency decreases with increasing inlet temperature, while the rule is completely opposite after entering the TB-mode. Because the dominant role of the two mixing zones in the flow field changes with the velocity, the inlet velocity has a significant impact on the mixing efficiency in the flow direction. The combustion simulation results show that the combustion efficiency in the DB-mode is almost always above 90% and the high-temperature zone is mainly concentrated downstream of the integrated strut. The radial temperature gradient increases and the combustion efficiency decreases in the TB-mode.

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.

Mode transition control law analysis of ammonia MIPCC aeroengine considering inlet–compressor safety matching

Inlet–Compressor (I–C) safety matching is a key technology for successfully implementing the mode transition of a turbine-based combined-cycle (TBCC) engine. To investigate the performance variation in Ammonia mass injection pre-compressor cooling (MIPCC) aeroengines (AMAs) during the mode transition, a model considering inlet–compressor engine safety matching was developed in this study. We provide a computation for the inlet unstart margin ξ which considers the positive shock wave assumption, I–C flow conservation equation, and combined inlet characteristics to assess the degree of I–C matching. Furthermore, the AMA characteristics indicate that the total pressure recovery coefficient of the hypersonic combined inlet has significantly affects the engine performance. The restraint ξ also ensures engine safety. A control law guided by ξ and the combustor outlet temperature (ξ–COT) was proposed to enhance the specific impulse and specific thrust of the AMA. Compared with the speed and COT control laws, the specific impulse and specific thrust were maximally improved by 2.93 % and 5.47 %, respectively, when ξ = 0.2. Both were further improved by 5.91 % and 10.49 %, respectively, when ξ = 0.1. Finally, the high-thrust performance results of the AMA, with the optimized control law ξ–COT during the mode transition process, were obtained.

Optimum Design of a Reusable Spacecraft Launch System Using Electromagnetic Energy: An Artificial Intelligence GSO Algorithm

Due to its advantages of high acceleration, reusability, environmental protection, safety, energy conservation, and efficiency, electromagnetic energy has been considered as an inevitable choicefor future space launch technology. This paper proposes a novel three-level orbital launch approach based on a combination of a traditional two-level orbital launch method and an electromagnetic boost (EMB), in whichthe traditional two-level orbital launch consists of a turbine-basedcombined cycle (TBCC) and a reusable rocket (RR). Firstly, a mathematical model of a multi-stage coil electromagnetic boost system is established to develop the proposed three-level EMB-TBCC-RR orbital launch approach, achieving a horizontal take-off–horizontal landing (HTHL) reusable launch. In order to optimize the fuel quality of the energy system, an artificial intelligence algorithmparameters-sensitivity-basedadaptive quantum-inspiredglowworm swarm optimization (AQGSO)is proposed to improve the performance of the electromagnetic boosting system. Simulation results show that the proposed AQGSO improves the global optimization precision and convergence speed. By using the proposed EMB-TBCC-RR orbital launch system and the optimization approach, the required fuel weight was reduced by about 13 tons for the same launch mission, and the energy efficiency and reusability of the spacecraft was greatly improved. The spacecraft can be launched with more cargo capacity and increased payload. The proposed novel three-level orbital launch approach can help engineers to design and optimize the orbital launch system in the field of electromagnetic energy conversion and management.

Intelligent ammonia precooling control for TBCC mode transition based on neural network improved equilibrium manifold expansion model

Precise control of the aeroengine precooling temperature is important for safeguarding the mode transition (MT) in turbine-based combined cycle (TBCC) propulsion. Ammonia mass injection pre-compressor cooling (Ammonia MIPCC) technology increases the maximum operating Mach number of aeroengines, presenting challenges in the design of controllers for complex control objectives. This paper introduces the combined inlet characteristics and the precooling scheme in the MT and proposes a neural network-improved equilibrium manifold expansion (NNEME) model. An intelligent ammonia precooling controller for the ammonia MIPCC aeroengine (AMA) was established based on NNEME and extended state observer (ESO). Furthermore, the final error confidence intervals for the NNEME model show a significant reduction of 37.499 % for rotor speed and 37.919 % for compressor outlet temperature, compared to the conventional EME model. By implementing ESO-based online compensation for modeling uncertainties, we achieved remarkable improvements in the control results. Specifically, the maximum steady-state errors for rotor speed and compressor outlet temperature were reduced by 60.839 % and 51.529 %, respectively, when compared to control results without ESO integration. Therefore, the proposed NNEME + ESO-based controller enabled precise control of the AMA precooling temperature during the MT phase. This demonstrates the effectiveness of our approach in enhancing control accuracy and stability.

Automatic Landing System Design for Unmanned Fixed-Wing Vehicles via Multivariable Active Disturbance Rejection Control

Landing control of unmanned aerial vehicles (UAVs) is challenging because of the strong nonlinear dynamics, multivariable, model uncertainties, wind variations, and sensor noise. Motivated by this fact, this paper investigates an automatic landing system (ALS) that includes trajectory generation and guidance law for the first flight test of a turbine-based combined cycle technology demonstrator. Specifically, the control scheme increases the original model’s order to generate a reasonable monotone-decreasing throttle reference flare trajectory by the pseudospectral method. Subsequently, the guidance law based on innovative multivariable active disturbance rejection control is designed to robustly track the reference altitude and velocity simultaneously with high accuracy. The multivariable extended state observer (ESO) incorporated decoupling algorithm enhances the estimation capability and accuracy of potential problem in cross-coupling dynamics compared to the traditional ESO. It is proven that the closed-loop error dynamic has bounded-input bounded-output stability and an explicit upper bound is given. Numerical simulation verifies that the presented approach has better robustness and higher tracking accuracy for external disturbances and parametric uncertainties than the existing benchmark autolanding controller. Finally, flight tests show that the proposed ALS can land the vehicle effectively and safely under severe wind conditions.

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.

Feasibility and Performance Analysis of High-Energy-Density Hydrocarbon-Fueled Turboexpander Engine

With the in-depth research on hypersonic aerodynamics and hypersonic propulsion technology, humans are growing closer to space travel. Recent studies have shown that the pre-cooled air-turborocket (ATR) or turboexpander engines are some of the potential propulsion methods for reusable space vehicles and single stage-to-orbit (SSTO) missions because they have a high specific impulse at low Mach numbers, which can overcome the problem of the “thrust gap” in turbine-based combined-cycle (TBCC) engines. The ATR engine needs an additional oxidizing agent and the turboexpander engine usually uses hydrogen as fuel, which has low energy density and poor safety. To address this problem, this paper proposed a high-energy-density (HED) hydrocarbon-fueled turboexpander engine, and its feasibility has been proven through a simplified thermodynamic model. Through detailed thermodynamic analysis based on the energy and pressure balance, this paper analyzed the performance characteristics of the engine to evaluate its capacity to work in a wide speed range at low Mach numbers. The results show that the endothermic hydrocarbon-fueled turboexpander engine has good specific impulse in Mach 0∼4 at an equivalence ratio of 0.7∼1.3, and the turboexpander engine can be combined with the dual-mode scramjet and become an efficient acceleration method for SSTO missions and the reusable spacecraft.

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.