Overview of the Energy Conservation and Sustainable Transformation of Aerospace Systems with Advanced Ejector Technology

Abstract

As an energy-saving fluid machinery component, the ejector holds significant potential for promoting energy conservation and sustainable transformation in aerospace. This review synthesizes recent progress, identifies persistent challenges, and outlines future directions for ejector technology in this field, addressing a gap in existing reviews. (1) In aero-engine systems, performance faces constraints from high-speed compression effects and flow losses. These systems require optimized design across a wide range of speeds. A mixed configuration incorporating a blade mixer achieved a 5~7% thrust increase under static conditions. (2) In high-altitude test facilities, transient start-up and flow instability under off-design conditions demand more precise models and control strategies. An alternative solution using a second throat exhaust diffuser reduced the start-up time by 50~70%. (3) In rocket-based combined cycle engines, development is limited by thermal choking, mode transition, and combustion-flow coupling issues. Optimization of the rocket layout and geometric throat increased the bypass ratio in ejector mode by 35% and improved the specific impulse by 12.5%. Future efforts should focus on constructing multi-physics coupling numerical simulation models for ejectors, analyzing unsteady flow behavior and thermal effects within ejectors, and developing performance optimization strategies based on intelligent control. These approaches are expected to enhance the engineering applicability and system efficiency of ejector technology in the aerospace field, which is increasingly focused on energy conservation and sustainable transformation.

1. Introduction

Amid the accelerating transformation of the global energy landscape, carbon emissions from the aerospace field have exceeded 1 billion tons [1]. The aerospace field urgently requires a dual breakthrough: not only a significant improvement in energy efficiency but also the adoption of new energy sources, both of which should be achieved through technological innovation. Optimizing energy efficiency in aerospace is crucial for global low-carbon development, as aerospace is a typical high-energy consumption field. Ejector technology is known for its inherent advantages, including zero mechanical energy loss, a compact structure, and adaptability to high temperature and high-pressure environments. These features enable ejector technology to meet the stringent demands of the aerospace field for a high thrust-to-weight ratio and energy cycle efficiency. Therefore, ejector technology has become one of the key pathways for driving energy conservation, emission reduction, and sustainable transformation in aerospace.

A typical ejector consists of diffusers, suction chambers, nozzles, and mixing chambers, as illustrated in Figure 1. The operating principle begins when a high-temperature, high-pressure motive fluid passes through the nozzles. During this stage, both potential and thermal energy are converted into kinetic energy. The resulting high-velocity, low-pressure jet enters the suction chamber, creating a local pressure lower than that of the suction fluid, which is thereby drawn into the system. The pressure difference between the two fluids generates a strong shear effect, which causes continuous mass and energy transfer between the suction fluid and the motive fluid. Therefore, the mixing process within the mixing chamber gradually achieves uniformity. As the mixed fluid passes through the diffuser, the cross-sectional area of the flow steadily increases. This results in a gradual decrease in velocity and an increase in pressure, as a portion of the kinetic energy is transformed back into potential and thermal energy. Ultimately, the diffuser outlet produces a mixed fluid with moderate temperature and pressure. In this configuration, a high-pressure motive fluid converts its energy into kinetic energy, which subsequently entrains and pressurizes a secondary suction fluid without mechanical input. The compact, passive design presented here is important because it supports the ejector’s suitability for aerospace applications. Characteristics such as minimal mechanical loss, structural simplicity, and reliable operation in extreme environments are key reasons for its use in systems demanding a high thrust-to-weight ratio and efficient energy cycling, underpinning its energy-saving potential. Because of these intrinsic advantage, ejector technology is employed not only in aerospace but also in various other fields, such as waste heat recovery [2], nuclear power plant safety engineering [3], rail transportation [4], petroleum refining [5], refrigeration cycles [6], and desalination [7]. Furthermore, ejector technology demonstrates significant value in energy recovery, pressure control, vacuum generation, and energy saving.

In the aerospace field, ejectors operate by ejecting fluid through a nozzle outlet into a generally cylindrical cavity, as illustrated in Figure 2. This figure identifies cross-section 1 and cross-section 2 as the nozzle inlet and outlet, respectively. Similarly, cross-sections 3 and 4 are designated as the ejector inlet and outlet, while cross-section 5 represents the region between the nozzle outlet and the ejector. The primary, high-speed flow accelerates from the nozzle inlet (Section 1) to the outlet (Section 2). This expansion creates a pronounced low-pressure zone around Section 5, which is the fundamental mechanism that inducts the secondary fluid from Section 4. The subsequent interaction, mixing, and transfer of momentum within the chamber, culminating at the ejector outlet (Section 3), are primarily governed by shear forces between the co-flowing streams. This schematic underscores a critical functional duality: the device acts not only as a fluid pump but also as a thrust augmenter or acoustic suppressor. The operating principle illustrated here is foundational, as it persists across diverse aerospace implementations. Therefore, a clear understanding of this scheme is essential for the critical evaluation of ejector performance in the specific systems.

The ejector is known for its compact design, reliable operation, and low maintenance costs. Moreover, its working mechanism relies on an innovative physical structure rather than additional energy input. This design addresses the critical demand for low consumption and high efficiency in new energy technologies. In the aerospace field, ejector technology is utilized in aircraft engine systems, high altitude test facility (HATF), rocket-based combined cycle (RBCC) engines, noise reduction, and infrared radiation suppression. Research on these applications not only directly impacts energy efficiency standards in the aerospace field but also fosters technological innovation in key areas, including the transformation of industrial processes for energy conservation and the integration of renewable energy systems.

While recent reviews have systematically summarized ejector advances in multiple civil and industrial fields—such as their function as thermal compressors and vacuum maintainers in multi-effect distillation for desalination [10,11], their integration into solar-driven refrigeration cycles and multi-generation systems for improved energy sustainability [12,13], their role in enhancing the efficiency of vapor-compression heat pumps and refrigeration cycles, their applications in waste heat recovery, braking energy recovery, hydrogen circulation, and active flow control for rail transportation [4]—these works have predominantly focused on moderate operating conditions, subsystem-level performance, or specific thermodynamic cycles. However, a review that addresses the unique roles, recent progress, and critical challenges of ejector technology within the extreme and dynamic operating environment of aerospace remains absent. To address this gap, this review examines the current research status of aircraft engine systems, HATF, and RBCC engines. It also explores the potential synergies, existing challenges, and future research directions for these technologies. Specifically, the analysis emphasizes their roles in promoting energy conservation, emission reduction, and the sustainable transformation of the aerospace field. This focused approach not only enhances understanding of research hotspots, progress, and limitations related to ejector technology in the aerospace field under global energy-saving and carbon-reduction goals but also fills a significant gap in existing academic literature.

2. Performance Parameters Related to Ejector Technology in the Aerospace Field

In aerospace, engine parameters including thrust and specific impulse, not only directly influence the aircraft’s power output but also are strongly linked to efficient energy utilization. Higher thrust enables lower fuel consumption for the same task, and a better specific impulse signifies more propulsion per unit of fuel. Improvements in these parameters collectively drive the system toward lower energy consumption. The following section details the key performance parameters related to ejector technology in the aerospace field and their roles in enhancing energy utilization efficiency.

The emphasis on key performance parameters has evolved with the distinct efficiency goals of different aerospace missions. In civil aviation, the bypass ratio has become paramount, driven by the need for fuel efficiency and noise reduction. Its value has increased systematically across engine generations, culminating in ultra-high-bypass designs [14,15]. For supersonic and hypersonic aircraft, the accurate ground simulation of their flight environment became paramount. This requirement advanced the development of HATF, where parameters such as pressure coefficient and thrust augmentation were established as one of the central design and test metrics [16,17,18]. The emergence of combined-cycle engines (e.g., RBCC) further necessitated the integrated optimization of these parameters to enable efficient multi-mode operation [19,20]. This historical progression underscores that optimizing ejector performance is intrinsically linked to the dominant efficiency and mission requirements of each aerospace era. The following subsections detail these key parameters and their roles in enhancing energy utilization.

The bypass ratio β is defined as the ratio of the mass flow rate of air through the outer bypass to that through the inner bypass of a turbofan engine. Tokudome [21] performed a quasi-one-dimensional analysis that examined the internal flow field and combustion process in an experimental ejector engine. The effect of the cross-sectional area ratio (AR) of the primary rocket exit to the mixer section on the bypass ratio was determined, as presented in Equation (1) [21].

$$\beta =\sqrt{{\displaystyle \frac{{\gamma }_a{R}_R{T}_R^{*}}{{\gamma }_R{R}_a{T}_a^{*}}}{\left\{{\displaystyle \frac{({\gamma }_R+1)}{2}}\left[1+{\displaystyle \frac{M{a}_{am}^2({\gamma }_a-1)}{2}}\right]\right\}}^{{\displaystyle \frac{({\gamma }_R+1)}{({\gamma }_R-1)}}}}\times M{a}_{am}\left(1-{\displaystyle \frac{A_{Rm}}{A_m}}\right)/\left({\displaystyle \frac{\stackrel{*}{P_R}/\stackrel{*}{P_a}}{A_m/A_{tR}}}\right)$$

(1)

where A_m, A_Rm, and A_tR are the areas of the mixer section, primary rocket exit, and primary rocket throat, respectively. Ma_am represents the mass of air entrained in the mixer. Equation (2) is obtained by simplifying Equation (1) [21].

$$\beta \propto {\displaystyle \frac{1-A_{Rm}/A_m}{C}},\hspace{1em}C\equiv {\displaystyle \frac{\stackrel{*}{P_R}/\stackrel{*}{P_a}}{A_m/A_{tR}}}$$

(2)

where it could be seen that the flow path area ratio A_Rm/A_m from the exit of the main rocket to the mixer section and the total pressure ratio $\begin{array}{c}*\\ P_R\end{array}/\begin{array}{c}*\\ P_a\end{array}$ between the primary flow and the secondary flow have a significant influence on the bypass ratio. The increase of $\begin{array}{c}*\\ P_R\end{array}/\begin{array}{c}*\\ P_a\end{array}$ should lead to the decrease in the bypass ratio. When other parameters are constant, reducing the total pressure ratio $\begin{array}{c}*\\ P_R\end{array}/\begin{array}{c}*\\ P_a\end{array}$ of primary and secondary flows is beneficial to improve the bypass ratio. A higher bypass ratio indicates that more outside air is introduced to participate in the combustion or propulsion process Therefore, the consumption rate of fuel is reduced, and the energy is reasonably distributed.

$$F_{\mathrm{ejector}}={\dot{m}}_{\mathrm{mix}}^{*}{u}_1-{\dot{m}}_s{u}_0$$

(3)

where ${{\dot{m}}^{*}}_{mix}$ is the total mass flow of the mixed fluid (kg/s), ${{\dot{m}}_s}^{*}$ is the total mass flow of the suction fluid (kg/s), u₁ is the speed of the nozzle exit (m/s), and u₀ is the flight speed in freestream condition (m/s) [22].

Specific impulse I_sp is a key parameter for assessing the fuel efficiency of propulsion systems. A higher specific impulse indicates greater impulse generated per unit mass of fuel consumed. Therefore, it can reduce the amount of fuel required and the overall aircraft weight for a given mission. This contributes to reduced energy loss during flight and enhances mission economy and sustainability. The specific impulse is defined as

$$I_{\mathrm{sp}}=F_{\mathrm{ejector}}/g_0{\dot{m}}_p$$

(4)

where g₀ is the gravitational acceleration of sea level (m/s²) and ${\dot{m}}_p$ is the total mass flow of the motive fluid (kg/s) [22].

The pressure coefficient C_p quantifies the back-pressure reduction achieved by the ejector system, as defined in Equation (5). A rise in the C_p corresponds to a greater pressure rise across the ejector and a lower pressure at the primary nozzle exit. Reducing the pressure at the exit of the primary nozzle could enhance the efficiency of the thermodynamic cycle in the turbojet engine. Improved cycle efficiency indicates a more complete conversion of fuel energy into propulsive work, thereby reducing energy losses. Therefore, fuel consumption for the same thrust output is lowered.

$$C_p={\displaystyle \frac{P_{amb}-P_p}{0.5\rho u_p^2}}$$

(5)

where P_amb is the ambient pressure (Pa), P_p is the motive fluid pressure (Pa), ρ is the density of the motive fluid and the suction fluid (assuming the same density of these two fluids, kg/m³), and u_p is the velocity of the motive fluid (m/s) [17].

Thrust augmentation, Φ, represents the increase in static thrust generated by the ejector system. By improving Φ, the engine could generate more thrust without additional fuel consumption. This can shorten the aircraft’s take-off distance and improves its climb rate. Additionally, fuel consumption can be reduced by appropriately lowering thrust output in the cruise phase to achieve an energy-saving effect. Φ is defined as follows:

$$\Phi ={\displaystyle \frac{Th}{{\dot{m}}_p{u}_p}}$$

(6)

where Th is the total thrust of the system (N), ${\dot{m}}_p$ is the mass flow rate of the motive fluid (kg/s), and u_p is the velocity of the motive fluid (m/s) [17].

Entrainment ratio (ER) is the ratio of suction fluid flow rate ${\dot{m}}_s$ to the motive fluid flow rate ${\dot{m}}_p$. The ratio indicates the effectiveness of fluid entraining and mixing in the ejector system, as shown in Equation (7) [4]. When the entrainment ratio increases, it enhances the ejector system’s capacity to entrain as well as mix surrounding fluid. This improves mixing and promotes more complete fuel-air integration. Therefore, the overall system’s energy utilization could be improved.

$$ER={\displaystyle \frac{\stackrel{.}{m_s}}{\stackrel{.}{m_p}}}$$

(7)

The improvement of the energy utilization mode is promoted from different aspects by the optimization of the above performance parameters. Enhancing thrust and specific impulse results in greater propulsion per unit of fuel, and the fuel consumption is reduced for the same mission. The engine’s dependence on its internal high-temperature and high-pressure motive fluid is reduced by the enhancement of the bypass ratio, which is achieved by introducing more outside air. Through this approach, the energy distribution is optimized. The efficiency of the thermodynamic cycle is improved by reasonably regulating the pressure coefficient, so that the fuel energy is more fully converted into propulsion work. The efficiency of energy transfer and conversion processes is ensured through the optimization of thrust gain and entrainment ratio, collectively establishing a performance foundation for energy conservation, emission reduction, and the sustainable transformation of aerospace systems.

The ejector technology demonstrates significant potential for energy conservation and engineering value in multiple application scenarios within the aerospace field, which is achieved through diverse action mechanisms and system integration approaches. This efficiency is based on the collaborative optimization of the aforementioned performance parameters. Details of the specific research progress in aircraft engine systems, HATF, and RBCC engines were elaborated in the following chapters. In the aircraft engine studied in Section 3, the ejector is operated by discharging a secondary flow and combining it with the primary flow’s momentum. This action optimizes thrust performance and improves flow matching. Consequently, the propulsion system’s total energy efficiency is enhanced. In the HATF detailed in Section 4, ejectors serve as a key component of the exhaust diffusion system whose geometric optimization and flow control directly determine the startup efficiency and operational energy consumption of the ground test facility. In the RBCC engine discussed in Section 5, flow organization and combustion control during ejector mode are vital, which are fundamental to achieving efficient energy conversion and enabling smooth transitions between multiple modes. Other applications of ejectors in aerospace are summarized in Section 6. These studies not only deepen the understanding of how ejectors work, but they also promote energy saving, emission reduction, and sustainable transformation in the aerospace field.

3. Research Progress of Ejector Technology in Aircraft Engine Systems

As flight technology continues to advance, integrating the design of aircraft and engines has become a critical challenge for improving propulsion efficiency and energy economy. As a key component of the engine system, the unified design of the ejector and the aircraft tail directly enhances overall energy efficiency by optimizing thrust performance and flow matching. The interaction between the ejector exhaust gases and the outside airflow alters the pressure distribution on the rear part of the fuselage. This affects the degree of gas expansion and disrupts the airflow around the fuselage, leading to increases in flight resistance. Additionally, approximately one-third of this drag is associated with the exhaust nozzle and fuselage tail. Integrating ejectors into the engine system, particularly through optimizing thrust performance, could significantly improve propulsion efficiency. This enhancement results in higher thrust output at constant fuel consumption or reduced fuel consumption at a constant thrust requirement, which indirectly improves energy efficiency.

The ejector nozzle consists of an annular shroud mounted around the engine’s primary nozzle (Figure 3). The exhaust fluid is enabled to continuously expand into a supersonic flow by ejecting the suction flow (Figure 4), which reduces premature mixing with the main flow. This structure is not only lightweight and simple but also maintains thrust enhancement across a wide flight range, thereby indirectly reducing fuel consumption.

3.1. Performance Study and Application of the Steady-State Ejector

The performance of the steady-state ejector, a key component operating under stable flow conditions, is crucial to the propulsion efficiency of aircraft engines. These components typically enhance overall propulsion performance by ejecting a secondary flow to improve fluid mixing. Steady-state ejectors can be mainly classified into momentum type, pressure type, and mixer type. The momentum type is characterized by the expulsion of secondary flows, which is achieved through direct momentum exchange along the high-speed main stream. In contrast, the pressure type relies on the entrainment of secondary flows by a low-pressure region generated through the expansion of the main flow. The mixer type refers specifically to the active enhancement of shear and mixing between the two fluids. This enhancement is accomplished by employing a special geometric structure, such as a lobe, to optimize the momentum transfer process.

Compared to other ejector types, the mixer ejector offers the advantage of achieving superior pumping and efficient mixing within an extremely short mixing pipe. Moreover, due to its low weight and manufacturing cost, the mixer ejector is the preferred choice for aircraft applications [24]. Among various mixer ejector configurations, the lobed design enhances momentum exchange and the mixing process between fluids through its unique geometry, leading to superior thrust gain [25]. Recent studies indicate that the lobed configuration yields a thrust gain of up to 9% during static testing of a micro turbojet engine [26], accompanied by a significant improvement in secondary mass flow rate. This section will later focus on the lobed configuration because of its potential to improve mixing efficiency and enhance system thrust.

3.1.1. Mixer Ejectors Adopted in Aircraft Engines

The mixer ejector is a critical component designed to enhance thrust, suppress noise, and improve the propulsion efficiency of aircraft engines through secondary flow ejection. This performance improvement directly reduces the engine’s thrust specific fuel consumption (SFC), offering significant potential for energy conservation. In order to fully realize this potential, it is necessary to conduct in-depth research into the complex internal flow mechanisms. High-precision research methods enable the precise capture of critical flow details, including shock wave structures and boundary layer separation. Their reliability in understanding the complex flow mechanisms within ejectors has been significantly enhanced through rigorous numerical verification and uncertainty analysis. The collaborative validation using high-resolution experimental data and comprehensive multi-physics models further strengthens both the accuracy of performance predictions and the engineering applicability of these models [27,28,29].

In recent years, researchers have adopted a combination of numerical simulations, theoretical modeling, and experimental validation to comprehensively study mixer ejectors across various configurations, ranging from traditional turbofan engines to adaptive cycle engines. The research primarily focuses on the thrust gain mechanism, flow ejection effect, the influence of geometric parameters, and performance optimization under different flight conditions. Consequently, engine thrust is enhanced while the fuel consumption is decreased. This shows the potential of such devices for boosting the propulsion efficiency and overall aircraft performance.

In terms of revealing the thrust gain mechanism, Presz et al. [30] investigated a thrust gain mechanism for an alternate blade mixer ejector using a combined experimental and numerical simulation approach. The findings indicated that the thrust gain ranges from 5% to 7% under static conditions, while the theoretically predicted thrust loss is approximately 4% during cruise conditions, which implies the thrust loss in actual flight is small. This small loss results from the low-energy boundary layer flow being drawn into the ejector, thereby reducing the external drag. The study indicates that ejector performance depends on flight states and installation environments. It is necessary to comprehensively evaluate the inlet flow and airframe interference effects to maximize the net thrust benefit and improve the overall energy efficiency in practical applications.

In alignment with the results obtained by Presz et al., a significant influence of flight conditions on ejector performance was also highlighted in the study by Khalid et al. Khalid et al. [31] established a two-dimensional axisymmetric model integrating a cylindrical sleeve ejector with a subsonic turbofan engine. Their study focused on the feasibility of secondary flow ejection, thrust gain, and boundary layer suction drag reduction. The result indicated that the net thrust gain reaches 24% at sea level static conditions, and net thrust drag reduction exceeds 10% during high-altitude cruise combined with boundary layer suction. This suggests a significant reduction in fuel consumption required to maintain flight during the cruise phase, demonstrating the energy-saving potential of ejectors for high-altitude long-endurance missions. However, net thrust gain decreases to about 2.3% at high subsonic Mach numbers due to substantial secondary flow ram drag. Furthermore, the two-dimensional model is difficult to fully capture the complex three-dimensional flow, and three-dimensional high-precision computational fluid dynamics (CFD) is required for further research in the future. In conclusion, the thrust gain mechanism is significantly affected by flight conditions and installation environment; Optimizing these factors is crucial to maximizing net thrust and improving overall energy efficiency.

The flow ejection effect enhances propulsion efficiency and operational economy by optimizing air ejection and matching, which has consequently attracted significant scholarly attention. Chen et al. [23] utilized the quasi-one-dimensional flow theory numerical modeling method to study a turbofan engine equipped with inlet ejectors. By optimizing flow matching, they reduced drag loss caused by the inlet overflow, directly decreased fuel consumption, and improved the range and economy of supersonic aircraft. The result showed that when inlet overflow is employed as secondary flow, the engine’s installed thrust increases by up to 28.5%, while fuel consumption decreases by 6%. Xu et al. [32] utilized a multi-precision simulation method based on a dynamically optimized and updated surrogate model to evaluate the performance and design of an adaptive cycle engine ejector. The result showed that in the third outer bypass, the suction relationship greatly influences airflow distribution, potentially causing a thrust loss of up to 22.6% during acceleration. This research reveals a complex coupling relationship between the ejector’s geometric adjustments and the engine’s working state, providing a key basis for subsequent control design and system integration.

In terms of the influence of geometric parameters on the performance of the ejector and system, He et al. [33] adopted the method of numerical simulation combined with orthogonal experimental design. For an axisymmetric ejector integrated into the tail of an aircraft, they examined the effects of geometric parameters, including throat AR and outlet AR, on ejection performance and thrust. The results demonstrated that the throat AR is highly significant, as it enhances the ejection flow ratio, thrust coefficient, and net thrust coefficient. Increasing the outlet AR slightly reduces the thrust coefficient. However, the net thrust coefficient could be substantially raised. Through the collaborative design of a surrogate model with an optimization algorithm, the performance was improved by 16.333% and 46.674% for the maximum thrust coefficient and maximum net thrust coefficient, respectively. This research reveals that optimizing ejector geometry by balancing thrust and tail drag is a key approach to enhancing performance. In summary, the mixer ejector contributes positively to improving the engine propulsion efficiency at different flight stages through mechanisms such as boosting thrust, optimizing flow matching, and reducing drag. Despite performance degrades at high speeds, the benefits of this technology are significant under static and subsonic conditions. The findings indicate that, with further optimization, design, and integration, this technology could enhance aircraft energy efficiency and reduce overall mission fuel consumption, demonstrating its tangible role in promoting energy conservation.

3.1.2. Lobed Mixer Ejector

The lobed mixer ejector, which enhances mixing through its lobe structure, is suitable for micro-engines and contributes to reduced energy loss and improved system efficiency. Its performance is influenced by geometric parameters such as the lobe expansion angle and overall configuration; optimizing these parameters is key to improving the ER and thrust gain. Related studies primarily focus on analyzing the impact of these geometric parameters on the ER, thrust gain, and flow structure, aiming to optimize the ejector’s aerodynamic performance and integration design. Through integrated system analysis, the use of ejectors has been shown to enable energy recovery and reduce system throttling losses, thereby improving thermodynamic cycle efficiency and overall performance, while also supporting higher power demands for thermal management. Furthermore, an integrated design enhances the system’s multifunctional capability, allowing it to handle different operational modes and maintain long-term stability under complex environmental conditions. Ultimately, this approach achieves energy conservation and reduces carbon emissions, demonstrating favorable environmental and economic benefits and highlighting its potential for future applications [34,35,36].

Research on lobe expansion angles across different parameter ranges indicates the existence of an optimal design interval. Performance improvements plateau or even decline when the expansion angles increase excessively. O’Sullivan [25] applied the numerical simulation to investigate the effects of lobe expansion angles ranging from 22° to 45° on the flow structure and performance of a lobed mixer. The research investigated boundary layer growth caused by changes in the expansion angle, as well as subsequent changes in streamwise vorticity intensity and total pressure loss. The findings indicated that boundary layer blockage is significantly enhanced when the expansion angle exceeds 30°, leading to a decrease in the effective expansion angle and effective lobe height. As a result, streamwise vorticity generation no longer increases with geometric expansion angle, while total pressure loss rises substantially. Therefore, within the studied parameter range, no significant performance gains occur when the lobe expansion angle exceeds 30°. This finding provides a critical basis for optimizing the lobed mixer’s expansion angle design.

Shan et al. [37] adopted the method of experiment to conduct in-depth research, which focused on the aerodynamic performance of a lobed ejection exhaust system for micro turbojet engines. The axisymmetric cross-section is illustrated in Figure 5. Their study analyzed the effects of lobe expansion angles from 5° to 17° on the ER and thrust gain. The experimental result indicated a monotonic rise in ER with increasing expansion angle, reaching a maximum increase of 22.35%. However, thrust gain initially increased, peaking at 3.55% at 8.5°, before declining to a loss of 1.34% at 17°. This research reveals a trade-off between thrust enhancement and flow loss in a lobed ejector for micro-engine applications. This finding indicates the existence of an optimal design point that balances thrust gain against additional flow loss to maximize net benefit, a consideration particularly important for energy consumption control in micro engines.

The thrust gain is effectively enhanced by the special geometric configuration of the lobed mixer. Presz [30] explored the lobed mixer ejector with the alternating lobed mixer concept by combining experiments and numerical simulations. This research detailed the mechanism by which lobed nozzles enhance jet mixing and improve secondary flow entrainment by generating streamwise vorticity. The result indicated that the ejector during cruise can draw in the low-energy boundary layer from the airframe, reducing ram drag and resulting in a cruise thrust loss of only about 4%. This verifies the effectiveness of the design in improving the thrust. Schmidt and Hupfer [26] investigated the design optimization and performance enhancement mechanisms of the lobed mixer ejector using combined experimental and numerical simulation methods. Their study focused on micro turbojet engines with a thrust range below 1000 N. Experimental tests involved standard ejectors, serrated ejectors, and two types of lobed mixer ejectors (LMN/LMN-SC) as primary ejectors, combined with a compact ejector (AR = 1.77). The result demonstrated that the lobed mixer ejector significantly improves ejection performance: the LMN achieved a 9% thrust gain in ejection mode, with a secondary mass flow ratio reaching 1.2. Furthermore, diffuser performance relies on adequate mixing. For example, the thrust gain of the LMN-SC is increased from 6.5% to 8% after the addition of a diffuser. This research confirms that the lobed mixer ejector enhances momentum exchange between the core flow and the surrounding flow, making it one of the key technical approaches for efficient thrust enhancement in the ejection systems of micro turbojet engines. The research not only reveals the inherent trade-off between thrust enhancement and flow losses but also highlights the importance of multi-parameter collaborative optimization to achieve efficient ejection performance.

Ejection performance and thrust gain can be significantly enhanced by using a lobed mixer ejector. However, the effect strongly depends on key geometric parameters, such as the lobe expansion angle, which has an optimal design range. Overexpansion may result in flow loss and reduced thrust. This conclusion indicates that energy dissipation caused by flow separation and insufficient mixing could be controlled or even reduced, while considerable thrust gains are achievable through refined geometric optimization. Therefore, the engine’s energy efficiency at the component level could be enhanced. Furthermore, the coordination between the lobe structure and the diffuser, as well as multi-parameter collaborative optimization, remains crucial for achieving efficient thrust. Future research should focus on studying the mechanisms of three-dimensional flow and the adaptability to multiple operating conditions to promote practical applications in micro engines. Regarding the thrust enhancement and energy-saving potential of ejectors with different configurations in aircraft engines,

Table 1. Research on thrust gain and energy efficiency characteristics of the ejector configuration in the aircraft engine.

Researcher (Year)	Operating Conditions	Key Findings and Thrust Gain	Energy Efficiency and the Significance of Saving Energy
Presz et al. (2002) [30]	The ratio of the outer shell outlet area to the main nozzle outlet area is approximately 2 (low area ratio) to 2.9 (moderately high area ratio).	The static thrust gain is 5% to 7%, and the cruise thrust loss is about 4%. In practice, the net gain is significant due to the reduction in drag caused by boundary layer suction.	The installation environment and flight state have a great impact on the net thrust, and the overall energy efficiency could be improved after optimization.
Khalid et al. (2010) [31]	Mach number range: static sea level to hypersonic (e.g., 0.4, 0.6, 0.8). Secondary flow inlet to main nozzle throat area ratio: 67%.	The net thrust enhancement is 24% at sea level. The net enhancement exceeds 10% with the integration of boundary layer suction at high altitude, and the enhancement is only 2.3% at high Mach numbers.	The potential for energy conservation is significant during long-duration flights at high altitudes, while performance significantly deteriorates at high speeds.
Chen et al. (2021) [23]	Mach number range: 0.6~1.5 (subsonic and supersonic). Nozzle design based on quasi-1D flow theory with main nozzle throat Mach number = 1.	The installed thrust is enhanced by 28.5%, fuel consumption is reduced by 6%, and the flow ratio error is less than 1%.	Overflow resistance is reduced by optimizing the flow matching, which leads to decreasing fuel consumption and improving fuel economy.
Xu et al. (2022) [32]	Mach number range: 0~2.0. Key diameters: main nozzle outlet 0.642 m, shell shoulder point 0.695 m, shell outlet 0.970 m.	The subsonic thrust coefficient is 0.9895. The coefficient is decreased with the Mach number in the acceleration state. Suction effects exert an impact on airflow distribution and fan stability.	Revealing the coupling relationship between geometric adjustment and working state, providing a basis for control rate design, and avoiding energy waste.
He et al. (2024) [33]	The flight Mach number is 1.05.	The throat AR has the greatest impact on the thrust coefficient. The maximum net thrust coefficient is increased by 46.674% after optimization.	The net thrust could be significantly enhanced by the collaborative optimization of geometric parameters, and the improvement of energy efficiency is contributed indirectly.
Shan et al. (2011) [37]	The variations in the beam expansion angle (5°, 8.5°, 13°, 17°) and the intake distance (10 mm, 15 mm, 20 mm).	The peak thrust gain is 3.55% when the expansion is 8.5°, the loss is 1.34% at 17°, and the thrust gain is 2.99% in the 5° configuration experiment.	There exists an optimal expansion angle that could balance the thrust gain and flow loss, which is particularly important for the energy consumption control of microengines.
Schmidt & Hupfer (2021) [26]	Main nozzle configurations: standard ring, sawtooth, wave-plate mixer variants. Combined with a ramjet nozzle area ratio of 1.77.	In ejection mode, the thrust enhancement reaches 9%, with a secondary mass flow ratio of 1.2. The enhancement increased to 8% after installing the diffuser.	Enhancing momentum exchange is the key path to achieving efficient thrust enhancement in micro engines, and it also improves propulsion efficiency.

summarizes the operating conditions, key findings, and their contributions to energy efficiency from various research studies.

3.2. Research on the Performance of an Unsteady (Pulse) Ejector

Unsteady ejectors operate using pulses. The type driven by a pulse detonation engine (PDE) relies on the high-energy jets produced by the detonation waves. However, the pulse jet-driven type utilizes vortex ring dynamics. The formation mechanism of this vortex ring structure resembles the phenomenon observed in self-excited oscillating jets. Specifically, an unstable shear layer is generated within the cavity by the high-speed primary jet, leading to the production of a series of vortex rings around this shear layer [38]. Although both involve unsteady flow, they differ in their methods of energy utilization. In aircraft engine systems, the PDE-driven ejector utilizes the high-energy pulse jet generated by the detonation wave, which can significantly improve thrust. However, controlling this jet is complex. Meanwhile, the pulse jet-driven ejector depends on vortex ring dynamics to eject secondary flow, making it more suitable for applications requiring rapid thrust adjustment.

3.2.1. Pulse Detonation Engine-Driven Ejector

The PDE-driven ejector is a device that utilizes a high-energy periodic jet generated by the detonation waves to eject secondary flows, thereby enhancing thrust. This ejector can be integrated into the propulsion system as a thrust augmentation device, improving thrust efficiency through the periodic jets induced by detonation waves. Alternatively, the ejector can be utilized to optimize energy utilization under specific flight conditions. However, control challenges arise due to its unsteady operating characteristics. Nevertheless, unique methods exist to utilize pulse energy and improve thrust generation efficiency, which is useful for enhancing energy utilization within specific propulsion cycles. The performance of the unsteady ejector is influenced by the matching relationship between its geometric parameters and the engine working state. It is also affected by the flow structure and the propagation characteristics of the detonation wave. Thrust gain and energy efficiency can be improved by optimizing these factors.

In research investigating the influence of geometric parameters (such as the length-to-diameter ratio, diameter, length, inlet shape, cone angle, and type of ejector) on thrust augmentation, Paxson et al. [39] adopted an experimental approach. Their study focused primarily on cylindrical ejectors driven by a pulse jet engine. They examined how geometric parameters, including ejector length and diameter, affected thrust gain as well as entrainment ratio. The experimental device is illustrated in Figure 6. The results indicated an optimal ejector diameter and length yielding a thrust gain of 1.8. The thrust gain was non-monotonically related to the mass ejection ratio, with the thrust not reaching its maximum even at the peak ejection ratio of 30 times the driver flow mass. This research indicates that optimizing the performance of unsteady ejectors depends greatly on jet dynamic characteristics. However, caution is advised when extrapolating these conclusions to PDE systems due to limitations related to the type of driving source. Complementing this, Allgood et al. [40] investigated the influence of the length-to-diameter ratio and internal expansion geometry on thrust gain through experimental methods for an axisymmetric ejector in a multi-cycle PDE. The consequence showed that the performance of divergent ejectors is markedly superior to that of cylindrical types, although the divergent ejector provides the additional thrust-bearing surface area. The divergent ejector achieved a maximum thrust gain of 65%, whereas the cylindrical ejector’s maximum gain was only 28%. As the length-to-diameter ratio rises, the cylindrical ejector’s performance improves. For the divergent ejector, however, the growth trend tends to be flat when the ratio exceeds 5. Along with the work of Paxson, the present study emphasizes that basic geometric parameters are of critical importance. However, the optimal design space is constrained by both the specific driving mechanism and the selected ejector type, for instance, cylindrical versus divergent.

Beyond length and diameter, equal emphasis is given to the design of the ejector inlet and its expansion profile. Shehadeh et al. [41] analyzed the influence of length, overlap distance, and inlet shape (straight edges and rounded corners) on thrust enhancement for a constant-diameter ejector driven by PDE through experimental methods. The consequence demonstrated that the maximum thrust gain for a 1.83 m straight inlet ejector is 24%, while performance improved significantly to 40% with a 0.914 m rounded inlet design under a positive overlap arrangement. The experiment was primarily conducted under static conditions, excluding the effects of actual incoming flow. The inquiry into expansion geometry was extended through the application of numerical methods. Yungster [42] applied numerical simulation to investigate the performance of a combined system comprising an oxyhydrogen PDE and an ejector. The findings demonstrated that a configuration with a convergent-divergent ejector could utilize the high-energy shock wave generated by the detonation wave to achieve thrust enhancement, with a maximum impulse gain ratio of 1.9. This ratio is significantly higher than that of a configuration equipped solely with a divergent ejector. The performance improvement is mainly attributed to the ejection of secondary airflow and the suppression of the low-pressure state within the tube by the ejector. Furthermore, the axial position of the ejector significantly impacts performance. However, real viscous effects and a wide range of geometric parameter variations were not fully explored. Therefore, further research on structural optimization and unsteady fluid–structure coupling is required for practical engineering applications. The applicability of these conclusions to real flight environments requires verification, and further study is needed to assess adaptability under dynamic incoming flow conditions.

The optimization of ejector geometry involves not only the enhancement of macroscopic thrust but also the management of internal flow dynamics and chemical reactions. Debnath [43] employed numerical simulation to investigate detonation wave propagation characteristics in hydrogen-fueled pulse detonation burners equipped with sheltered ejectors. The findings reveal that an ejector length-to-diameter ratio of 2.39 most effectively enhances vortex formation, enabling detonation wave propagation speeds up to 2310 m/s. Furthermore, while the stoichiometric gas mixture produces a maximum thrust of approximately 37.82 N, it also generates the highest NOx emissions, demonstrating the essential trade-off between thrust maximization and pollution control. In addition, Debnath [44] conducted numerical simulations on modified shrouded ejectors in pulse detonation combustors, with different cone angles (−4°, 0°, +4°) for the ejectors. This research examined how these ejectors influence detonation wave propagation and thrust in hydrogen-air and kerosene-air mixtures. The consequence demonstrated that the +4° ejector produces the strongest starting vortex ring, leading to a fully developed detonation wave within 0.032 s. In the hydrogen-air mixture, the detonation wave velocity reaches 2550 m/s, and the thrust increment is 38 N. Both of these values are better than those for kerosene oil. This study highlights the advantages of hydrogen in detonation performance. However, simplifications were introduced in the single-step chemical reaction model and the large eddy simulation method employed in this study. These simplifications affect the description of actual multi-step reactions and turbulation-chemical interactions. Collectively, Yungster and Debnath have illustrated that advanced ejector designs, such as those featuring convergent-divergent shapes and optimized cone angles, can significantly improve wave propagation and thrust. However, these studies also highlight the complex trade-offs and model limitations that must be addressed for practical application.

Research on factors such as vortex ring formation, inter-stage flow, and energy transfer influencing flow structures has been extensive. Peng et al. [45] focused on the ejector system combining an air-breathing PDE with a convergent nozzle. They adopted experimental methods to investigate the effects of single-stage and multi-stage ejector structures. The findings showed that the performance of multi-stage ejectors is inferior to that of the single-stage ejectors. This performance difference is significantly influenced by the overlap ratio between stages and the flow area. The anticipated benefits of the multi-stage structure were not realized, indicating that the balance between flow loss and energy transfer efficiency requires further optimization. Debnath et al. [46] applied the numerical simulation method to investigate the detonation wave propagation characteristics and thrust enhancement mechanism in a hydrogen and oxygen mixture within a pulse detonation combustor with an ejector. The result indicated that the vortex ring structure formed at the ejector inlet plays a key role in secondary flow mixing and thrust enhancement, achieving a maximum thrust of 36.82 N. This study demonstrated that ejectors not only accelerate detonation wave development but also optimize the flow field to enhance performance. However, the model used assumed ideal chemical equilibrium, which limited its ability to accurately describe non-equilibrium reaction kinetics and three-dimensional turbulent transient evolution. Additionally, systematic quantification of geometric parameter optimization was lacking. Therefore, research on vortex ring formation and inter-stage flow effects indicates that single-stage ejectors outperform multi-stage ejectors. The vortex ring structure enhances secondary flow mixing and thrust, but flow losses and energy transfer efficiency still require improvement. Moreover, the universality of these conclusions is limited by the simplified numerical models employed.

The performance of PDE ejectors significantly depends on the matching of geometric configuration with operating parameters, and optimal design can yield substantial thrust gains. Although current research primarily focuses on static conditions, the demonstrated thrust enhancement suggests that more impulse can be generated per unit of fuel consumed. In essence, this indicates an improvement in energy conversion efficiency, offering a promising direction for future high-performance and low-consumption propulsion systems. Performance can be effectively enhanced through the use of expansion structures, rounded inlets, and optimized length-to-diameter ratios. However, the multi-stage layout has not shown clear advantages in existing studies. Most research to date relies on static conditions and simplified models. The effects of three-dimensional transient flows, real chemical reactions, and incoming flight streams remain insufficiently explored. Further research should be focused on unsteady flow mechanisms to support reliable engineering design and applications.

3.2.2. Ejector Driven by Pulsed Jet

For the ejector driven by pulsed jets, vortex ring structures generated by periodic pulses are utilized to entrain secondary flow and augment thrust. Through this mechanism, mixing and entrainment are enhanced, while the potential for additional thrust is offered alongside optimized energy utilization. The pulsed-jet-driven ejector can be configured by adjusting its shape and pulsed jet frequency to produce greater thrust with improved fuel efficiency in aircraft engines, making it suitable for aircraft that require frequent thrust adjustments. The performance of the ejector powered by pulsed jets is collectively determined by its geometric parameters and pulsed dynamic characteristics.

On the one hand, existing studies have systematically explored the influence of ejector geometric parameters—including diameter, length, length-to-diameter ratio, and inlet shape—on thrust augmentation and flow field structures. Wilson et al. [47] carried out experimental investigations on an unsteady ejector powered by a resonance tube, focusing on how diameter, length, and nose radius influence thrust augmentation. Their results revealed an optimal diameter matching the vortex ring, achieving a maximum thrust augmentation of 1.32. Furthermore, empirical correlations developed for steady ejectors are found to remain valid under unsteady operating conditions. This study demonstrated the considerable impact of geometric parameters and vortex rings on performance. However, the detailed physical mechanisms of unsteady injection require further explanation through CFD. Additionally, the experimental results are constrained by the pulse characteristics of the resonance tube. Wilson [48] employed a combined theoretical modeling and experimental validation approach to investigate thrust augmentation characteristics in pulsed ejectors, focusing on simplified models based on vortex ring dynamics and their predictive capabilities. They indicated that maximum thrust enhancement occurred at a specific length-to-diameter ratio (L/D ≈ 6.86), with experimentally measured thrust augmentation ranging from 1.33 to 1.39. Experimental values obtained with the annular nozzle were found to be lower than the theoretical limit of 2.1. This deviation in the model is attributed to flow unsteadiness and viscous losses. Wilson [49] performed experimental studies that examined how ejector radius influences the performance of unsteady ejectors. The results showed that an optimal combination of ejector radius and length, with the optimal radius determined to be approximately 0.84 times the sum of the vortex ring radius and the core radius. Additionally, the maximum thrust gain is obtained at a pulse length-to-diameter ratio of 6.9. Collectively, these studies by Wilson et al. demonstrate that geometric optimization is crucial, with performance highly dependent on matching ejector dimensions to the vortex ring structure. Nevertheless, the unsteady mechanisms and model prediction discrepancies should be further clarified by CFD and experimental studies.

On the other hand, it has been confirmed that the dynamic characteristics of the pulses (such as pulse length, frequency, and Strouhal number St) and their interaction with the flow structure have a crucial impact on performance. Choutapalli et al. [50] investigated the performance of pulsed jet ejectors through experimental methods, as illustrated in Figure 7. The results demonstrated that optimizing pulsed dynamic conditions can achieve a maximum thrust augmentation of about 1.9. However, the one-dimensional theoretical model fails to predict this declining trend because it does not consider vortex dynamics. Pulsed jets outperform steady jets due to efficient mass and momentum exchange, though high frequencies may amplify broadband noise. Wilson [48] carried out experimental studies to examine the effects of the pulse length-to-diameter ratio and frequency on thrust enhancement in unsteady ejectors. The results demonstrated that as the pulse length-to-diameter ratio approaches the formation number (N = 6.9), the thrust enhancement coefficient reaches a peak. Furthermore, the optimal ejector radius is approximately 0.84 times the size of the vortex ring structure. This study highlights the significant role of matching pulsed dynamic characteristics with vortex ring flow structures. However, the measured thrust enhancement levels are lower than those reported in some previous literature, suggesting that performance is constrained by the characteristics of the driving source. This indicates that optimization criteria should be combined with specific pulse generation mechanisms. Choutapalli et al. [51] performed experimental work to examine the potential application of rectangular pulsed jet thrust enhancers (PETA) in aircraft that take off and land vertically. They focused on the impact of the AR and St on thrust augmentation (Φ). The results revealed that when AR is 11, and St is 0.11, thrust augmentation reaches a maximum value of 1.85, which surpasses that of steady-flow enhancers of the same size due to enhanced entrainment and mixing efficiency. However, their study at Mach 0.3 within a limited frequency range necessitates further verification at higher Mach numbers. These studies have provided important bases for revealing the physical nature of unsteady ejection and guiding performance optimization.

The pulsed jet ejector achieves maximum thrust augmentation under specific AR and pulsed length-to-diameter ratios. Its performance depends significantly on the matching between vortex ring dynamics and geometric parameters. This thrust enhancement mechanism indicates that by utilizing the inherent characteristics of unsteady flow, fluid momentum can be transferred to the aircraft with high efficiency, thereby reducing energy waste during propulsion. This approach is particularly energy-saving for flight missions requiring frequent thrust adjustments. The optimal ejector size is strongly linked to the characteristic scales of the vortex ring. However, excessive ARs or frequency deviations result in performance degradation. Existing studies are mostly based on experiments conducted under normal temperature and a single Mach number, and have not systematically examined issues such as thermal effects, three-dimensional flow structures, noise, and frequency domain adaptability in practical engineering environments. Recent research on ejector technology has shifted from fixed geometric configurations toward active controllability. By adjusting key parameters such as nozzle throat area in real time and implementing intelligent closed-loop strategies like adaptive PID control, ejectors can achieve optimized performance under varying conditions, offering more flexible and efficient fluid control solutions for thermal management and thrust modulation in aero-engine systems [52,53,54]. Future research should combine active control strategies with multi-parameter coupling analysis with real pulse sources to achieve greater depth and applicability.

4. Application and Energy-Saving Optimization of Ejector Technology in HATF

The HATF is an indispensable ground testing equipment for evaluating the performance of aerospace propulsion systems. Its operational efficiency directly affects the research and development cycle, experimental costs, and energy consumption. In the context of global energy conservation and carbon reduction efforts, improving the energy efficiency of the HATF has become one of the key technical challenges. The second throat ejector diffuser (STED) is a core component in high altitude testing. Optimizing its performance not only improves flow stability and startup reliability but also greatly reduces the operational energy consumption of the whole ground experimental system. Through pre-vacuum treatment, geometric optimization, and flow control methods, the STED can significantly shorten startup time. Moreover, it can decrease the energy consumption of auxiliary systems and enhance the overall energy efficiency of the experimental process. The STED ensures test accuracy while providing key technical support for the energy-efficient operation of large ground experimental facilities.

Before constructing a flight model and launching a satellite, it is necessary to ensure the normal operation of the propulsion subsystem. Therefore, performance tests of the upper stage engine within the propulsion subsystem should be conducted on the ground. The utilization of HATF aims to accurately evaluate the performance characteristics of the upper stage engine. Figure 8 presents a schematic diagram of the flow field for the STED-type high altitude simulator. In this device, the dynamic pressure of the combustion gas undergoes a series of complex expansion processes, forming shock waves that facilitate pressure recovery. These processes proceed along the diffuser, creating a vacuum around the nozzles within the vacuum chamber. The STED is one of the most commonly used devices for high altitude testing, and its components are shown in Figure 9.

4.1. Studies on the Flow Characteristics and Stability of the Ejector in the HATF

4.1.1. Flow Mode Conversion and Separation Mechanism of the Ejector

Previous research has primarily examined the flow characteristics of the second throat exhaust diffuser (STED) in bell ejectors and thrust-optimized parabolic (TOP) ejectors under varying back pressure conditions, using experimental and numerical simulation approaches. Particular emphasis has been placed on their influence on flow stability and the startup process. This influence arises from the conversion mechanisms between free shock wave separation and restricted shock wave separation, as well as the hysteresis effect and the pre-evacuation technology. These complex flow phenomena and their transformation mechanisms pose significant challenges to theoretical modeling and numerical simulation. Traditional isotropic turbulent models often fail to accurately capture key transition points and hysteresis in flow simulations. Current research has therefore adopted non-steady methods and anisotropic models to better resolve time-dependent dynamics and complex flow features. Supported by hybrid-scale simulations and data-driven techniques, this comprehensive approach significantly advances the understanding and prediction of flow transition mechanisms [56,57,58,59].

In 2019, Fouladi et al. [60] employed a combined experimental and numerical simulation approach to investigate the high-altitude simulation performance of the STED in the bell ejector. The results showed that, for the bell ejector, the non-dimensional hysteresis range is four times that of the cylindrical ejector. Pre-pumping vacuum resulted in a 50% to 70% decrease in the start-up time of the diffuser and eliminated abnormal transition. STED technology showed potential to reduce ejector load through hysteresis, possibly decreasing start-up time, vacuum system operation, and energy use. However, experimental conditions differed from the actual thermal tests, necessitating performance verification. In 2020, Lee et al. [61] employed a high-precision transient numerical simulation method to investigate the flow separation mechanism of the TOP ejector of the KSIV-II third-stage rocket engine during high-altitude experiments. The results demonstrated that the entrainment ratio for the 15 mm gap between the ejector outlet and the diffuser inlet significantly influences the separation mode. The stable flow field during the startup stage smooths the transition, and the gas film cooling effect (an 18% reduction in wall temperature) during the shutdown stage delays flow separation. This study revealed a unique asymmetric separation evolution path in the high-altitude experimental environment, providing a basis for the design of the ejector-diffuser gap. However, more performance considerations under various actual operating conditions are necessary for practical applications. These studies highlight different aspects of ejector performance, with Fouladi et al. emphasizing startup time reduction and Lee et al. focusing on flow separation control.

Further studies refined the understanding of transient processes. In 2023, Afkhami et al. [62] conducted a study on the STED transient startup process of a parabolic nozzle for TOP under pre-vacuum conditions by combining experimental and numerical simulation methods. The results demonstrated that a high nozzle pressure slope combined with pre-vacuum could eliminate restricted shock wave separation and shorten startup time. Extending the pre-vacuum area to the diffuser end enables instantaneous startup, with minimal impact from the initial pressure. This promising optimization is nevertheless limited by nozzle dynamics and geometry, making compatible system design essential for practical application. In the same year, Fouladi et al. [55] employed a combined experimental and numerical simulation approach to investigate the startup transient process of the STED for the TOP ejector. The results indicated that increasing the ejector pressure to 34 bar could reduce startup time by about 23% and suppress pressure fluctuations. However, the non-starting condition exhibits periodic fluctuations due to dynamic changes in the recirculation vortex. These oscillations are a characteristic of the initial startup phase in vacuum ejectors, arising from asymmetric entrainment and unsteady recirculation bubbles induced by the primary jet deflection. The suppression of these fluctuations at a critical primary pressure is attributed to the onset of secondary flow choking, which decouples the vacuum chamber from downstream disturbances [63]. This study revealed the complex flow mechanism of the TOP ejector impeller during diffuser startup but was limited to a cold air medium and a specific configuration. Moreover, it did not consider the thermal coupling effect present in actual thermal experiments. Both studies in 2023 explored STED applications for TOP, yet they addressed different optimization strategies and faced distinct limitations.

The conversion of the ejector’s flow mode is influenced by factors such as back pressure, pressure slope, and geometric gap. Pre-pumping vacuum and a high-pressure slope can effectively shorten the start-up time of the diffuser and suppress flow instability. This not only enhances the reliability and repeatability of the experimental process but also reduces unnecessary flow mode transitions, thereby lowering the additional energy required to maintain stable flow. As a result, significant energy-saving benefits can be achieved in high-frequency, long-duration ground experiments. However, current research primarily focuses on cold flow conditions and simplified geometric configurations, failing to fully address thermodynamic effects, multiphase flow, and three-dimensional asymmetric flow present in real thermal tests. Therefore, it is necessary to combine full-scale thermal experiments for verification to improve the predictive ability of the ejector’s separation mechanism and enhance the reliability of future engineering applications.

4.1.2. Startup Behavior of the Diffuser and Unsteady Flow

Unsteady phenomena during start-up, such as shock train oscillations and pressure fluctuations, not only threaten system stability but can also cause frequent cycling of the vacuum system or extended operation under high pressure. Consequently, this leads to additional energy waste. The startup behavior of the diffuser and the unsteady flow are influenced by shock train oscillations, pressure fluctuations, and geometric parameters. Through optimizing these factors, flow stability can be enhanced and operating energy consumption reduced.

In research on the interaction between shock waves and boundary layers, as well as the transmission characteristics of pressure fluctuations, this dynamic interaction phenomenon impacts the initial flow stability and operational reliability of diffusers in HATF. Otterstatter [64] employed a combined method of numerical simulation and theoretical analysis to investigate supersonic flow and shock wave phenomena in the ejector-diffuser system used in HATF. The study focused on the pressure recovery process within the supersonic diffuser, which involves the interaction between the flow and the diffuser wall surface, as well as the possible formation of oblique shock waves. The results indicated that the designed supersonic diffuser could achieve sufficient static pressure recovery through its convergent surface, wall friction, and a weak oblique shock wave. This design allows the outlet pressure to reach ambient pressure while ensuring that the entire diffuser outlet flow remains supersonic. Yeom et al. [65] adopted an unsteady numerical simulation approach to examine the flow dynamics of the supersonic diffuser under conditions close to the minimum startup pressure. The results showed that as the pressure approaches the minimum startup pressure, the shock wave train within the diffuser displays periodic oscillations, leading to substantial fluctuations in the vacuum chamber pressure. Vacuum chamber pressure stabilizes slightly above startup values. This study revealed that unsteady flow under critical conditions limits steady-state analysis accuracy for the diffuser. Numerical simulations therefore supplement missing experimental details, aiding in energy loss prevention. Sung et al. [66] utilized a combined numerical simulation and experiments approach to investigate the startup characteristics and flow field evolution mechanisms of the rocket exhaust diffuser in high altitude simulations. The results indicated that as the combustion chamber pressure approaches the minimum startup value, periodic pressure oscillations appear in the vacuum chamber. Furthermore, when the pressure rises to 44 bar, the fluctuation is efficiently suppressed. This indicates that appropriately increasing the working pressure can suppress oscillations and stabilize the flow field, thereby reducing the auxiliary energy consumed to compensate for pressure fluctuations. Higher pressure raises energy consumption, necessitating balance. Near the critical starting pressure, shock wave oscillations emerge in the supersonic diffuser. Design optimization with suitably increased working pressure stabilizes flow, reducing energy loss.

Regarding the regulatory effect of geometric optimization on flow stability, the design of geometric parameters impacts the flow behavior of diffusers as well as the total performance of high-altitude simulation tests. Kumaran et al. [67] investigated the performance of the second throat diffuser system in HATF. The results indicated that the diameter of the second throat is the most sensitive parameter affecting the starting stagnation pressure. An optimal throat diameter ratio (approximately 4.11 times the nozzle throat diameter) that results in the lowest starting required stagnation pressure. Radial and axial clearances do not alter starting pressure but cause significant backflow during unstarted states in large AR cryogenic engines. The resulting slow pressure buildup shortens preparatory time, while optimization reduces no-load energy consumption. Kumar et al. [63] adopted the CFD method to conduct design and optimization studies for large AR ejectors and the second throat diffuser system. The findings indicated that through optimizing structural parameters, including the length-to-diameter ratio of the diffuser and the diameter of the throat, a configuration was achieved. This configuration enables stable flow and pressure recovery when the back pressure is below 200 mbar. This study confirmed that the optimized design’s performance is highly dependent on precise outlet pressure control, necessitating the incorporation of a corresponding pressure regulation mechanism to maintain operational stability. Moreover, the applicability of the turbulence model used under extremely high Mach numbers and complex shock wave interactions requires further verification. These results indicate that the energy cost associated with back pressure regulation must be comprehensively considered during the design and control processes. Bharate et al. [68] employed numerical simulation methods to investigate the transient startup process of the second throat vacuum ejector in the HATF. The results showed that increasing the convergent angle of the diffuser (from 10° to 20°) leads to a 30% decrease in the vacuum contribution rate in the fourth stage and a 15% increase in startup mode pressure. Additionally, moving the nozzle position down by 25 mm completely eliminates the fourth stage. This study explained vacuum generation regulation via shock wave migration and under-expansion transition, providing a basis for ejector design. While current modeling retains turbulence limitations, optimizing throat ratio and convergent angle improves diffuser startup and lowers energy consumption. Nevertheless, backpressure control and model applicability require evaluation.

The startup process of the diffuser exhibits significant unsteady characteristics. Shock wave oscillations and pressure fluctuations in the vacuum chamber are particularly pronounced at the critical state. Geometric parameters, especially the throat diameter and AR, have a significant impact on the startup performance. An optimal throat diameter ratio exists that can minimize startup pressure requirements. Optimizing the geometric design can directly lower the energy threshold required for startup and reduce the load on the compressor unit or vacuum pump. This approach is one of the key paths for achieving energy-efficient operation of ground experimental facilities. Current research primarily relies on axisymmetric models and specific turbulence models. However, they have not fully covered three-dimensional flow effects, real gas working medium, or wider adaptability to varying operating conditions. Therefore, it is necessary to develop high-precision multi-physics coupling methods and combine them with experimental verification to enhance the predictive ability for the startup and operational behavior of the diffuser in the future.

In conclusion, the characteristics of shortened startup time, improved flow stability, and optimized energy efficiency were summarized and compared in

Table 2. Comparison of geometric optimization in HATF.

Researcher (Year)	Operating Conditions	Key Findings and Initiating Behaviors	Energy Efficiency and Engineering Significance
Fouladi et al. (2019) [60]	Geometric design parameters: nozzle expansion ratio 34.707, diffuser inlet area ratio 45.593, second throat area ratio 1.7.	Pre-vacuuming shortens the start-up time of the diffuser by 50~70% and eliminates abnormal transition; the numerical simulation shows that the abnormal increase in the vacuum chamber pressure by 17.6% is due to the restricted shock wave separation.	The significant reduction in the start-up time directly lowers the energy consumption of the vacuum system, demonstrating the remarkable advantages of the pre-vacuuming technology in enhancing the experimental efficiency.
Lee et al. (2020) [61]	Geometric parameter: nozzle expansion ratio of 94.5.	During the startup stage, there is a three-stage transition. The gap between the ejector and the diffuser affects the separation mode, and an abnormal transition is observed during the shutdown stage.	The asymmetric flow path specific to high altitude experiments has been revealed, providing a basis for gap design and avoiding energy loss caused by unstable flow.
Afkhami et al. (2023) [62]	Geometric parameters: nozzle expansion ratio 34.7, diffuser inlet area ratio 45.6, second throat area ratio 1.2, diffuser outlet area ratio 4.	Combining high nozzle pressure slope with pre-vacuum could eliminate restricted shock wave separation, significantly shortening the startup time; the initial pre-vacuum pressure value has a relatively minor impact on the startup.	The optimization of the startup is jointly constrained by the dynamic pressure of the nozzle and its geometric structure, and a systematic matching is required to achieve energy-saving operation.
Fouladi et al. (2023) [55]	Geometric parameters: nozzle expansion ratio 34.7, diffuser throat area ratio 1.7.	Increasing the ejector pressure to 34 bar could shorten the startup time by approximately 23% and suppress pressure fluctuations; the non-starting condition shows periodic fluctuations.	Pressure increase could suppress oscillations and reduce the auxiliary energy consumed to compensate for the fluctuations, but the higher pressure itself brings additional energy consumption that should be considered.
Kumaran et al. (2010) [67]	Geometric and flow parameters: nozzle area ratio 55, exit Mach number 4. Geometric ranges: diffuser-to-nozzle throat diameter ratio 3.42~10.27, radial clearance 0.0174~0.087, axial clearance −0.087~0.087.	The second tracheal diameter is the most sensitive parameter. There exists an optimal tracheal diameter ratio (approximately 4.11 times the nozzle tracheal diameter) that results in the lowest startup pressure; radial and axial clearances cause unstart backflow.	The optimal tracheal diameter ratio could directly reduce the startup energy demand and decrease the load of the compressor unit or vacuum pump, which is one of the design goals for achieving energy-saving operation.
Kumar et al. (2016) [63]	Geometric parameters: diffuser-to-primary pipe outlet height ratio 2.82, diffuser length-to-diameter ratio range 2.5~12.8.	Optimizing the diffuser structure enables stable flow at back pressures lower than 200 mbar, allowing for complete expansion within the ejector.	The ability to operate stably at lower back pressures reduces reliance on the downstream vacuum system, contributing to a reduction in the overall energy consumption of the experimental system.
Jo et al. (2021) [69]	Geometric parameters: diffuser inlet length ratio 0 or 1, second throat length ratio 3~8, expansion section length range 0~644 mm.	The inlet length ratio = 1 ensures the full development of the plume; the second tracheal length ratio $L_{st}/D_{st}$ ≥ 7 is required to ensure the full development of the supersonic flow.	Define key geometric thresholds to avoid additional energy consumption caused by unstable phenomena such as flow separation and shock oscillation.

based on the relevant literature.

4.2. Research Progress on Design and Optimization Methods of the Ejector

4.2.1. Geometric Optimization of the Nozzle and Diffuser of the Ejector and Its Impact on Energy Efficiency

Geometric optimization of the throat and diffuser is a key part of the STED design in HATF. Its purpose is to establish and sustain a stable supersonic flow with minimal energy input. The key challenge lies in minimizing flow resistance, pressure loss, and the required auxiliary driving power while satisfying performance requirements. Existing studies have mainly analyzed the effects of key geometric parameters on startup characteristics, flow stability, and pressure recovery performance. These investigations have revealed the importance of optimizing the throat diameter ratio, determining the minimum inlet length, and establishing the design criteria for the diffusion section to maintain supersonic flow, suppress backflow, and ensure stable operation. The research provides a theoretical basis for the design and performance improvement of the STED.

Research on geometric parameter optimization has demonstrated its importance for system performance. In 2010, Kumaran et al. [67] conducted a study on the performance of the STED in the HATF using numerical simulation methods. The results showed that the second throat’s diameter is the most sensitive parameter affecting the starting stagnation pressure. Moreover, an optimal throat diameter ratio leads to the lowest required starting stagnation pressure. Minimizing starting pressure is essential for energy conservation. Though radial and axial clearances do not alter starting pressure, their induced backflow during unstarted states causes energy loss, making structural optimization to suppress backflow critical. In 2021, Jo et al. [69] employed numerical simulation methods to analyze the flow and heat transfer characteristics of a diffuser with a supersonic second throat, applied in a high altitude simulation. The results demonstrated that an inlet length ratio $L_d/D_d$ = 1 could ensure the full development of the plume and maintain low pressure in the vacuum chamber, while a ratio $L_d/D_d$ = 0 would cause a sudden change in the flow path and deteriorate flow field stability. The length ratio of the second throat $L_{st}/D_{st}$ should be ≥7 to ensure the full development of the supersonic flow, and the expansion section length $L_s$ ≥ 322 mm to achieve a smooth transition from supersonic to subsonic flow. This study provides a design basis for the second throat diffuser and cooling system optimization, although real gas reactions and complex turbulence interactions were not fully included. Key geometric thresholds define stable, efficient flow boundaries, and following these design principles prevents extra energy use. While both studies focus on geometric optimization for efficiency, Kumaran et al. identify a key parameter for starting pressure, whereas Jo et al. establish specific dimensional ratios for flow stability.

Further studies have explored system-level performance optimization and validation. In 2016, Kumar et al. [63] employed the CFD method to conduct design and optimization studies for large AR ejectors and the second throat diffuser system. The results demonstrated that optimizing the diffuser’s structural parameters yields a configuration that enables stable flow and pressure recovery when the back pressure is lower than 200 mbar. Lower back pressures enable stable vacuum ejector-diffuser operation, reducing energy consumption. However, optimized design performance relies on outlet pressure control and requires regulation. Turbulence model applicability under high Mach numbers and complex shock waves remains unverified. In 2023, Kim et al. [57] employed experimental verification methods to study the optimized design of the second throat supersonic diffuser for the high-altitude test environment of an upper-stage liquid rocket engine. The results showed that the inlet in the contraction section of the diffuser should maintain a uniform velocity in the x-direction, ensuring good agreement with quasi-one-dimensional predictions. Geometric parameters show little impact on thrust, but an overly short inlet disrupts flow with high y-direction velocity. While diffuser self-suction proves limited in deep throttling, a combined system can flexibly allocate pumping tasks to enhance overall energy efficiency. These works illustrate a complementary approach, progressing from numerical analysis of stability conditions to experimental insights into system integration for potential energy savings.

The geometric parameters of the diffuser have a decisive influence on its performance. An optimal throat diameter ratio exists that can minimize the startup pressure. The inlet length ratio, throat length ratio, and diffusion section length should meet specific thresholds to ensure flow uniformity and stability. Through fine geometric optimization, the energy demand for operating the HATF can be reduced at the source, representing one of the most direct and effective technical approaches to achieving energy conservation and consumption reduction. An optimized geometric design means that, under the same test conditions, the load on the driving system (such as compressors and vacuum pumps) can be reduced, and the running time shortened. Thus, the overall energy consumption during the experiment can be significantly reduced. Current research primarily relies on cold flow conditions and simplified models and has not fully covered the influence of real working fluids, chemical reactions, and three-dimensional asymmetric flows. Therefore, it is necessary to combine multi-physics field coupling analysis, high-precision turbulence models, and thermal experiments for verification to improve the performance prediction ability and engineering design reliability of the diffuser in complex conditions in the future.

4.2.2. Design of a Multi-Stage Ejector System and Vacuum Maintenance Performance

The design of the ejector system and its multi-stage structure are crucial technologies in high altitude simulation experimental facilities. These efforts focus on optimizing the configuration of multi-stage ejectors, understanding their working mechanisms, and evaluating their performance characteristics under various conditions. Current research has systematically analyzed load balancing, working medium selection, pressure oscillation mechanisms, and system stability after engine startup in two-stage ejectors. The design of the multi-stage ejector system aims to efficiently establish and maintain the required vacuum environment through inter-stage coordination and optimization of the working medium. Its performance directly affects energy consumption in high altitude experiments, especially the working medium flow and the energy consumed by the driving ejector itself.

Early studies established foundational insights into ejector system operation and optimization. In 2012, Kumaran et al. [70] conducted a study using numerical simulation methods to investigate the performance of a two-stage ejector system for high altitude tests of large AR satellite thrusters. The principle is shown in Figure 10. The results showed that the downstream ejector plays a more critical role in preserving the vacuum of the test chamber compared to the upstream ejector. The optimal operating stagnation pressure range was found to be 18 to 22 bar. Moreover, using low-molecular-weight working fluids such as hydrogen could reduce the mass flow rate by 73.5%. The two-stage ejector maintained a vacuum at high working fluid cost. Practical use therefore demands optimized downstream design, careful working fluid selection for performance-cost balance, and precise geometric control to prevent backflow or momentum loss. In 2013, Kumaran et al. [71] conducted steady-state and transient performance studies of the HATF for a large AR rocket engine with numerical simulation methods. The research demonstrated that in the engine’s no-flow state, the external ejector should rely on nitrogen flow to establish a vacuum during full-flow operation. Rocket exhaust maintains vacuum via self-ejecting, reducing or halting external ejector flow. This self-pumping lowers external energy demand by utilizing the engine’s own exhaust. The study clarified the ejector’s working mechanism and optimization potential, though the simplified spray model inadequately captured local phase change and heat exchange.

Recent research has shifted focus towards system behavior under off-design conditions. In 2025, Kim et al. [72] conducted a study with numerical simulation methods to investigate the performance of the ejector system in the HATF of rocket engines. The results indicated that the vacuum pressure of both the integrated and independent models stabilized at 0.04 bar, but the initial pressure fluctuation in the integrated model occurred 0.09 s later. After engine start-up, the ejector could increase the vacuum pressure to 0.5 bar within 0.5 s and maintain stability. The ejector system maintains operation via self-regulation below design pressure, yet initial vacuum establishment depends critically on pressure conditions. Stable startup in practice therefore requires meeting or exceeding the design injection pressure, demonstrating the system’s energy adaptability under non-design conditions. Also, in 2025, Chao et al. [73] employed a combined experimental and theoretical approach to investigate the performance prediction of multi-stage steam ejector systems under non-design operating conditions. The aim was to simulate engine performance tests at an altitude of 42 km (with a vacuum pressure of approximately 220 Pa). The results demonstrated that the established model could accurately predict system performance, with a maximum secondary flow rate prediction error of 15.78% and a maximum ejector outlet pressure prediction error of 5.25%. However, the prediction accuracy for cooling water flow was unsatisfactory. Moreover, calibration based on experimental data was required for four key parameters in the model, namely the isentropic efficiency coefficient of secondary flow, the friction loss coefficient of mixed flow, the isentropic efficiency coefficient of main steam, and the loss coefficient of main steam. These results indicated that the method was more suitable for performance evaluation of existing systems, and its applicability to general design should be expanded. Kim et al. emphasized defining an operational threshold critical for startup stability, while Chao et al. concentrated on establishing a predictive model dependent on empirical calibration.

In these studies, optimizing ejector design and selecting appropriate working fluids are crucial for maintaining a vacuum environment. By utilizing the system’s own energy and precisely controlling operating parameters, energy efficiency can be significantly improved, thereby reducing external energy consumption requirements.

The multi-stage ejector system can achieve load balancing through parameter optimization. The downstream ejectors play a crucial role in maintaining the vacuum. Using low-molecular-weight working fluids, such as hydrogen, can significantly reduce mass flow consumption. Prioritizing the optimization of key components, selecting efficient working fluids, and fully utilizing the engine’s exhaust energy can significantly reduce the demand for external energy assistance in system design. Therefore, the energy utilization economy of the entire high altitude simulation experiment system can be improved. The energy-saving effect is reflected not only in reduced electrical energy or fuel consumption but also in a significant reduction in working fluid consumption and optimization of the system’s operation mode. The system may experience periodic pressure oscillations under low-pressure conditions, but the engine can maintain stable operation through the self-ejecting effect after startup. Current research primarily relies on specific configurations and simplified models. However, there remains insufficient adaptability to multiphase flow, heat exchange, and a wider range of operating conditions. Therefore, it is necessary to develop high-precision multi-physics field coupling models combined with experimental validation to enhance both reliability and economy in future system design.

5. Research and Application of Ejector Technology in RBCC Engines

The RBCC engine is a highly promising propulsion system that integrates air-breathing ramjet engines with rocket engines through structural integration and thermodynamic cycle coupling. This hybrid engine combines the advantages of both air-breathing and traditional rocket propulsion, enabling it to launch from zero initial velocity and accelerate into orbit while featuring reusable capabilities. Therefore, it greatly lowers costs and improves efficiency for space transportation missions [74,75,76]. The rocket ejector mode forms a distinct and essential phase of the RBCC engine, applicable in subsonic and low supersonic speed ranges, for example, flight Mach numbers under about 2.0 [77]. During static or low-speed conditions, the surrounding air is drawn into the engine by the jet effect produced by the rocket. This air then undergoes after-burning to augment thrust, achieving self-starting functionality from zero velocity. The air-augmented ejector performance continuously improves due to enhanced turbulent effects as the flight Mach number increases. In practice, during the multi-mode operation typical of aerospace missions, the performance of the ejector mode significantly influences the overall load throughout the entire process. Even within the Mach number range of 0 to 2, fuel consumption increases by more than 50% [78]. Effective thrust enhancement is closely related to the rational control of fuel consumption. By optimizing energy conversion efficiency in the ejector mode, it is possible to reduce the fuel consumption required per unit of thrust while maintaining power output. This capability underscores the technology’s value in terms of energy conservation. Therefore, analyzing the ejector mode is crucial for understanding how RBCC engines can achieve efficient energy utilization across the flight envelope. The flow path principles and characteristic cross-sectional descriptions of the typical RBCC engine ejector mode are illustrated in Figure 11.

5.1. Research on the Application of Ejector Technology in RBCC Engines

5.1.1. Theoretical and Experimental Research on the Characteristics of Ejector Mode

Theoretical and experimental research on the characteristics of the ejector mode primarily focuses on a systematic exploration of the flow characteristics, combustion processes, and performance optimization of RBCC systems operating in ejector mode. Existing studies have employed a combined approach involving experimental validation, numerical simulation, and thermodynamic analysis to thoroughly investigate the influence of key factors, such as rocket operating parameters, geometric configurations, mixing and combustion mechanisms, and thermal choking effects on performance indicators, including ER, thrust augmentation, and specific impulse. These investigations provide crucial theoretical support for the design and optimization of RBCC systems. Improvement in these performance indicators directly reflects an enhancement in system energy utilization. For instance, an increase in specific impulse indicates that a unit of fuel generates more thrust, thereby reducing the total energy consumption for the mission. Similarly, an increase in the ER indicates high-efficiency utilization of ambient air for propulsion, reducing reliance on costly onboard propellants and enhancing the economic efficiency of energy utilization at its source.

Early foundational studies explored key parameters affecting RBCC ejector mode performance. In 2000, Lehman [79] conducted experimental research on the rocket ejector mode of RBCC. The results indicated that the rocket mixture ratio has a relatively small impact on the mixing length, and the flow tends to become uniform at a distance of approximately 39.2 times the nozzle height from the outlet of the rocket ejector. This study demonstrated that jetting performance could be effectively optimized by adjusting the rocket conditions and geometric configuration. However, differences in mixing efficiency under different configurations have significant implications for actual engine intake design and modal transition strategies. In 2002, Han [22] conducted numerical simulation methods to investigate the influence of the main flow molecular weight on injection performance in the rocket-jet ramjet RBCC. The ideal experimental assumption is shown in Figure 12. The results indicated that a low molecular weight main flow generates high specific impulse, a high bypass ratio, and low thrust, whereas a high molecular weight main flow produces higher thrust but lower specific impulse. This suggests that selecting an appropriate working fluid molecular weight can balance thrust and fuel utilization efficiency. The advantage of low molecular weight working fluids in terms of specific impulse means that a unit of fuel generates more propulsive work, thereby reducing overall fuel consumption. In 2007, Kanda [80] conducted experimental studies on the combustion chamber performance of the rocket-ramjet RBCC under ejector mode. The results demonstrated that the average combustion efficiency under standard conditions is 0.8. Increasing the rocket mixture ratio enhances jet suction performance, while excessive combustion chamber pressure leads to a decrease in suction capacity. Enhanced combustion efficiency improves fuel utilization, reducing incomplete combustion losses and overall consumption. The study emphasized the difficulty of maintaining effective ejection under high rocket chamber pressure and recognized coordination between rocket and ramjet modes as a key challenge. Lehman focuses on geometric mixing development, Han on fundamental fluid property trade-offs, and Kanda on combustion interactions.

Subsequent research analyzed thrust optimization and internal flow limitations. In 2015, Yang et al. [81] employed ideal thermodynamic cycle analysis to study the maximum thrust characteristics of hydrogen fuel RBCC under rocket-ejector mode. The results demonstrated that engine thrust initially increases gradually and then rapidly decreases as the injection ratio increases. There exists an optimal injection ratio at which thrust reaches its maximum value. This optimal ER maximizes thrust per fuel unit by avoiding energy waste from improper ratios, though prediction accuracy may be limited by excluded factors such as wall friction and specific heat variation. In 2017, Wang et al. [82] employed a quasi-one-dimensional numerical simulation method to investigate the influence of thermal choking in the combustion chamber on the injection process when RBCC operates in ejector mode. The results indicated that thermal choking significantly increases back pressure at the outlet of the mixing chamber, thereby inhibiting injection airflow. This flow inhibition may reduce combustion efficiency and increase fuel consumption per unit thrust. Therefore, design optimization is necessary to mitigate its impact and maintain system energy utilization efficiency. The simulation of real and complex flows remains challenging due to the simplified treatment of shock wave structures and mixing processes in this model. In 2020, Lin et al. [77] combined thermodynamic cycle analysis with three-dimensional CFD numerical simulation to investigate the influence of the main rocket jet on the thermodynamic cycle performance of RBCC under jetting mode. The results indicated that the jetting capability can be substantially boosted through elevating the primary rocket chamber pressure and the ejector’s expansion ratio. The bypass ratio and specific impulse increased by 35.5% and 12.5%, respectively. The combined effect of these two factors leads to reduced fuel consumption for the engine under the same mission. However, the simplified treatment of flow separation and combustion instability in this model still requires further verification. Yang et al. and Lin et al. propose strategies to enhance performance, whereas Wang et al. identify a mechanism that constrains it.

Recent work employs integrated approaches for performance optimization and practical evaluation. In 2024, Luo et al. [83] employed a combined numerical simulation and thermodynamic modeling approach to research on the ejector mode of RBCC. The results indicated that increasing the AR of the mixing section and reducing the rocket pressure ratio can enhance thrust gain and specific impulse. This trade-off calls for balancing thrust enhancement and energy consumption during design, identifying through optimization a performance-efficient equilibrium. However, under higher Mach numbers, the model’s prediction accuracy still needs improvement. In 2025, Nie et al. [84] conducted a study on the sea-level thrust gain of kerosene-fueled RBCC under ejector mode by combining experimental methods, theoretical modeling, and numerical simulation. The results demonstrated that utilizing a geometric throat can achieve combustion chamber blockage, resulting in a theoretical thrust gain of 25.2% and an actual thrust gain of 15.9%. Optimizing combustion organization reduces energy loss while increasing thrust efficiency. Practical engineering must also address flow resistance and thrust loss from the fuel support plate, requiring further structural optimization to approach theoretical performance. Luo et al. analyzes parametric trade-offs, while Nie et al. evaluates a specific design concept and its implementation challenges.

Existing research has revealed that the intrinsic mechanism of thrust performance under ejector mode is constrained by multiple factors, providing important theoretical support for the design and optimization of RBCC.

In summary, the performance of the ejector mode can be optimized by adjusting parameters such as rocket chamber pressure, ejector expansion ratio, and mixing section AR. However, actual performance remains constrained by factors including thermal choking, flow resistance, and combustion efficiency. Current research is predominantly based on idealized assumptions or specific operating conditions, limiting its adaptability to real three-dimensional flow phenomena, multimodal transition processes, and broad-range operational conditions. Therefore, it is necessary to further develop high-precision multi-field coupling models and combine them with systematic experimental verification to enhance the reliability of performance predictions and the effectiveness of design optimization for the ejector mode in practical engineering applications. The key findings on improving energy utilization efficiency through different methods, as reported in the relevant literature, were summarized in

Table 3. Overview of optimization research on energy utilization efficiency of RBCC engine in ejector mode.

Researcher (Year)	Operating Conditions	Key Findings and Initiating Behaviors	The Significance and Limitations of Energy Efficiency
Lehman (2000) [79]	Simulated flight at Mach 1 and 1.9. Nozzle geometry: throat height 2.54 mm, exit height 15.2 mm. Tested DAB and SMC configurations.	Achieve thermal congestion and maintain a stable mixing length	Verify the feasibility of the ejector mode; however, the difference in mixing efficiency affects the actual energy efficiency
Han (2002) [22]	Free-stream Mach number range: 0.01 to 3.0. Ejector area ratio range: 6 to 30.	Low molecular weight working fluid has a higher specific impulse, while high molecular weight provides greater thrust.	There is a trade-off between thrust and specific impulse. Low-molecular-weight working fluids could help improve fuel utilization efficiency
Kanda (2007) [80]	Entrance throat: 46.3 mm (contraction ratio 2.04). Exit throat: 74 mm (contraction ratio 1.13). Average exit cross-section Mach number: 1.0.	The average combustion efficiency is 0.8. Increasing the mixture ratio could enhance the intake performance	The increase in combustion efficiency reduces the loss of unburned fuel. However, high chamber pressure could inhibit the intake capacity
Yang et al. (2015) [81]	Analysis based on mainstream conditions at Mach numbers 2.0, 2.5, and 3.0.	There exists an optimal injection ratio that maximizes the thrust	Avoid excessive or insufficient jet ratios that lead to energy waste, and improve the efficiency of energy conversion
Wang et al. (2017) [82]	Flight Mach number: 2.0. Key geometric parameters: inlet throat area 0.14 m2, secondary flow inlet area 0.15 m2, mixing tube outlet area 0.20 m2.	The heat congestion increases the back pressure and suppresses the air ejection	The design requirements to be optimized to reduce the inhibition of thermal congestion on the intake flow and to avoid energy loss
Lin et al. (2020) [77]	Main rocket chamber pressures: 9.4 MPa and 25.6 MPa. Corresponding nozzle expansion ratios: 6 and 20.	The bypass ratio has increased by 35.5%, and the exhaust gas pressure ratio has risen by 12.5%	Significantly enhance fuel utilization efficiency and power output, reflecting the optimization of energy conversion efficiency
Luo et al. (2024) [83]	Flight Mach number range: 0 to 2.0. Geometric parameters: inlet contraction ratio 3.22, rocket nozzle area ratio 5.00, mixing section area ratio 26.87, nozzle throat area ratio 0.875.	Thrust gain and specific impulse increase, but there is a trade-off among the targets	Multi-objective optimization requires balancing thrust and energy consumption to avoid energy waste caused by the pursuit of a single performance metric
Nie et al. (2025) [84]	Sea-level static condition (Mach 0). Throat area ratio: 1.8. Mixing section length: 4 × hydraulic diameter of rocket nozzle outlet.	Theoretical thrust gain is 25.2%, but the actual value is 15.9%	The geometric throat optimization combustion system reduces energy loss, but the actual gain is limited by the flow resistance

.

5.1.2. Ejector Geometric Parameters and Multi-Objective Optimization Design

Geometric parameters and multi-objective optimization design are crucial aspects in the performance analysis of RBCC systems. These aspects primarily focus on how parameters such as the ejector’s geometric configuration, rocket layout, and area contraction influence performance indicators, including compression ratio, intake flow rate, combustion efficiency, and thrust. By appropriately adjusting key rocket parameters, implementing multi-stage injection strategies, and optimizing the flow channel configuration, the ejector’s mass flow rate and thrust performance can be enhanced [85]. Current research mostly employs methods combining numerical simulation, neural network optimization, and experimental methods to systematically analyze the coupling mechanism of parameters such as the main thrust chamber size, pipeline length-to-diameter ratio, AR, and rocket spatial layout on bypass ratio, compression ratio, and thrust efficiency. The overall energy efficiency of the system can be enhanced through geometric optimization, which can be achieved without additional energy input by optimizing the flow path and energy distribution methods. Multi-objective optimization strives to coordinate the competing relationships among different performance targets and enhance overall system performance under constraint conditions.

RBCC engines can enhance compression ratio, intake flow rate, and combustion efficiency by optimizing geometric parameters such as contraction, expansion, and throat design. In 2007, Etele [86] conducted research on the performance of the ejector in the RBCC that improves the compression ratio through area contraction. The results indicated that reducing the outlet area by 12% to 25% significantly increased the compression ratio. Specifically, the conical-cylindrical configuration increased the compression ratio by approximately 30% at a 25% contraction rate, while the air injection flow rate decreased by less than 30%. Area contraction constrains the flow path, reducing expansion loss and converting pressure into thrust. Though flow rate slightly decreases, the increased compression ratio improves energy efficiency. This confirms geometric constraint feasibility but may lower injection flow. In 2008, TomiOka [87] conducted experimental research on the performance of a rocket-ramjet RBCC model under the ejector mode. The results demonstrated that replacing the original constant-area mixing section with an expansion channel featuring a 3.1° expansion angle increased the intake flow rate by 40%. This increase implies that more air participates in combustion, which could reduce the relative fuel consumption ratio. Meanwhile, the expansion channel reduces energy loss by minimizing flow separation. This study confirms that geometric adjustments can effectively improve intake performance in ejector mode. However, it also reveals the complexity of balancing pressure rise, enhanced mixing, and maintaining inlet blockage conditions with the expansion channel. In 2020, Dong et al. [88] employed numerical simulations to examine flow and mixing characteristics within a convergent-divergent mixing duct in RBCC ejector mode. Their results showed that suitably raising the contraction ratio greatly improves the interaction between shock waves and the supersonic mixed layer, promoting rapid and thorough mixing of the primary and secondary flows. Thorough mixing enhances combustion efficiency, reducing fuel loss and consumption while maintaining thrust. This work offers a design basis for the ejector mixing section, highlighting geometric optimization to control pressure loss. In 2023, Ye et al. [89] conducted a study of the mechanism of the geometric throat in a variable-geometry RBCC combustor through numerical simulations combined with ground direct-connection experiments. The results indicated that the geometric throat precisely controls the choking position and produces thrust 8.7% higher than that of the thermal throat structure. Additionally, it exhibits a higher total pressure recovery coefficient and improved combustion efficiency. An increased total pressure recovery coefficient implies reduced airflow energy loss, while enhanced combustion efficiency directly reduces fuel waste. The combined effect enables the engine to generate higher thrust with the same fuel consumption, demonstrating that the optimized throat geometry significantly impacts performance. In conclusion, structural adjustments such as area contraction, flow-passage expansion, and geometric throat optimization can substantially improve the compression capacity, intake performance, and combustion efficiency of RBCC engines, resulting in increased thrust and reduced fuel consumption. However, it is necessary to balance key parameters, including flow rate, mixing, and total pressure loss, to achieve optimal performance.

Innovative geometric configurations, such as rocket exhaust layout, multi-stage injection strategies, and plug-type ejector, can effectively enhance the injection performance and thrust of RBCC engines. In 2005, Etele [90] utilized numerical simulation methods to study the influence of different rocket exhaust layouts on ejector performance in RBCC engines. The results demonstrated that a configuration in which 75% of the total rocket exhaust mass is injected by a ring-shaped rocket achieved a compression ratio of up to 2.47, with the mixing area covering 95% of the outlet area. This annular layout improves mixing uniformity and promotes complete combustion, thereby reducing the energy waste. A higher compression ratio indicates improved energy efficiency, as the rocket layout enhances jet mixing and compression. However, the turbulence model requires validation to accurately predict shear layer expansion. In 2022, Yan [91] conducted research on improving RBCC performance in ejector mode by combining numerical simulations with experimental verification. The results indicated that increasing the oxygen-fuel mixing ratio in the primary flow significantly enhances ejector capability. Adopting a multi-stage ejector method increased the ejector mass flow rate by 80% and thrust by 8%. Multi-stage ejectors reduce energy concentration losses by distributing energy use across stages. Active adjustment of ejection strategies and flow configuration alleviates multi-mode performance conflicts, though balancing performance requires dynamic parameter optimization across flight stages. In 2025, Song [92] conducted a detailed numerical study on the influence of the expansion ratio of plug-type ejectors on RBCC performance in ejector mode. The results indicated that introducing a plug-type ejector significantly enhances engine thrust in ejector mode, with total thrust increasing by 6.28% to 38.91%. Multi-stage ejectors reduce concentrated energy loss by distributing energy across stages. Adjusting strategies and flow configuration alleviates multi-mode conflicts, yet balancing performance requires dynamic parameter optimization across flight stages. These works reveal the competitive relationship between different geometric parameters and their influence mechanisms on ejector performance, providing an important theoretical basis and method support for multi-objective optimization design. In summary, optimizing rocket exhaust layout, adopting multi-stage ejector methods, and introducing plug-type ejector can significantly enhance the mixing ability, ejector mass flow rate, and thrust of RBCC engines. Nevertheless, challenges remain, including improving turbulence model accuracy, dynamically optimizing design parameters, and selecting appropriate expansion ratios.

Multi-objective optimization methods can effectively coordinate the geometric parameters and operating strategies of RBCC ejectors, but they should balance the competing relationships among different performance indicators. In 2005, Jahingir and Huque [93] employed a method combining numerical simulation and neural network optimization to conduct a multi-objective design optimization study of the intake duct/ejector system in RBCC. Their results demonstrated that a smaller main thruster (with an AR of 15.66) could eject more secondary flow, achieving a higher bypass ratio and improved mixing thrust efficiency. However, higher length-to-diameter ratios and AR raise bypass ratio but reduce compression and thrust efficiency. A smaller main thruster enhances secondary flow use and mixing, lowering fuel overconsumption. Optimized balance prevents energy loss from excessive flow resistance, with the study revealing coupled geometric effects and emphasizing balanced multi-objective optimization. In 2009, Pastrone et al. [94] employed a hybrid evolutionary algorithm combined with the control volume method for numerical simulation to conduct a multi-objective optimization study on the performance of the ejector-ram-jet engine in RBCC under low Mach number conditions. The results indicated that, under given constraints, this algorithm can efficiently obtain Pareto optimal solutions, with a convergence efficiency exceeding 90%. Multi-objective optimization prevents energy waste by balancing performance metrics such as thrust and fuel consumption. While the control volume method is computationally efficient, its simplified flow assumptions limit accurate prediction of non-equilibrium ejector behavior. In 2019, Klink [95] conducted a study on the full-process modal transformation and system design of the RBCC propulsion system’s ejector to rocket mode using numerical simulation and multi-objective optimization methods. The results demonstrated that the ejector should maintain a lower throttling level during the low-altitude and low-speed stage to effectively control the maximum dynamic pressure, while the ram mode requires high throttling operation to provide continuous acceleration. By optimizing the scaling factors of length, thrust, and specific impulse combined with throttling and angle of attack control, the orbital transformation task could be achieved. In summary, multi-objective optimization algorithms enable balancing key indicators such as thrust efficiency and fuel consumption in RBCC systems, significantly improving overall performance. However, it is necessary to consider the limitations of the calculation model assumptions and to account for competing goals such as compression ratio and bypass ratio in parameter design.

Geometric parameters significantly influence the performance of RBCC ejectors. Optimizing area contraction, flow channel expansion, and throat design can increase the compression ratio, intake flow rate, and combustion efficiency. Nevertheless, these modifications also impose constraints on the injection flow rate and total pressure loss. Configurations such as rocket exhaust layout and multi-stage injection strategies can further improve mixing efficiency and thrust, but challenges related to turbulence model accuracy and expansion ratio matching should be considered. Multi-objective optimization methods can coordinate competing indicators such as thrust efficiency and fuel consumption, aiming to optimize overall system performance within given constraints. However, their effectiveness is limited by model assumptions and variable ranges, and they often fail to capture complex physical phenomena such as non-equilibrium flow. Current research primarily focuses on specific operating conditions. Therefore, integrating high-fidelity simulations with experimental validation is essential to enhance the engineering applicability of design schemes in the future.

5.2. Working Characteristics of Rocket Ejector and Technology of Enhanced Mixed Flow

5.2.1. Working Characteristics and Modal Conversion Mechanism of the Rocket Ejector

The working characteristics of the RBCC engine in ejector mode and the mechanism of multi-mode transition are crucial for optimizing the overall performance of the RBCC engine. The working characteristics in ejector mode primarily include key performance parameters such as air intake volume, mixing efficiency, combustion performance, and thrust. Existing studies mainly explore the influence of different experimental conditions on these working characteristics. Regarding the multi-mode transition, a smooth transition from ejector mode to ram mode can be achieved through coordinated regulation of rocket thrust and fuel control, maintaining thrust continuity and reducing energy loss.

The working characteristics of the RBCC engine in ejector mode are influenced by factors such as thruster configuration, fuel conditions, and mixer configuration. In 2001, Cramer et al. [96] utilized experimental methods to study the effects of single-thruster and dual-thruster configurations on the RBCC engine’s performance. They focused on analyzing air intake volume, mixing efficiency, and static pressure changes. The results demonstrated that the dual-thruster configuration significantly improved ejector performance: air intake volume increased by approximately 15%, mixing distance was reduced by about 50%, and static pressure improved by 26% compared to the single-thruster configuration. This investigation, however, was limited to only one of three spacing configurations for the dual-thruster system, and the Raman spectroscopy measurements used were based on averaged signals. Consequently, the dynamic impact of unsteady effects during the mixing process could not be adequately captured. In 2014, Lin [97] investigated the RBCC engine’s working characteristics in ejector mode with a lean fuel main rocket. The results indicated that utilizing a lean fuel main rocket could significantly increase the bypass ratio. This improvement is attributed to the lower pressure in the inlet and isolation section under lean fuel conditions, which enhances the ejector’s suction capacity. However, an increase in combustion chamber pressure can inhibit the bypass ratio’s growth. Optimizing secondary fuel injection parameters improves bypass ratio without reducing chamber pressure. Although simulation trends agree with earlier experiments, this study employed a simplified reaction model, so broader validation remains necessary. In 2016, Shi et al. [98] employed a three-dimensional viscous reactive flow numerical simulation to study the RBCC intake duct’s working characteristics under ejector mode. Their results indicated that the rocket jet’s jet effect positively enhances air intake at low speeds. However, as flight speed increases, the rocket jet’s restriction effect may disrupt choked flow in the intake duct. Enhancing air intake at low speeds improves combustion efficiency, while flow matching at high speeds should avoid excessive energy loss. This model simplified the treatment of combustion organization and turbulent chemical reactions. Therefore, high-fidelity simulations and experimental validation are recommended for future work. In 2024, Chen et al. [99] employed an experimental and numerical simulation approach to investigate the impact of the blocking phenomenon on RBCC engine performance in ejector mode. They analyzed the flow structure, mixing and combustion characteristics, and thrust performance of expansion-type and uniform cross-section mixers under different blocking conditions. The results demonstrated that inlet blocking limits secondary air ejection flow rate and causes significant total pressure loss. Although the ejector volume of the uniform-section mixer is relatively low, outlet blockage enhances mixing and combustion intensity, resulting in an 18.87% increase in specific impulse compared to traditional rockets. This study indirectly assessed mixing and combustion processes through wall pressure and thrust measurements. However, in-depth analysis of key details, such as the component distribution within the flow field, remains a subject for future investigations. Therefore, the working characteristics of RBCC engines in the ejector mode are jointly determined by thruster configuration, fuel conditions, and mixer configuration, which directly affect the injection capability, mixing efficiency, and combustion performance. Key issues such as the matching of high and low-speed flows should be addressed.

The RBCC multi-modal transition research mainly focuses on the influence of thrust regulation and fuel control strategies on the smoothness of transitions and the continuity of thrust. In 2014, Huang et al. [100] conducted a comprehensive review of multi-modal transitions in RBCC engine systems, systematically analyzing three typical transition processes: ejector/ramjet, turbine/ramjet, and ramjet/ultra-combustion impulse. For the RBCC injection/impulse transition, they elaborated that by optimizing the synergy between rocket thrust and fuel injection, a smooth transition near Mach 2.6 within the optimal range could be achieved. The results indicated that reasonable strategies could effectively shorten transition time and maintain thrust continuity. This research also reveals common problems faced by each transition process, including contradictions such as intake system startup, combustion efficiency, and shock wave control. Moreover, it highlighted the need to develop seamless multi-mode transitions within a single-flow channel. In 2019, Shi et al. [101] conducted a systematic review on the injection mode of RBCC engines, focusing on the key mechanisms underlying the transition from injection mode to compression mode. The study deeply analyzed the characteristics of the “range-type” and “point-type” conversion methods, emphasizing that the core lies in coordinating embedded rocket thrust regulation, secondary fuel injection, and the formation of downstream thermal choking to achieve dynamic balance. The results demonstrated that a successful mode requires maintaining stable thrust transition while ensuring stable suction and combustion chamber pressure in the ejector. The study emphasized that future research should deepen the study of the coupling mechanisms among mixing, combustion, and thermal choking under real flight conditions to further enhance engine performance. Also, in 2019, Klink [95] conducted the modal transition process of RBCC propulsion with numerical simulation and multi-objective optimization methods. The results indicated that the timing of transitions between different propulsion modes is crucial for flight performance and dynamic pressure control. Low throttling operation of the ejector at low Mach numbers is the key factor limiting maximum dynamic pressure, while high throttling in ram mode provides the necessary thrust for subsequent acceleration. Nevertheless, the linear control strategy employed in this study did not account for aerodynamic heating effects under actual flight conditions. Therefore, the feasibility of the optimized trajectory obtained requires further verification in engineering practice. In 2020, Shi [19] conducted an experimental study investigating the conversion mechanism of a divergent kerosene fuel RBCC combustion chamber transitioning from ejector mode to ram mode under low total temperature inflow conditions. The tests verified both stepwise and progressive rocket control strategies. The results indicated that stepwise control directly shuts down the rocket during mode transition, enabling self-sustained ram combustion. However, it strongly relies on combustion enhancement facilities such as fuel supports and cavities for coordinated organization. In contrast, progressive control significantly reduces and maintains the rocket mass flow rate, utilizing the stable flame effect of the rocket plume to improve engine operational reliability. Nevertheless, it imposes lower requirements on combustion chamber layout but reduces specific impulse performance. Through coordination of rocket thrust regulation and fuel injection, as well as by adopting stepwise or progressive control strategies, smooth mode transitions are achievable. However, further research is needed regarding applicability to actual flight conditions and the thermal coupling mechanism involved.

The working characteristics of the RBCC engine are mainly influenced by thruster configuration, fuel conditions, and mixer configuration. Optimizing these factors could effectively enhance key performance parameters such as air intake, mixing efficiency, and combustion performance. Regarding mode transition, thrust regulation, and fuel control strategies are crucial for achieving a smooth transition. Both stepwise and progressive rocket control methods have their respective advantages and disadvantages. However, the thermal coupling mechanisms and dynamic matching under actual flight conditions require further research. Current work is mostly based on ground experiments and simplified models, lacking a comprehensive understanding of the real flight environment. Therefore, it is necessary to combine high-precision numerical simulations with flight tests to improve the engine’s adaptability and reliability under complex conditions in the future.

5.2.2. Mixing Enhancement and Flow Control Technologies for Rocket Ejectors

Research on ejector flow and mixing enhancement technologies primarily focuses on improving the ejector mixing efficiency of RBCC engines through geometric configuration optimization and flow control strategies. Existing studies have systematically analyzed the impacts of key parameters—such as annular aerodynamic spikes, vane nozzles, mixing section length, and secondary flow throat configurations—on flow separation suppression, mixing enhancement, and combustion performance. These investigations have revealed the operational mechanisms by which different configurations improve the ejector coefficient, reduce critical backpressure, and enhance mixing effectiveness. The core objectives of enhanced mixing and flow control are to minimize total pressure losses and thermal energy wastage caused by flow separation and non-uniform mixing, thereby improving the thermodynamic efficiency of propulsion systems. Rapid and uniform mixing is a critical prerequisite for ensuring complete chemical reactions and efficient energy release.

Earlier work examined specific aerodynamic devices and parametric geometry studies. In 2010, Boccaletto [102] conducted a numerical simulation study on a rocket ejector incorporating an annular aerodynamic spike configuration. The results demonstrated that this configuration, acting as a secondary flow ejector, could effectively suppress flow separation while consuming only approximately 5% of the mainstream mass flow rate. This enables the ejector to maintain fully attached flow conditions under low-pressure ratios. Suppressing flow separation reduces energy loss, while low secondary flow consumption avoids excessive diversion. This structure demonstrated potential for improving symmetric flow transformation under transient conditions. However, experiments used primarily cold air, thus real thermal performance and the added structural complexity require further assessment. In 2020, Dong [88] employed numerical simulation methods to systematically investigate mixing enhancement and flow control technologies within converging-diverging mixing ducts of RBCC engines operating in ejector mode. The study examined the effects of geometric parameters on flow structures and mixing characteristics. These parameters involved contraction ratio, throat position, and convergence angle. The findings indicated that suitably increasing the contraction ratio enhances interactions between shock waves and supersonic mixing layers, promoting rapid and complete mixing of the primary and secondary flows. Reducing the throat position improves mixing layer growth rates and mixing efficiency in upstream areas, but its impact weakens in downstream far-field regions. Conversely, increasing the convergence angle shows limited improvement in mixing performance while causing a sharp decline in total pressure recovery. Boccaletto’s work centered on a specific aerodynamic device, while Dong’s research investigated the effects of broader geometric parameters.

Later studies considered specific operational variables and an innovative nozzle concept. In 2021, Gu et al. [103] conducted experimental and numerical research on a rocket ejector system with a limited mixing length. The results demonstrated that decreasing the mixing section length reduces the system’s critical backpressure. When the mixing section length ratio is set to 6, the ejector coefficient achieves 3.432, and the total pressure ratio attains 36.4. Increased total pressure ratio reflects reduced aerodynamic loss, with improved ejector coefficient enhancing air induction and lowering relative fuel consumption. Despite flow non-uniformity at the mixing-section exit, this configuration meets RBCC requirements, while the 5D position marks a critical threshold for total pressure loss. In 2021, Gu et al. [104] conducted experimental and numerical investigations on a rocket ejector system equipped with a secondary flow throat, with the operational principle illustrated in Figure 13. The results revealed that rocket equivalence ratios exceeding 1.1 significantly degrade air-breathing performance, while increasing the equivalence ratio improves pressurization capability at the cost of reduced air induction capacity. Excessively high equivalence ratios result in incomplete fuel utilization and increased energy waste, whereas moderate improvements in pressurization performance reduce flow losses. The study emphasized that inlet throat design must align with the flow rate requirements of the rocket ejector system to prevent performance degradation and identified an optimal airflow range that minimizes total pressure losses. In 2022, Wang et al. [105] employed three-dimensional numerical simulation methods to investigate the enhancement effects of vane-shaped nozzles on mixing and combustion performance in RBCC engines. The results demonstrated that high-intensity streamwise vortex structures generated by vane nozzles lead to a 56% improvement in thermal mixing efficiency and an approximately 6.1% enhancement in combustion efficiency, albeit with a reduction in the total pressure recovery coefficient. The study highlights the potential of vane nozzles in RBCC engine applications while noting that their comprehensive performance under high-temperature and high-pressure conditions requires further optimization. Gu et al. analyzing the impact of discrete operational parameters, and Wang et al. assessing an active design intended to modify the internal flow structure.

Mixing enhancement technologies can significantly improve ejector performance. However, these technologies are frequently accompanied by increased total pressure losses or elevated structural complexity, necessitating a trade-off between performance gains and engineering costs. Current research predominantly relies on cold-flow conditions or simplified models, with limited validation of high-temperature and high-pressure effects under real combustion environments, as well as long-term operational reliability. Future efforts should integrate multi-physics coupled analyses combined with thermal experimental validation to further optimize the comprehensive efficiency and engineering applicability of mixing enhancement technologies.

6. Other Applications of Ejectors in the Aerospace Field

As a device that utilizes high-speed primary flow and entrains secondary flow to enhance momentum and energy exchange, the ejector demonstrates significant application potential in the aerospace field. In addition to the aforementioned application, ejectors effectively reduce the temperatures of high-temperature components and exhaust gases by injecting cold airflow, thereby significantly enhancing engine thermal protection and infrared radiation suppression. Precise thermal management not only extends component lifespan but also reduces additional energy consumption caused by performance degradation due to overheating. In noise control, ejectors exhibit a suppressive effect on broadband noise by enhancing mixing and reducing exhaust velocity, particularly in the mid-to high-frequency components. Moreover, smooth airflow organization indirectly reduces energy losses caused by turbulent disturbances. In terms of fuel economy, ejectors contribute to widespread fuel consumption reduction across different engine types by optimizing flow matching and momentum transfer, thereby improving the conversion efficiency of fuel energy into propulsive power. In recent years, multiple studies have systematically evaluated the performance and underlying mechanisms of ejectors in these areas through experimental and numerical simulation methods, providing crucial evidence for their application in engineering practice.

6.1. Application of Ejectors in Component Cooling and Infrared Radiation Suppression

For aviation turbine engines, raising the temperature at the turbine’s front end is a crucial measure to enhance the thrust-to-weight ratio. However, prolonged operation under high-temperature conditions significantly reduces the service life of hot components in these engines. Thus, to ensure the dependable operation of these components, effective cooling measures have to be applied. Among these, cooling that employs air as the cooling medium serves as the primary method for thermal protection in aero engines. Currently, the cooling air is mainly sourced from the aviation engine’s compressor. Extracting compressed air for cooling directly reduces the airflow available for combustion, leading to aerodynamic and thermal losses within the power unit. Hence, it is essential to cool heated components while minimizing the volume of air drawn from the compressor. The use of ejectors in aviation engines enhances cooling efficiency and reduces the amount of air extracted from the compressor by injecting cold air to cool hot components. Reducing air extraction allows more high-pressure air to participate in combustion, improving fuel combustion completeness, preventing energy waste caused by excessive cooling air, and simultaneously reducing compressor power loss due to additional loads. Investigating the effects of different ejector configurations on cooling efficiency, infrared radiation characteristics, and propulsion performance is a significant research direction in the study of stealth technology and thermal management for aviation engines.

In cooling efficiency research, performance can be enhanced by optimizing flow and improving mixing through different ejector configurations. In 2000, Loka [106] employed an integrated numerical simulation and experimental approach to examine the aerodynamic and cooling performance of an exhaust ejector system for tiltrotor aircraft. The results showed that the configuration without guide vanes exhibited superior cooling performance, with a 0.6% reduction in pressure loss and a significant increase in ventilation flow rate, enabling effective injection of cold air to cool engine components. This efficient cooling approach reduces component performance degradation caused by insufficient cooling. In contrast, the configuration with guide vanes experienced a 29.3% reduction in effective flow area due to flow separation, which decreases cooling efficiency. However, this study does not consider the impact of flow unsteadiness on cooling stability under different flight attitudes. In 2011, Shan [37] utilized a combined numerical simulation and experimental approach to study the cooling performance of a lobed ejector exhaust system for a micro-turbojet engine. The results showed that ejectors effectively draw in ambient air, causing the peak wall temperature of the ejector to drop from 837 K to 616 K, and they also significantly lowered the overall plume temperature. Lower temperatures reduce heat transfer losses, enhancing energy efficiency. Experimental wall temperatures remain below simulated values due to additional cooling from radiation and conduction, indicating that accurate simulation requires coupling with thermal radiation models. In 2023, Zhang et al. [107] conducted a numerical simulation study on the cooling performance of a multi-nozzle ejector for aviation engines, focusing on the impact of nozzle quantity on the injection ratio and cooling efficiency. The results revealed that the four-nozzle configuration achieves the highest injection ratio at an expansion ratio of 2 with a fixed throat area, representing a 34% increase compared to a single nozzle. This demonstrates the advantage of a multi-nozzle layout in enhancing mixing and cooling efficiency. A higher injection ratio means that a unit of cooling medium can carry away more heat, reducing the overall consumption of the cooling medium.

Ejectors suppress infrared radiation intensity (especially in specific wavebands) by reducing wall and gas temperatures. In 2015, Shi [108] employed numerical simulation methods to investigate the cooling function and infrared radiation characteristics of an axisymmetric ejector for turbofan engines. The results demonstrated that the ejected airflow reduces the length of the high-temperature gas core zone from 10 to 8.6 times the outlet diameter. The infrared radiation intensity at a 50° azimuth decreases by up to 22%, and the proportion of the low-temperature zone on the outer bypass wall increased from 5% to 50%. However, the cooling effect on high-temperature components such as the center cone is weak. This targeted cooling reduces thermal stress losses caused by local overheating while avoiding energy waste from excessive cooling. The infrared suppression effect is not significant at azimuth angles ranging from 0° to 15°, indicating that the effectiveness of the approach is highly dependent on the observation angle and wall temperature distribution. In 2021, Chen [23] utilized numerical simulation methods to study the cooling effect and infrared characteristics of a turbofan engine equipped with an inlet ejector. The results showed that the ejector utilizes excess intake flow as a secondary flow to cool high-temperature components, significantly reducing the wall temperature of the ejector and lowering infrared radiation intensity by approximately 78% under afterburner conditions. Cooling with excess flow not only avoids the requirement for additional extraction of high-pressure air but also reduces drag losses caused by spillage, thereby achieving secondary energy utilization. The study verifies accuracy through comparisons between theoretical models and CFD simulations, but it did not analyze local cooling effects on specific components such as turbine blades. In conclusion, ejectors suppress infrared radiation intensity by reducing wall and gas temperatures. In particular, cooling with excess flow can minimize energy losses. However, the suppression effectiveness is dependent on the observation angle and temperature distribution.

However, while an ejector can reduce infrared radiation, it may simultaneously lead to thrust loss and increased fuel consumption. Therefore, it is necessary to strike an optimal balance by weighing the stealth benefits against propulsion efficiency. In 2019, Chen et al. [109] investigated the infrared radiation suppression capability of a turbofan engine equipped with an ejector. Their results demonstrated that the ejector significantly reduces wall and gas temperatures by introducing a secondary flow to cool the primary flow, resulting in more than a 50% reduction in infrared radiation intensity within the 3~5 µm wavelength band. Additionally, the ejector sleeve provides physical shielding for high-temperature components. Increased injection ratios reduce thrust, especially at high speeds where stealth requires trade-offs. Optimal energy distribution is therefore necessary to balance these competing design objectives. In 2024, Chen et al. [110] employed a quasi-one-dimensional theoretical analysis to study the cooling and infrared suppression characteristics of ejectors in turbofan engines. The findings revealed that the ejector effectively reduces wall and exhaust temperatures by surrounding the high-temperature primary flow with a secondary flow, leading to an approximately 20% reduction in infrared radiation intensity. The improvement plan was found not only to enhance cooling performance but also to cause a slight increase of approximately 0.5% in fuel consumption and a reduction of about 1.5% in thrust. These outcomes indicated an energy allocation trade-off between cooling capacity and the overall economy and thrust performance of the engine. Therefore, further optimization of this balance through meticulous design is required. Therefore, the aforementioned studies have elucidated the significant effects of cold air ejection in reducing wall and gas temperatures and suppressing infrared radiation, while clarifying the trade-off relationship with thrust loss and increased fuel consumption. These findings provide critical theoretical foundations for thermal protection and stealth design in aviation engines.

These studies have demonstrated the notable efficacy of cold air ejection in reducing wall and gas temperatures, as well as suppressing infrared radiation. Meanwhile, they have clarified the trade-off between cold air ejection and its adverse effects, such as thrust loss and increased fuel consumption. The findings provide critical theoretical foundations for thermal protection and stealth design in aero-engine systems.

Ejection cooling technology is an effective approach for enhancing the lifespan and infrared stealth performance of thermal components in aero-engines. By ejecting cold air, infrared radiation can be significantly suppressed by the reduction in wall and gas temperatures. Multi-nozzle configurations, inlet ejection, and other reductions contribute to improving ER and cooling uniformity. However, this technology involves trade-offs, including thrust loss and increased fuel consumption. Moreover, the infrared suppression effect exhibits angular dependence, and the cooling efficacy for localized high-temperature regions, such as the central cone, remains limited. In practical applications, a multi-objective optimization approach is required to balance cooling efficiency, infrared suppression, and propulsion performance, thereby avoiding a decline in overall energy efficiency due to excessive cooling or structural complexity. Current research predominantly relies on numerical simulations, highlighting the need for further exploration into adaptability under dynamic operating conditions, localized thermal assessments, and multi-physics coupling modeling. Future efforts should focus on integrating cooling efficiency, infrared suppression, and propulsion performance to achieve an optimal balance between engineering feasibility and comprehensive performance through multi-objective optimization.

6.2. Application of Ejectors in Fuel Economy Optimization

Improvements in fuel economy are typically accompanied by the maintenance or enhancement of thrust, indicating that fuel consumption optimization is achieved without sacrificing power performance. Continuous optimization of fuel efficiency in aeronautical propulsion systems represents a crucial engineering challenge and a key research focus. In this context, the advantages of ejectors in integrated energy utilization are effectively leveraged. As typical propulsion devices, micro-turbojet and turbofan engines have fuel consumption performance that directly influences the economic viability and environmental adaptability of aircraft. Ejectors have demonstrated potential to improve propulsion efficiency and reduce fuel consumption by enhancing momentum transfer and airflow mixing through secondary flows. The reduction in fuel consumption directly reflects an enhancement in the conversion efficiency of fuel energy into propulsive work. Consuming less fuel for the same mission implies more efficient utilization of unit energy. In recent years, multiple studies have systematically analyzed the influence of different ejector configurations on fuel consumption, SFC, and propulsion efficiency across various engine types and operating conditions through experimental and numerical simulation approaches. These efforts provide theoretical foundations and data support for understanding the performance enhancement mechanisms of ejectors and their practical applications.

In 2021, Chen [23] employed numerical simulation methods to investigate the installation performance of a turbofan engine equipped with an inlet ejector during flight missions, focusing particularly on the ejector’s impact on the engine’s fuel consumption. The results demonstrated that, compared to a conventional mixed-flow turbofan engine, the engine equipped with an inlet ejector achieves a reduction in fuel consumption across multiple mission segments. Specifically, the TSFC is decreased by 6.014% during the supercruise state and by 2.36587% to 3.119% during acceleration and turning maneuvers. The ejector improves efficiency by utilizing intake spillage as secondary flow, reducing spillage drag, and optimizing engine-intake flow matching. This reduces fuel needed to overcome drag and minimizes energy losses from flow separation, collectively enhancing fuel economy. In 2021, Schmidt [26] utilized a combination of experimental and numerical simulation methods to study the effects of four different primary nozzles (a standard annular nozzle SN, a chevron nozzle CVN, and two lobed mixer nozzles LMN and LMN-SC) combined with an ejector on thrust and fuel consumption in a 180 N-class small turbine engine. The results showed that all configurations exhibited reductions in fuel consumption: a 5% relative reduction with the standard nozzle combined with an ejector, a 4% reduction with the CVN, and the LMN performs best, achieving a 7% reduction. The LMN-SC with a cutout design shows a 5% reduction without a diffuser, which increased to 7% when a diffuser is installed. It is also stressed that the extent of fuel consumption improvement is directly linked to the mixing capability of the primary nozzle. However, relying solely on mixing capability is insufficient to guarantee a reduction in fuel consumption; a comprehensive balance between thrust efficiency and flow losses is also necessary. In 2023, Cican et al. [9] conducted a systematic experimental study on the performance of a micro-turbojet engine equipped with an ejector, focusing on changes in fuel consumption. The study compared the fuel consumption rate (Qc) and SFC of a baseline ejector and an optimized ejector under three typical operating conditions (idle, cruise, and maximum thrust). The results indicated that after installing the ejector, fuel consumption decreases under all operating conditions. At idle, Qc decreased from 6.88 L/h to 6.59 L/h, and S decreased from 1.213 kg/Nh to 1.083 kg/Nh. During the cruise, Qc decreased from 10.19 L/h to 9.89 L/h, and S decreased from 0.612 kg/Nh to 0.584 kg/Nh. At maximum thrust, Qc decreased from 22.00 L/h to 21.82 L/h, and S decreased from 0.241 kg/Nh to 0.230 kg/Nh. These data demonstrate that the ejector enhances airflow mixing and momentum transfer, thereby improving propulsion efficiency and reducing fuel consumption per unit of thrust, and improving fuel economy. Notably, the reduction in fuel consumption is not accompanied by a decrease in thrust. Instead, thrust increases under all operating conditions, indicating that the ejector improves fuel economy without sacrificing power performance. These studies reveal the mechanisms by which ejectors improve fuel efficiency by utilizing intake spillage, enhancing airflow mixing and momentum transfer, optimizing flow matching, and reducing spillage drag. These findings provide important theoretical and practical bases for energy-saving design in propulsion systems.

Ejectors show significant potential for optimizing fuel efficiency in aero-engines. They can improve flow matching between the engine and the intake, enhance propulsion efficiency by utilizing intake spillage, and promote better airflow mixing and momentum transfer. These effects collectively contribute to reduced fuel consumption and improved fuel economy. However, the actual fuel-saving effects of ejectors depend heavily on their geometric design, flight conditions, and the characteristics of the primary nozzles. Further optimization and analysis of adaptability to varying operational conditions are necessary to ensure consistent performance benefits across a broader range of scenarios.

6.3. Application of Ejectors in Noise Control

The transportation industry stands as one of the primary sources of noise pollution, with the aviation field being a significant component within this domain. Commercial aircraft engines are a major noise source, with ejector flow constituting the primary contributor to noise from turbojet engines [111,112]. In recent years, there has been a growing focus on identifying effective solutions to curb the impact of noise pollution on the environment and human health. To mitigate aviation noise pollution, measures such as CVNs, acoustic liners, and advanced high-performance engine designs have become common approaches for reducing noise levels from turbojet engines. As both a conceptual and structural solution, ejectors can effectively reduce fuel consumption and sound pressure levels (SPL) in turbojet engines. By optimizing airflow mixing for noise reduction, ejectors avoid the increase in aerodynamic drag typically caused by additional noise-reduction devices, thereby reducing the excess fuel consumed to overcome this drag and indirectly enhancing energy utilization efficiency. This device ejects fluid through a nozzle into a primarily cylindrical cavity, utilizing the high-speed primary flow to generate induced secondary flows through momentum and energy transfer [113].

Existing research has primarily focused on a series of studies examining the noise reduction characteristics of ejectors with different configurations through experimental methods. Since the 1950s, NASA and its predecessor, the National Advisory Committee for Aeronautics, have conducted sustained research on ejector noise control [114], with a particular emphasis on the application of ejectors and their combinations with mixers for noise reduction. In 2019, Zaman et al. [115] investigated the noise reduction performance of a rectangular mixer-ejector nozzle with an 8:1 length-to-diameter ratio using experimental methods. The results indicated that the tonal noise present in the baseline configuration could be eliminated and that the pumping efficiency of the ejector could be more than tripled by installing mixing vortex generators. Additionally, a reduction in noise amplitude is observed in the low frequency range of 5 to 30 Hz. The improvement in pumping efficiency implies that the ejector could achieve the same airflow entrainment effect with less energy consumption, thereby balancing noise reduction with energy utilization efficiency. However, incomplete mixing in high-velocity regions and vortex-generated flow instability limit overall SPL reduction to about 1 dB in the best case. Though mixing enhancement eliminates tonal noise, greater broadband reduction still requires improved outlet flow uniformity and low-frequency instability suppression. In 2021, Schmidt et al. [26] employed a combined experimental and numerical simulation approach to study the thrust and thermal management performance of ejector nozzles for micro-turbojet engines. The results demonstrated that up to a 9% increase in thrust and a significantly reduced exhaust gas temperature could be achieved by employing a configuration that combines a lobed mixer primary nozzle with an ejector. Simultaneously, the ejected airflow is considered to have potential for noise reduction. The reduction in exhaust gas temperature decreases the generation of thermal noise, while the thrust enhancement implies lower energy consumption per unit of thrust, achieving a synergy between noise reduction and energy efficiency improvement. This study demonstrated that the ejector produced significant effects not only on thrust but also on thermal management. However, specific noise reduction data and spectral analysis were not provided in the investigation. Consequently, the acoustic advantages of the ejector have not been experimentally quantified, and systematic verification through specialized acoustic testing and optimization remains necessary. In 2023, Cican et al. [9] investigated the impact of ejector installation on the noise emissions of a Jet Cat P80 micro-turbojet engine with experimental methods. The results revealed that the overall SPL decreased under all operating conditions after the ejector was installed, with an average noise reduction of approximately 1 dB at maximum operating conditions and a local maximum reduction of 1.7 dB. However, the noise level at one measurement point slightly increased at idle conditions, which may be related to the easier propagation of mechanical noise caused by the shorter nozzle of the ejector. These studies highlight the synergistic potential of ejectors in eliminating tonal noise and enhancing pumping efficiency. Moreover, they clarify the complexity of their noise reduction effects, which are influenced by configuration, operating conditions, and measurement orientation. These findings provide important experimental evidence for aviation noise control.

In summary, reducing noise decreases energy loss caused by vibrations, while maintaining overall performance stability ensures the efficient conversion of energy into thrust. As a structural solution for reducing aviation engine noise, the ejector demonstrates potential by enhancing mixing through the ejection of a secondary flow by the primary flow. However, its effectiveness depends on operating conditions and measurement positions. Although its broadband noise reduction capabilities remain limited, its capabilities to eliminate tonal noise, improve pumping efficiency, and provide synergistic benefits in thrust and thermal management indicates its potential for energy conservation in multi-objective optimization. Future efforts should focus on enhancing broadband noise reduction performance and engineering applicability through flow field optimization, improved structural adaptability, and specialized acoustic testing. Ultimately, the goal is to achieve noise control without compromising the system’s overall energy efficiency.

7. Summary and Prospects

Energy efficiency and environmental sustainability have consistently been central concerns in the aerospace field. The application and promotion of ejector technology not only yield considerable economic and social benefits but also hold major importance for advancing energy conservation, emission reduction, and the sustainable transformation of the aerospace field. This review examines recent progress in ejector technology application and research within aerospace. Existing technical achievements were summarized, current research shortcomings were identified, and future research directions were defined.

In aerospace propulsion systems, energy efficiency and operational sustainability are significantly enhanced by advanced ejector technology through various mechanisms that directly contribute to reduced fuel consumption, lower operational energy demands, and improved overall system energy management. Within aircraft engines, an integrated design comprising a hybrid ejector and a lobe mixer achieved a notable thrust gain under static conditions. Moreover, when high altitude boundary layer suction technology was incorporated, the net thrust gain exceeded 10% [31]. For high altitude simulation experiment facilities, the traditional diffuser was replaced by an STED. This modification, combined with pre-vacuum treatment and geometric optimization, markedly reduced the start-up time and substantially improved the energy-saving operational capability of the ground-based experiment system. In thermal management applications, the lobe ejector decreased the wall temperature by over 26% through the ejection of cold air [37]. This approach not only enhanced the cooling efficiency of components but also improved infrared stealth performance. In RBCC engines, optimization of the ejector throat and rocket layout led to significant improvements in both the bypass ratio and specific impulse, demonstrating efficient energy conversion in multi-mode propulsion. Collectively, ejector-based technological approaches advance the transformation of aerospace systems toward high efficiency, energy savings, and low carbon emissions.

Based on existing research results, this paper provided references for engineering applications of ejector technology. In aircraft engines, ejector efficiency depends on flight state. During low-speed stages, a waveguide-type ejector increases thrust. The expansion angle of the ejector has an optimal design range. During cruise stages, integrated design utilizes drag reduction potential. For HATF, integration of a pre-vacuum system and optimization of the second throat diameter ratio reduce startup energy consumption effectively. In the ejector mode of RBCC engines, selection of the main working fluid involves a trade-off between thrust and specific impulse. A contraction-expansion-type mixed section combined with gradually changing rocket flow control improves working stability. When ejectors apply to thermal management or noise suppression, evaluation of the compromise impact on thrust is necessary. Attention should focus on noise reduction effects mainly concentrated in specific frequency bands. In summary, engineering decisions should be based on mission objectives. Through coupling simulation and experiments, targeted optimization of ejector configuration, integration method, and control strategy should be carried out.

However, several challenges persist in current research.

(1) Regarding improvements in aircraft thrust performance, the performance of the hybrid ejector is compromised under high subsonic and cruise conditions. Performance attenuation degradation occurs due to inlet stagnation resistance, secondary flow losses, and three-dimensional flow effects, which result in a net thrust gain of only about 2.3% [31]. The thrust gain of the lobe hybrid ejector exhibits a non-monotonic relationship with the expansion angle. Excessive expansion induces flow losses, with thrust loss reaching 1.34% at a 17° expansion angle [37]. Although PDE-driven ejectors demonstrate remarkable performance under static conditions, their adaptability to real flight inflow conditions has not been fully verified.

(2) In high altitude simulation experimental facilities, the starting process of the diffuser displays strong unsteady characteristics. Shock train oscillation causes pressure fluctuations within the vacuum chamber under critical conditions. Geometric parameters, such as the throat diameter, significantly influence start-up performance. However, most current models rely on axisymmetric assumptions and specific turbulence models, which lack adaptability to three-dimensional effects and real gas working fluids. Although pre-vacuum technology can shorten start-up time, its effectiveness is constrained by the dynamic pressure characteristics and geometric structure of the nozzle.

(3) In the ejector mode of RBCC engines, thermal choking increases the back pressure in the mixing chamber and inhibits the air ejection flow. Actual performance is limited by combustion efficiency and flow resistance. Multi-objective optimization reveals a trade-off relationship among thrust gain, specific impulse, and other performance indicators. Predictions at higher Mach numbers still exhibit deviations. Although the geometric throat can enhance thrust, the actual gain remains lower than theoretical values due to flow resistance and structural complexity.

(4) In other aerospace fields, ejector cooling can reduce wall and gas temperatures while suppressing infrared radiation. However, challenges such as thrust loss and increased fuel consumption remain. The rectangular mixer-ejector configuration can eliminate tonal noise, but the overall SPL reduction is limited to approximately 1 dB [9]. When the ejector is installed in micro-turbine engines, general noise reduction is achieved, though local noise may increase during idle due to enhanced mechanical noise propagation.

Future research should prioritize the following directions.

(1) Multi-physics collaborative technology should be developed by integrating CFD, structural mechanics, and thermal management models. This integration will help reveal the working mechanisms of ejectors in complex environments. Additionally, an intelligent throttling and thermal management cooperative system should be designed to achieve real-time matching between geometric parameters and working conditions.

(2) Improved transient start-up control strategies should be established to enhance the stability of the HATF under critical conditions. Additionally, a high-precision prediction model must be developed to optimize diffuser adaptability during off-design operations.

(3) Research on coordinated control of RBCC engines across multiple modes should be deepened. The transition process between ejector mode and ramjet mode requires further optimization. New fuel ejection strategies and combustion organization methods should be explored to improve performance and reliability across a wide operating range.

(4) Multi-objective optimization of the propulsive ejector should be conducted concerning noise reduction, infrared suppression, and fuel economy. A balance must be struck between stealth requirements, thrust performance, and energy consumption indicators. Lightweight, high-temperature-resistant materials and innovative structures should be developed to enhance engineering applicability.

The technical roadmap for future ejector research in the aerospace field is presented in

Table 4. Technical Roadmap for Future Ejector Research in Aerospace Field.

Challenges	Research Pathways	Key Technologies	Expected Outcomes
Performance Degradation at High Speeds: Net thrust gain of hybrid ejectors is significantly reduced under high-subsonic and cruise conditions. Excessive expansion angles in lobed mixers lead to flow losses and thrust reduction.	Multi-Physics Integration and Intelligent Control: Develop coupled simulation frameworks and adaptive control systems to enable real-time optimization across varying flight conditions.	1. Coupled CFD–Structural–Thermal Models 2. Adaptive Geometric Control Systems 3. Digital Twin Platforms for System Validation	Wide-Speed-Range Adaptive Ejectors: Enable sustained high thrust gain and low aerodynamic loss across the entire flight envelope, improving overall propulsion efficiency.
Unsteady Startup Behavior in Altitude Test Facilities: Diffuser startup is characterized by shock train oscillations and pressure fluctuations. Pre-vacuum effectiveness is limited by nozzle dynamics and geometric constraints.	Transient Modeling and Advanced Control: Establish high-fidelity transient models and robust control strategies to stabilize startup and off-design operation.	1. Real-Fluid and 3D Unsteady Flow Models 2. Turbulence–Combustion Interaction Models 3. Active Flow and Pressure Control Algorithms	Fast-Start, Low-Energy Altitude Test Systems: Achieve rapid, stable diffuser startup with minimal auxiliary energy consumption, enhancing test facility efficiency.
Mode Limitations and Performance Trade-offs in RBCC Engines: Thermal choking in ejector mode restricts secondary air entrainment. Performance is constrained by combustion efficiency, flow losses, and multi-objective trade-offs.	Multi-Mode Coordination and Enhanced Combustion: Optimize mode transition processes and develop advanced mixing and combustion strategies for wide-range operability.	1. Mode Transition Control Logic 2. Enhanced Mixing Devices (Lobed Nozzles, Geometric Throats) 3. Advanced Combustion Schemes.	Seamless Multi-Mode RBCC Propulsion: Realize efficient and smooth transitions between propulsion modes, ensuring high reliability and reduced fuel consumption across the flight trajectory.
Performance Compromises in Auxiliary Functions: Cooling and infrared suppression lead to thrust penalty and increased fuel consumption. Noise reduction is limited and may exacerbate mechanical noise propagation.	Multi-Objective System Optimization and Advanced Materials: Pursue integrated design optimization balancing stealth, thermal management, and propulsion, supported by new materials and structures.	1. Multidisciplinary Design Optimization 2. Lightweight, High-Temperature Composite Materials 3. Innovative Cooling and Acoustic-Lined Structures	High-Efficiency, Low-Signature Exhaust Systems: Deliver optimized performance integrating effective cooling, significant noise and infrared signature reduction, and maintained propulsion efficiency.

. Through these technical approaches, ejector performance can be advanced toward greater efficiency, intelligence, and reliability. Such developments will subsequently support energy conservation, emission reduction, and the sustainable transformation of aerospace systems.