Rocket-based Combined-cycle Engine Research Articles

Mixing enhancement assisted by dual plate cavity in supersonic flow

Fast and efficient mixing of high-speed shear flow formed by fuel and oxidizer is of great importance for the improvement of rocket-based combined cycle engine performance. Nevertheless, the existence of compressibility effects of high-speed flow significantly inhibits the growth process of a mixing layer. Moreover, a finite-length combustor of the engine calls for more effective enhanced-mixing strategies to complete mixing in a shorter streamwise distance. To this end, in present paper, the strategy called dual plate cavity (DPC) is proposed to promote mixing. Three cases including the benchmark, front-DPC and back-DPC cases are selected to perform the comparative study. By means of high-order direct numerical simulations, the structure evolution characteristics and turbulence intensity distributions are researched. The index of velocity thickness is utilized to assess the mixing layer growth. The results indicate that with the introduction of DPC, the mixing process is dramatically promoted. The penetration behavior of newly found T-shaped structures into the upper main stream can engulf more fluid into the mixing region. Specifically, in the back-DPC case, the coexistence of both large-scale and small-scale structures in the far flow field can improve the turbulence intensity. The spatial correlation analysis results show that with the influence of DPC, the structure sizes are much larger than that of the benchmark case in the same streamwise position. Meanwhile, the contour line equal to 0.5 possesses property of distortion for the back-DPC case. The drastic pulsation of a mixing layer edge can obviously promote the mixing process. Through exploration of the enhanced-mixing mechanisms, this work indicates that the proposed DPC strategy is a good candidate for efficient mixing, and in the future, more detailed work including three-dimensional simulations concerning the strategy optimization is suggested to be performed.

Numerical study on throttling characteristics of RBCC engines in ejector mode

Investigating the throttling characteristics of Rocket-Based Combined Cycle (RBCC) engines in ejector mode is crucial for adapting to varied flight mission demands. This study employs validated RANS numerical methods to investigate throttling effects on an experimental kerosene-fueled RBCC engine under two representative flight conditions: ground take-off (Mach 0) and supersonic flight (Mach 2.0). Results indicate that, at Mach 0, throttling reduces the entrained secondary stream and increases mixing efficiency, but exacerbates mixing losses, which cannot be offset by the thrust produced by the secondary stream due to its weak working capacity. Therefore, specific impulse performance deteriorates during throttling. Moreover, the optimum mixture ratio remains at 2.4, consistent with conventional rockets. Conversely, at Mach 2.0, throttling mitigates the primary stream’s squeezing effect to augment the entrained secondary stream. Simultaneously, throttling enhances mixing efficiency to improve the utilization of residual fuel. Consequently, the optimum mixture ratio shifts from 2.4 to below 1.4, which contributes to harnessing the work capacity of the secondary stream generated by ram effects, and significantly improves specific impulse performance. This study emphasizes the distinct throttling characteristics at different Mach numbers and provides pivotal insights for developing more efficient RBCC engines in ejector mode.

Numerical Investigation on the Effect of Fuel-Rich Degree in the RBCC Engine under the Ejector Mode

The ejector mode of the Rocket-Based Combined-Cycle (RBCC) engine is characterized by high fuel consumption. This study is aimed at investigating the influence of the rocket fuel-rich degree on the RBCC engine’s performance under the ejector mode combined with simultaneous mixing and combustion (SMC). Numerical simulations were conducted for various rocket mixing ratios (Φ=1.6~3.2) under subsonic (Maf=0.9) and supersonic (Maf=1.8) flight conditions. It was observed that a high fuel-rich degree in the rocket plume negatively impacts the eject performance under all conditions. However, it improves the overall performance (Isp) at high flight Mach numbers (Maf). For supersonic conditions, increasing the fuel-rich degree promotes greater fuel participation in combustion, thereby enhancing RBCC engine performance. Nevertheless, the subsonic-supersonic mixing layer exhibits low evolution, resulting in a decrease in reaction efficiency from 29.2% to 12.0% as the Φ decreases from 3.2 to 1.6. Consequently, there is an inefficient utilization of fuel. To optimize RBCC engine performance, the rocket fuel-rich degree can be appropriately increased. However, this increase should be limited to prevent fuel wastage arising from low reaction efficiency. Under subsonic conditions (Maf=0.9), the low kinetic energy of captured air leads to the occurrence of “negative thrust surface” and “wall impact” phenomena, which hinder the efficient and stable operation of the RBCC engine. Consequently, adjusting the fuel-rich degree alone cannot promote specific impulse (Isp), and a low fuel-rich degree is considered an ideal strategy when combined with adjustable nozzle technology.

Effects of excess oxidizer coefficient on RBCC engine performance in ejector mode: A theoretical investigation

This study investigates the effects of the primary rocket’s excess oxidizer coefficient (EOC) on the performance of a kerosene-fueled rocket-based combined cycle (RBCC) engine operating in ejector mode using the diffuser and afterburning (DAB) cycle. Employing complete mixing and chemical equilibrium assumptions, a quasi-one-dimensional model is developed to simulate combustion processes in the primary rocket chamber, mixer, and afterburner. Experimental data and full three-dimensional simulation validate the model’s accuracy. Results indicate that increasing EOC mitigates thermal choking effects at the mixer exit, decreasing mixing and Rayleigh losses, enhancing the entrainment ratio, and providing more oxygen for secondary combustion. However, the gas mixture’s compression ratio and total temperature decrease, rendering secondary combustion more susceptible to thermal choking in a constant cross-sectional afterburner. As the flight Mach number increases from 0.8 to 2.5, the optimum EOC for maximum specific impulse shifts from 0.725 to 2.375. While the optimum EOC for maximum thrust remains consistently around 2.0. The findings underscore the necessity of a higher EOC to fully unlock the DAB cycle’s thrust augmentation potential.

Experimental and numerical study on RBCC engines performance in simultaneous and combustion cycle

The simultaneous mixing and combustion (SMC) cycle is a simple yet effective approach to augment the specific impulse of rocket-based combined cycle (RBCC) engines in ejector mode. However, the thrust performance of RBCC engines in the SMC cycle has not been adequately explored. This study investigates the effects of the flight Mach number, mixer shapes, and mixture ratio (MR) on the thrust performance of an axisymmetric kerosene-fueled RBCC engine. Ground free-suction experiments were conducted, employing a supersonic air ejector to simulate varying flight backpressures. Wall pressure distributions and axial thrusts were obtained. To eliminate undesirable interactions between MRs and backpressures inherent in experiments, numerical simulations were performed. The experimental results show that as flight Mach number increases, specific impulse improves rapidly, and specific impulse augmentation is achieved in the transonic regime. Simulations conducted at H4.0km/Mach0.88 with the expansion mixer reveal that as MR decreases, both mixing efficiency and combustion intensity increase. The balance between specific impulse gain from the secondary stream and loss from the primary stream is achieved at MR = 2.2, with the highest specific impulse being achieved at MR = 2.0. These insights provide valuable guidance for advancing the development of more efficient RBCC engines using the SMC cycle.

A simplified chemical model for RBCC engines operating in ejector mode

The ejector mode in rocket-based combined-cycle (RBCC) engines involves complex mixing and combustion processes between the primary and secondary streams and second fuel, which poses a challenge to numerical simulation. This paper introduces an efficient model to predict RBCC engine performance in ejector mode. Given the slow growth rate of compressible mixing layers with strong compressibility effects and high chemical reactivity of immediate species within the primary stream, the combustion intensity is primarily governed by the mixing process. The model converts residual chemical energy within the primary stream into suitable mass flow rates of H2 and CO and simplifies the intricate chemical kinetics into single-step models. The model was tested on hydrogen- and kerosene-fueled RBCC engine experiments involving diffusion and afterburning (DAB) and simultaneous mixing and combustion cycles. Simulation results align well with experimental data regarding wall pressure distributions, entrainment ratio, and axial thrust. The model suggests intense combustion occurs in both the mixer and afterburner, even in the DAB cycle, which notably reduces the entrainment ratio. Additionally, early combustion within the mixing process incurs extra Rayleigh loss and impairs pressurization performance. The secondary fuel exhibits a jet-wake stabilized combustion mode under a high-temperature and low Mach number environment. In conclusion, the proposed model provides an efficient way of predicting RBCC engine performance in the ejector mode.

Drag reduction characteristics of recirculation flow at rocket base in an RBCC engine under ramjet/scramjet mode

To reduce the drag generated by the recirculation flow at the rocket base in a Rocket-Based Combined Cycle (RBCC) engine operating in the ramjet/scramjet mode, a novel annular rocket RBCC engine based on a central plug cone was proposed. The performance loss mechanism caused by the recirculation flow at the rocket base and the influence of the plug cone configuration on the thrust performance were studied. Results indicated that the recirculation flow at the rocket base extended through the entire combustor, which creates an extensive range of the “low-kinetic-energy zone” at the center and leads to an engine thrust loss. The plug cone serving as a surface structure had a restrictive effect on the internal flow of the engine, making it smoothly transit at the position of the large separation zone. The model RBCC engine could achieve a maximum thrust augmentation of 37.6% with a long plug cone that was twice diameter of the inner isolator. However, a shorter plug cone that was half diameter of the inner isolator proved less effective at reducing the recirculation flow for a supersonic flow and induced an undesirable flow fraction that diminished the thrust performance. Furthermore, the effectiveness of the plug cone increased with the flight Mach number, indicating that it could further broaden the operating speed range of the scramjet mode.

Combustion organization strategies for a variable geometry RBCC combustor under low total temperature conditions

A rocket-based combined cycle (RBCC) engine experiences a low Mach number phase during flight operations. Through combustor geometry adjustment technology, the engine can combust more efficiently under low-temperature inflow conditions during this phase, thereby improving the engine efficiency. In this paper, we investigate the fuel blending and combustion processes in rocket ramjet and pure ramjet modes under Ma 2.5 low-temperature inflow conditions, analyze the fuel combustion and flow and engine performance changes under different operating modes, and propose a new method for centralized heat release combustion organization under low-temperature inflow conditions. The method is based on the RBCC engine combustor with a geometry throat and relies on both direct-connect experiments and numerical calculations. The results show that (1) when the primary rocket operates, its jet reacts with the inflow in the shear layer and ignites the kerosene fuel injected by the pylons, forming a high-temperature zone downstream of the fuel pylons. When the primary rocket stops working, this flame is stabilized in the low-speed reflux region at the exit of the rocket by injecting a small equivalent ratio of fuel in the isolator to mix with the air. This approach can be a good substitute for the rocket jet to stabilize the flame. (2) When the primary rocket operates, the secondary fuel is mainly concentrated in the area from the fuel pylons to the geometric throat. After the rocket is turned off, the interval of heat release from combustion is also concentrated in the combustor from the fuel pylons to the geometric throat. (3) After the primary rocket stops working, the performance of the combustor improves significantly.

Developing characteristics of subsonic-supersonic shear layers induced by different lobed mixers in the RBCC

The rapid mixing of air and rocket gas is pivotal for enhancing combustion performance in the ejector mode of rocket-based combined-cycle engines. This study experimentally investigated developing characteristics of subsonic-supersonic shear layers induced by sinusoidal- and rectangular-lobed mixers using a subsonic-supersonic shear layer experimental system and an ice-cluster-based planar laser scattering technique. Both sinusoidal- and rectangular-lobed mixers evidently bolstered mixing through the promotion of transverse flow and vortex interactions. Under equivalent conditions, the sinusoidal-lobed mixer offered superior mixing effects. The mixing enhancement of a sinusoidal-lobed mixer exhibited a positive correlation with the wavelength, with excessively small wavelengths potentially impeding mixing. Based on the dimensionless wavelength ratio, a criterion was identified for evaluating the mixing enhancement of the sinusoidal-lobed mixer: a positive mixing enhancement occurred when the dimensionless wavelength ratio was at least 1. Notably, in this study, the shear layer thickness growth rate increased by 12 % when the dimensionless wavelength ratio was 1.905. These findings can offer valuable insights for improving combustion performance via enhanced mixing.

Thermodynamic performance modeling, optimization and numerical simulation of RBCC ejector mode

Ejector mode is a unique and critical phase of wide-range rocket-based combined cycle (RBCC) engine. In this paper, a quasi-one-dimensional thermodynamic performance modeling method, with more detailed model treatments for the inlet-diffuser system, primary/secondary flow interaction, and pressure feedback matching, was developed for operating characteristics studies and multi-objective optimization analysis of the ejector mode of an actual RBCC engine. A series of three-dimensional simulations of separate inlet and full flowpath was completed to validate the modeling study and provide further insight into the operating characteristics. The primary/secondary equilibrium pressure ratio functions a significant effect on ejector mode performance, a higher performance augmentation can be obtained by lower rocket pressure ratio, larger mixing section area ratio, smaller throttling throat and higher equivalence ratio, within an appropriate range. The positive performance augmentation can be realized at low flight Mach conditions, the coordination and trade-off relationships between specific impulse, performance augmentation ratio and thrust-to-area ratio during ejector mode are present by the Pareto-front from MOP analysis. It is further verified by CFD simulation that, the operating back-pressure at the exit of inlet-diffuser system functions a decisive influence on the airbreathing characteristics, the pressure feedback and matching should be well-controlled for secondary flowrate and performance augmentation. The thermodynamic modeling analysis results are basically consistent with those of numerical simulation, to validate the rationality and effectiveness of the modeling method.

Thrust performance of the rocket-based combined-cycle engine under ejector mode

The Rocket-Based Combined-Cycle (RBCC) engine integrates the preponderance of ramjet and rocket engines. It can perform excellently at a lower Mach number than a ramjet while consuming less fuel than the rocket. The higher specific impulse under lower flight Mach conditions guarantees the competitiveness and application prospect of an RBCC engine for reusable space transportation and hypersonic cruise vehicles. With a wide range of Mach numbers, the flow choking between primary and secondary streams in the inner flow passage of the engine becomes complicated. The flow choking not only affects the mass flow ratio between the air stream and rocket plume but also determines the thrust performance of the engine. However, the relationship between flow-choking states and thrust performance has not been revealed. This investigation aimed to provide a better design reference for the RBCC trajectory design and a basis for the RBCC engine geometry design so that the thrust performance under different flow choking states was studied. The findings indicate that the engine is more favorable in the supersonic regime status than others. At the premise of the lower total pressure air stream condition, the thrust of the RBCC engine is attributed to the rocket pressurization effect. On the contrary, its thrust is ascribed to the air stream when the air stream total pressure is higher. Besides, the thrust augmentation of the RBCC engine is essentially due to the larger entrainment ratio. There is a rocket mass flow rate range in different mixer diameters, named region B in this paper, in which the air mass flow rate can be boosted with the rocket mass flow rate and the specific impulse ratio is higher than the rocket specific impulse. Significantly, the RBCC design state should be maintained in region B to meet the operational requirements of engines under the precondition of an insufficient flight Mach number.

Numerical study on the dynamic process of ramjet/scramjet mode transition in the integrated full flow path for RBCC engine

As the inflow velocity increases, rocket-based combined cycle(RBCC) engine typically require a ramjet/scramjet mode transition within the flight range Ma∞ = 5.5 to 6.0. However, during the mode transition process, the reduction of the fuel mass flow rate usually causes the combustion intensity in the engine to drop sharply in a very short time, which will further lead to severe thrust fluctuations. So some technical methods need to be applied in this process to realize smooth mode transition. In this paper, different throttling strategies (kerosene mass flow rate dominant and rocket mass flow rate dominant) were designed and employed to study the dynamic mode transition process in the integrated RBCC engine flow path through three-dimensional numerical simulation for flight Ma∞ = 5.5 and Ma∞ = 6.0, respectively. The results show that when RBCC engine operates under the condition of large mass flow rate of fuel, the wall flow separation zone caused by oblique shock wave and combustion cannot be effectively improved by the bleed block and the inlet is still in the limit working state. Under flight Ma∞ = 5.5 condition, with the reduction of fuel equivalence ratio, RBCC engine successfully transited from ramjet mode to scramjet mode, and the flow separation zone near the bleed block disappeared. At this time, the high temperature gas of the rocket has no obvious effect on the thrust gain of the engine. By adjusting fuel mass flow rate in a small range and rocket mass flow rate in a large range, stable mode transition can be achieved, with relatively small time consumption and minimum thrust fluctuation during mode transition. With the increase of flight altitude, under flight Ma∞ = 6.0, the incoming flow speed increases, but the air capture ability of the inlet decreases, which is not conducive to the mixing of fuel and air combustion. When Ma∞ = 5.5, the mode transition by reducing the fuel equivalence ratio will cause the RBCC engine to produce huge thrust fluctuations, and the range of the high-pressure zone in the combustor will be greatly reduced. In order to ensure the stable operation of the RBCC engine, the rocket needs to operate at a larger mass flow rate.

Numerical study on the dynamic process of ramjet/scramjet mode transition in the integrated RBCC full flow path through multistage fuel injection strategies

The rocket-based combined-cycle (RBCC) engine usually completes the transition process from ramjet to scramjet mode under the Ma∞ = 5.5–6.0. The dual-mode (ramjet/scramjet) transient process is an important technology for engine operation. CFD modeling based on Reynolds average Navier-Stokes was for studying the influence of fuel injection position, high-temperature rocket jet and fuel throttling time on the combustion flow structure and transient characteristics of RBCC engine during the mode transiton . The results showed that whether RBCC engine could successfully achieve stable mode transition was mainly influenced by flame stability in the combustor. When the rocket shuts down, it was more suitable to choose the fuel injection arrangement which distributed downstream of the combustor to realize relatively smooth mode transition. The analysis shows that in the mode transition process, using the synergistic combustion-supporting effect of the cavity and the strut is beneficial to improve the fuel combustion efficiency and reduce the thrust fluctuation. Although the high temperature rocket jet with low mass flow rate cannot maintain the fuel combustion in the isolator at high Mach number, it can expand the high temperature combustion area width in the central flow of RBCC engine and increase the secondary fuel combustion efficiency. Therefore, keeping the rocket running at low mass flow rate can reduce the thrust drop caused by insufficient fuel heat release. When adopting a mode transition strategy that throttles the fuel linearly over time, the engine can go through the thrust dip more smoothly, but this correspondingly increases the time required for the internal flow field to stabilize.

Study on SABRE and rocket-based combined cycle engine

Since the 20th century, countries have been conducting researches on near-spacecraft and have proposed a series of representative space vehicle concepts. This paper reviews the brief history of the development of reusable near-spacecraft and focuses on two combined cycle propulsion modes, SABRE and RBCC, with the goal of improving the performance of future near-spacecraft propulsion systems. An improving idea of combining the RBCC with the SABRE is proposed, and its possibility is discussed by analyzing the specific impulse curve. Two combining configurations of them and design principles are obtained by citing the literature and studying similar configurations. The two designs are also compared and the more optimized one is suggested, so as to offer some inspiration for future studies.

Investigation on geometric throat matching characteristics of variable geometry rocket-based combined cycle combustor

One of the most effective ways to realize multi-mode and high-performance matching operation of rocket-based combined cycle (RBCC) engine is to adopt a variable geometry combustor. In this study, direct-connect experiments and three-dimensional numerical simulations were used to investigate the matching characteristics between the isolator and the geometric throat of the variable geometry rocket-based combined cycle combustor under Ma 6 condition. The influences of equivalence ratio and geometric throat height variations on the combustor matching characteristics were obtained. The study results show that: (1) The pressure in the combustor and the outlet pressure of the isolator increase gradually with the increase of equivalence ratio, which causes the pre-combustion shock train position in the isolator to move forward gradually. Meanwhile, the combined injection of secondary fuel with the isolator and the pylon is beneficial to the improvement of combustor pressure, thus further improving the engine performance. (2) Different geometric throat heights have a substantial impact on the isolator and combustor matching when the equivalent ratio of the combustor is fixed. Under the conditions of Ma 6 operating at an altitude of 24 km inflow, the starting position of the pre-combustion shock train in the isolator is precisely at the isolator entrance, and the corresponding geometric throat height is 2.10H. The isolator and the combustor can complete matching operation at this point which the geometric throat height is the minimal throat height for Equivalent ratio = 1.0. (3) The reduction in geometric throat height can result in an increase in combustor pressure but a loss in combustor heat release and combustion efficiency of fuel combustion, which are reduced by 21.1% and 7.2% respectively. (4) The heat release and combustion efficiency in the combustor will decrease with a decrease in equivalent ratio at the minimal throat height condition where the isolator and the combustor can complete matching operation. This happens when the equivalent ratio and the geometric throat height change synergistically.

Experimental study of rocket plume expansion in the rocket-based combined-cycle engine under the ejector mode

In this paper, we utilized high-speed schlieren technology to investigate the impact of expansion degrees on the evolution of subsonic-supersonic mixing layers. Specifically, we explored characteristics of supersonic and subsonic flows with Mach numbers of 1.27 and 0.75. Our findings revealed that over-expansion of the supersonic flow facilitated transition, causing the generation position of Kelvin-Helmholtz vortexes to shift forward by 35.3%∼45.1%. Additionally, the growth rate of visual thickness was observed to increase to 2.7∼3.5 times in the incipient stage of the mixing layer under these conditions. We employed root-mean-square analysis to observe the rapid increase of turbulence intensity and the expansion of the strong turbulent region in the incipient stage (within x/H=0∼1.5), from under-expansion to over-expansion conditions. Our results indicate that the turbulent region and intensity play a crucial role in the evolution of mixing layers. Spatial correlation analysis revealed larger turbulent structures under over-expansion effects, with structures at x/H=1.0 being over 8.2% larger than those under weak under-expansion conditions. This turbulent mechanism explains why the over-expansion of rocket plumes can enhance mixing layers. Therefore, designing the rocket in the over-expansion state is an effective strategy for rocket-based combined-cycle engines under the ejector mode. Moreover, we noted that mixing layers and Kelvin-Helmholtz vortexes deflected to the low-pressure flow side, and pressure-unmatched conditions would cause flatten turbulent structures in subsonic-supersonic mixing layers.

Effect of the heat release on ejector-to-ramjet dynamic mode transition process in the rocket-based combined cycle engine

Effect of the heat release on ejector-to-ramjet dynamic mode transition process in the rocket-based combined cycle engine

A Mixing Process Influenced by Wall Jet-Induced Shock Waves in Supersonic Flow

With the development of hypersonic air-breathing propulsion systems, such as the supersonic combustion ramjet (Scramjet) and rocket-based combined cycle (RBCC) engines, the mixing process of supersonic airstream with fuel in the engine combustor has been drawing more and more attention. Due to the compressibility effects, the mixing process in a supersonic condition is significantly inhibited. In the present paper, the novel strategy of wall-jet induced shock waves (WJISW) is put forward to realize mixing enhancement. The interaction process between WJISW and the supersonic mixing layer is researched and the enhanced-mixing mechanism is revealed, employing large eddy simulation (LES) methods. The fine vortex structures of the flow field are well captured and presented, utilizing the numerical schlieren technique. Detailed visualization results indicate that WJISW in a low frequency condition can result in the ‘region action mode’ (RAM) never reported before. The drastic dynamic behaviors including growth, deformation, and distortion in the interaction region can undoubtedly promote the mixing of upper and lower streams. The Reynolds stress distributions along the streamwise x-direction suggest that more intense fluctuations can be achieved with a low frequency WJISW. Moreover, a sharp increase in mixing layer thickness can be realized in the interaction region. The dynamic mode decomposition (DMD) analysis results show that the mixing layer evolution process is dominated by the mode induced by WJISW, which leads to the coexistence of both large- and small-scale structures in the flow field. The entrainment process corresponding to large-scale vortices and the nibbling process corresponding to small-scale vortices can obviously promote mixing enhancement. It is suggested that the present proposed strategy is a good candidate for enhanced-mixing with application to Scramjet and RBCC combustors.

Numerical study on the lobed nozzles for enhancing the mixing and combustion performance of rocket-based combined cycle engine

Numerical study on the lobed nozzles for enhancing the mixing and combustion performance of rocket-based combined cycle engine

Ejector mode performance improvement of rocket-based combined cycle engine designed for dual-ramjet mode

Rocket-based combined cycle engine integrating multi-mode operation accomplishes an excellent performance under various flight speed conditions. Generally researchers selected dual-ramjet mode as the cruise and design point with adjusting heat addition method to improve the performance under other modes. To analyze the difficulties during the multi-mode design, a unilateral expansive flowpath through central rocket ejection was investigated to analyze the effects of the mixing ratio of oxygen and fuel and the mass flow rate of the primary rocket plume. The results indicate that the mixing ratio increase can significantly strengthen the entrainment amount. There is an optimal operating rate to obtain the highest entrainment ratio due to Fabri choking limit induced by a confined path width. However, the excessively divergent afterbody based on dual-ramjet results in the under-expansion issue of the exhaust gas and negative thrust of the nozzle. The multistage ejection method can effectively overcome this shortage by compressing the exhaust gas. But with the flight speed increasing, the existence of upstream bow shock induces high inlet resistance. Adding extra duct can effectively solve the problem by decreasing inlet convergent ratio and sucking more secondary flow for combustion.