Air-breathing electric hybrid engines have emerged as a promising concept for future atmospheric-to-orbit (ATO) vehicles and very low Earth orbit (VLEO) missions. The primary advantage of this technology lies in its ability to reduce dependence on carried propellant by utilizing ambient atmospheric particles as the working fluid, thereby improving mission efficiency and extending operational lifespan. This review synthesizes the core principles of air-breathing propulsion, electric propulsion, hybrid engine architecture, and mission-level performance analysis, with emphasis on how these systems may enable next-generation space access. The review also examines the principal engineering challenges involved in realizing this concept. These include intake capture efficiency, aerothermal loading, onboard power availability, multi-mode transition stability, and system resilience across varying flight conditions. Performance enhancement strategies such as plasma-assisted combustion, magnetohydrodynamic (MHD) flow control, intake compression and shock optimization, and precooler thermal management are discussed as enabling technologies. In addition, thermodynamic analysis, multi-objective optimization frameworks, and mission-level design methodologies are addressed. The review concludes that air-breathing electric hybrid propulsion remains in an early but promising phase of development, offering a compelling research direction for efficient and reusable ATO systems.

Revolutionary rather than evolutionary changes in propulsion systems are most likely to decrease cost of space transportation and to provide a global range capability. Hypersonic air-breathing propulsion is a revolutionary propulsion system. The performance of scramjet engines can be improved by the AJAX energy management concept. A magneto-hydro-dynamics (MHD) generator controls the flow and extracts flow energy in the engine inlet and a MHD accelerator downstream of the combustor accelerates the nozzle flow. A progress report toward developing the MHD technology is presented herein. Recent theoretical efforts are reviewed and ongoing experimental efforts are discussed. The latter efforts also include an ongoing collaboration between NASA, the US Air Force Research Laboratory, US industry, and Russian scientific organizations. Two of the critical technologies, the ionization of the air and the MHD accelerator, are briefly discussed. Examples of limiting the combustor entrance Mach number to a low supersonic value with a MHD energy bypass scheme are presented, demonstrating an improvement in scramjet performance. The results for a simplified design of an aerospace plane show that the specific impulse of the MHD-bypass system is better than the non-MHD system and typical rocket over a narrow region of flight speeds and design parameters. Equilibrium ionization and non-equilibrium ionization are discussed. The thermodynamic condition of air at the entrance of the engine inlet d

The focus of the NASA / Marshall Space Flight Center (MSFC) Advanced Reusable Technologies (ART) project is to advance and develop Rocket-Based Combined-Cycle (RBCC) technologies. The ART project began in 1996 as part of the Advanced Space Transportation Program (ASTP). The project is composed of several activities including RBCC engine ground testing, tool development, vehicle / mission studies, and component testing / development. The major contractors involved in the ART project are Aerojet and Rocketdyne. A large database of RBCC ground test data was generated for the air-augmented rocket (AAR), ramjet, scramjet, and ascent rocket modes of operation for both the Aerojet and Rocketdyne concepts. Transition between consecutive modes was also demonstrated as well as trajectory simulation. The Rocketdyne freejet tests were conducted at GASL in the Flight Acceleration Simulation Test (FAST) facility. During a single test, the FAST facility is capable of simulating both the enthalpy and aerodynamic conditions over a range of Mach numbers in a flight trajectory. Aerojet performed freejet testing in the Pebble Bed facility at GASL as well as direct-connect testing at GASL. Aerojet also performed sea-level static (SLS) testing at the Aerojet A-Zone facility in Sacramento, CA. Several flight-type flowpath components were developed under the ART project. Aerojet designed and fabricated ceramic scramjet injectors. The structural design of the injectors will be tested in a simulated s

(DMR) under Air Breathing Propulsion Project. The scramjet engine weighed 3,277 kg at lift-off. Critical technologies that have been successfully demonstrated Reusable Launch Vehicle–Technology Demonstration Programme is a series of technology demonstration missions that has been conceived by ISRO as a first step towards realising a Two Stage To Orbit (TSTO) reusable launch vehicle, in which the second stage is a spaceplane. For this purpose, a winged reusable launch vehicle technology demonstrator (RLV-TD) has been configured. The RLV-TD acted as a fly Push… Reusable Launch Vehicle–Technology Demonstration Programme is a series of technology demonstration missions that has been conceived by ISRO as a first step towards realising a Two Stage To Orbit (TSTO) reusable launch vehicle, in which the second stage is a spaceplane. For this purpose, a winged reusable launch vehicle technology demonstrator (RLV-TD) has been configured. The RLV-TD acted as a flying test bed to evaluate various technologies like powered cruise flight, hypersonic flight, and autonomous landing using air-breathing propulsion. Application of these technologies would bring down the launch cost by a factor of 10. This project has no connection with the Avatar spaceplane concept by India's Defence Research and Development Organisation. Pushpak (Sanskrit, ISO: Puṣpaka, lit. 'Little Flower', Namesake: Pushpaka Vimana) (also known as RLV-TD or Reusable Launch Vehicle Technology Demonstrator) is India's first uncrewed flying testbed developed for the ISRO's RLV Technology Demonstration Programme started in 2012. It is a scaled down prototype of an eventual two-stage-to-orbit (TSTO) reusable launch vehicle. In January 2012, the design of ISRO's reusable launch vehicle was approved by the National Review Committee and clearance was granted to build the vehicle. The vehicle was named 'Reusable Launch Vehicle-Technology Demonstrator' (RLV-TD). ISRO aims to bring down the cost of payload delivery to low Earth orbit by 80% from existing $20,000/kg to $4,000/kg. The RLV-TD was developed with an objective to test various aspects such as hypersonic flight, autoland, powered cruise flight, hypersonic flight using the air-breathing engine propulsion and Hypersonic Experiment. A series of four RLV-TD test flights have been planned by ISRO: HEX (Hypersonic Flight Experiment) - Completed LEX (Landing Experiment) - Completed REX (Return Flight Experiment), later renamed to OREX (Orbital Return Flight Experiment) - Planned SPEX (Scramjet Propulsion Experiment) - Planned A team of 750 engineers at the Vikram Sarabhai Space Centre, National Aerospace Laboratories, and Indian Institute of Science worked on the design and development of RLV-TD and the associated rocket. RLV-TD underwent 120 hours of wind tunnel, 5,000 hours of computational fluid dynamics and 1,100 runs of flight simulation tests. RLV-TD has mass of 1.75 tonnes, wingspan of 3.6 meters and overall length of 6.5 meters (excluding the rocket). The vehicle had 600 heat-resistant tiles on its undercarriage and it features delta wings and angled tail fins. Total cost of the project was ₹95 crore (equivalent to ₹137 crore or US$14.2 million in 2023). Future planned developments include testing an air-breathing propulsion system, which aims to capitalise on the oxygen in the atmosphere instead of liquefied oxygen while in flight. The aerodynamic characterization research was conducted at the National Aerospace Laboratories' 1.2m Trisonic Wind Tunnel Facility. The Reusable Launch Vehicle Hypersonic Flight Experiment or RLV HEX was the first test flight in the RLV Technology Demonstration Programme. HEX was successfully launched on 23 May 2016. RLV-TD consists of a fuselage (body), a nose cap, double delta wings and twin vertical rudders. It has active control surfaces called Elevons and Rudders. Apart from the twin rudders it is similar in shape and operation to a small Space Shuttle Orbiter. TDV uses 600 or so heat resistant silica tiles and Flexible External Insulation, nose-cap is made out Carbon-Carbon composite with SiC coating. The leading edges of twin rudders are Inconel-718, wing leading edges of 15CDV6. HEX was the first test flight of a reusable launch vehicle developed by India. The test flight objectives included: Validating the aerodynamic design characteristics during hypersonic flight Characterize induced loads during the hypersonic descent through the atmosphere Assess the performance of the carbon fibre used in construction of the nose of the vehicle Demonstrate first stage separation sequencing The vehicle was tracked during its flight from ground stations at Sriharikota and a shipborne terminal. The total flight duration from launch to splashdown lasted about 773.6 seconds. The unit was not planned to be recovered. ISRO plans to construct an airstrip greater than 4km long in Sriharikota island in the "near future". Critical technologies such as autonomous navigation, guidance & control, reusable thermal protection system, and descent mission management were validated in this flight. Simulating the exact conditions of a Space Re-entry vehicle's landing - high speed, unmanned, autonomous, precise landing from the same return path Validating the landing parameters such as the ground relative velocity, the sinking rate of landing gears and precise body rates as might be experienced by an orbital re-entry space vehicle on its return path After the successful completion of the mission, S. Somanath, chairman of ISRO, said to the media that they are currently planning to conduct more such landing tests to check the readiness of software and hardware under different conditions. The reported test will include the vehicle being dropped from an altitude of about 4.5 kilometres and at a lateral difference following which the vehicle must automatically guide itself for a landing. The test would be retroactively referred to as RLV-LEX-01.

The French Space Agency (Centre National d'Etudes Spatiales-CNES), the UK Space Agency (UKSA), Reaction Engines Limited (REL) and the French Aerospace Lab (Office National d'Etudes et de Recherches Aérospatiales-ONERA) share a common interest for future space launch architectures and technologies. Preliminary results of a joint study are presented in this paper with the main objective of assessing and quantifying the potential advantages and benefits of air-breathing propulsion technology developed by REL named SABRE (Synergetic Air-Breathing Rocket Engine) for future launch architectures. In

This review provides a comprehensive analysis of acceleration mechanisms utilized in air-breathing electric propulsion, focusing on their fundamental principles, advantages, and the latest technological advancements. These thrusters, which utilize atmospheric gases to generate plasma and produce thrust, hold significant promise for very low Earth orbit missions due to their potential for high-efficiency propulsion. Central to their operation are the different mechanisms of acceleration. The review systematically categorizes the various acceleration mechanisms and discusses the physical principles behind these mechanisms, their integration into air-breathing propulsion architectures, and recent experimental efforts aimed at performance optimization. 3205 npjmicrogr NPJ Microgravity NPJ Microgravity Nature Publishing Group PMC13230546 13230546 13230546 41922351 10.1038/s41526-026-00573-5 The quest and opportunities for air-breathing propulsion Duppada Guru Sankar 1 ✉ # Taploo Anmol 1 # Zhuang Taisen 1 Spinelli Jake 1 Keidar Michael 1 ✉ 1 George Washington University, Washington, DC USA ✉ Corresponding author. # Contributed equally. If material is not included in the article’s Creative Commons licence and your intended use is not permitted by statutory regulation or exceeds the permitted use, you will need to obtain permission directly from the copyright holder. To view a copy of this licence, visit http://creativecommons.org/licenses/by/4.0/ . Abstract This review provides a comprehensive analysis of acceleration mechanisms utilized in air-breathing electric propulsion, focusing on their fundamental principles, These thrusters, which utilize atmospheric gases to generate plasma and produce thrust, hold significant promise for very low Earth orbit missions due to their potential for high-efficiency propulsion. Central to their operation are the different mechanisms of acceleration. The review systematically categorizes the various acceleration mechanisms and discusses the physical principles behind these mechanisms, their integration into air-breathing propulsion architectures, and recent experimental efforts aimed at performance optimization. Another significant limitation of chemical propulsion is its relatively low specific impulse (Isp), which typically ranges from 300 to 450 s depending on the specific propellant combination and engine design. To overcome these limitations 2 – 4 , the focus has shifted toward electric and plasma-based propulsion systems, particularly those capable of utilizing environmental resources in situ, thereby enabling extended operational periods, and reducing overall mission costs. A tradeoff study between a satellite with an air-breathing electric propulsion (ABEP) and a traditional (Xenon-based) electric propulsion systems has been performed. In addition to these developments, the operation of traditional thrusters on molecular propellants such as air molecules has been explored. While conventional electric thrusters like Hall-effect thrusters and ion engines are typically optimized for noble gases, their operation with molecular propellants introduces additional challenges, including energy losses due to the ionization of molecules like N₂ and O₂. Advances in grid materials, such as carbon-carbon composites and molybdenum alloys, as well as improved grid geometries, are ongoing to address these issues 41 , 42 .Furthermore, the same gridded acceleration technique can, in principle, be adapted to air-breathing ion engines, where ionization and acceleration are performed using ambient atmospheric gases. Recent more detailed studies suggested that air-breathing thruster can operate at altitudes of 80-100 km without the need for air compression, utilizing a scramjet-like configuration 46 . To that end, efficient air ionization approach based on electron beam ionization has been developed and implemented 47 . In addition, it has been discovered that air-breathing thruster operation at these altitudes provides conditions for self-neutralization 48 .The ultimate requirement for electric propulsion systems, especially considering limited onboard resources, is to attain a high thrust-to-power (T/P) ratio 49 , which can be expressed as: 1 T P ≈ 2 V ex − Vo where V ex is the exhaust velocity and Vo is the initial velocity of the air entering the inlet, i.e. flight speed. It should be noted that Eq. Figure 8 shows the relationship between Isp and plasma-generated thrust for an air-breathing MPD thruster operating under hypersonic electric propulsion conditions 50 . The results show a systematic increase in Isp with increasing plasma thrust, reflecting more effective acceleration as discharge energy and current density increase. Variations in acceleration voltage and pulse repetition rate shift the operating points along this Isp–thrust envelope, with higher repetition rates generally enabling higher Isp at a given thrust level. In this altitude range of 65-120 km, the availability of dense incoming flow allows the MPD thruster to operate in a regime where both thrust and Isp can be increased simultaneously, as reflected by the upper envelope of the data in Fig. 8 . This behaviour supports the suitability of air-breathing MPD concepts for hypersonic propulsion and drag-compensation missions at extremely low altitudes. Fig. 8 Specific Impulse as a function of thrust due to plasma. Isp is calculated based on experimental data presented in Ref. 50 . It should be noted that the reported thrust and thrust-to-power values for Aether HT5k, PPS1350, and RIT 10 thrusters correspond to ideal laboratory operating conditions, in which neutral particles are supplied using mass flow controllers or particle flow generators at densities several orders of magnitude higher than those available in VLEO. These results should therefore be interpreted as demonstrations of the acceleration stage rather than complete air-breathing propulsion performance. ABEP in refs. 2 – 5 .

British company Reaction Engines Limited, using SABRE, a combined-cycle, air-breathing rocket propulsion system. The vehicle design is for a hydrogen-fuelled Skylon was a series of concept designs for a reusable single-stage-to-orbit spaceplane by the British company Reaction Engines Limited, using SABRE, a combined-cycle, air-breathing rocket propulsion system. The vehicle design is for a hydrogen-fuelled aircraft that would take off from a specially built reinforced runway, and accelerate to Mach 5.4 at 26 kilometres (85,000 ft) altitude (compared to In… The Skylon was to be a fully reusable SSTO vehicle, able to achieve orbit without staging, which was intended to be used principally as a reusable launch system. Proponents of the SSTO approach have often claimed that staging involves a number of inherent complications and problems due to complexity, such as being difficult or typically impossible to recover and reuse most elements, thus unavoidably incurring great expense to produce entirely new launch vehicles instead; therefore, they believe that SSTO designs hold the promise of providing a reduction to the high cost of space flights. Operationally, it was envisioned for the non-crewed Skylon to take off from a specially strengthened runway, gain altitude in a fashion akin to a conventional aeroplane and perform an ascent at very high speeds, in excess of five times the speed of sound (6,100 km/h or 3,800 mph), to attain a peak air-breathing altitude of roughly 28 kilometres (92,000 ft). At that altitude the engine would switch to rocket mode, using onboard oxidiser, and ascend to orbital altitude and velocity. Payloads would typically be deployed prior to the vehicle's re-entry into the atmosphere, upon which it would conduct a relatively gentle descent before performing a traditional landing upon a runway. Skylon was a series of concept designs for a reusable single-stage-to-orbit spaceplane by the British company Reaction Engines Limited, In 2000, the firm completed work with University of Bristol testing the precooler. From 2007 to 2009 Reaction worked with University of Bristol and Airborne Engineering on Project STERN (Static Test Expansion/Deflection Rocket Nozzle), which tested Reaction's engine ignition system, a Reaction designed air breathing hydrogen rocket engine, and investigated the flow stability and behaviour of Dr Neil Taylor's expansion deflection nozzle design via multiple test-firings by Airborne Engineering. An expansion deflection nozzle is capable of compensating for the changing ambient pressure encountered while gaining altitude during atmospheric flight, thus generating greater thrust and thereby efficiency. Work on STERN was continued in project STRICT (Static Test Rocket Incorporating Cooled Thrust-chamber), which investigated the stability of the engine's exhaust flow and the dissipation of the generated heat into the engine walls. The results and designs delivered by both the STRICT and STERN projects were subsequently declared by Reaction to have been "a great success". Static testing of the engine precooler began in June 2011, marking the start of Phase 3 in the Skylon development programme, In April 2012, Reaction announced that the first series of the precooler test programme had been successfully completed. On 10 July 2012, Reaction announced that the second of three series of tests has been completed successfully, and the final series of tests would begin the following month after the testing facilities had been upgraded to allow testing of −150 °C (−238 °F) temperatures. ESA's propulsion division audited the precooler tests during mid-2012 and found the results satisfactory. In 2011, Reaction stated that a preproduction prototype of the Skylon could be flying by 2016, and the proposed route would be a suborbital flight between the Guiana Space Centre near Kourou in French Guiana and the North European Aerospace Test Range, located in northern Sweden. Pre-orders were expected in the 2011–2013 timeframe, coinciding with the formation of the manufacturing consortium. In December 2011, Alan Bond stated that Skylon would enter into service by 2021–2022 instead of 2020 as previously envisaged. In November 2012, Reaction announced that it would begin a three-and-a-half-year project to develop and build a test rig of the SABRE engine to prove its performance in both the air-breathing and rocket modes. Reaction's precooler was tested in 2022 and 2024. Work ceased later in 2024, when the company entered administration. In July 2025, ESA announced that the precooler would be used in their new INVICTUS research programme, which aims to demonstrate hypersonic flight in the atmosphere at Mach 5. The Skylon was to be a fully reusable SSTO vehicle, able to achieve orbit without staging, which was intended to be used principally as a reusable launch system. Proponents of the SSTO approach have often claimed that staging involves a number of inherent complications and problems due to complexity, such as being difficult or typically impossible to recover and reuse most elements, thus unavoidably incurring great expense to produce entirely new launch vehicles instead; therefore, they believe that SSTO designs hold the promise of providing a reduction to the high cost of space flights. Operationally, it was envisioned for the non-crewed Skylon to take off from a specially strengthened runway, gain altitude in a fashion akin to a conventional aeroplane and perform an ascent at very high speeds, in excess of five times the speed of sound (6,100 km/h or 3,800 mph), to attain a peak air-breathing altitude of roughly 28 kilometres (92,000 ft). At that altitude the engine would switch to rocket mode, using onboard oxidiser, and ascend to orbital altitude and velocity. Payloads would typically be deployed prior to the vehicle's re-entry into the atmosphere, upon which it would conduct a relatively gentle descent before performing a traditional landing upon a runway.

We report on the results of numerical-simulation investigations of ignition characteristics of hydrocarbon-fuel blends expected from thermal cracking of typical jet fuels, at conditions relevant to high-Mach-number, air-breathing propulsion. A two-point-continuation method was employed, with a detailed description of molecular transport and chemical kinetics, focusing on the effects of fuel composition, reactant temperature, additives, and imposed strain rate. It captured the entire S-curve that describes the processes of vigorous burning extinction, and ignition. The results demonstrate that

Determination of specific fuel consumption of air-breathing engines is one of the problems of modeling their performance. As a rule, the estimation error of the specific fuel consumption while calculating air-breathing engine performance is greater than that of thrust. In this work, this is substantiated by the estimation error of the fuel-air ratio, which weakly affects thrust but significantly affects the specific fuel consumption. The presence of a significant error in the fuel-air ratio is explained by the use of simplified methods, which use the dependence of enthalpy as a function of mix