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Aerospace
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Space Propulsion: Advances and Challenges (3rd Volume)
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Space Propulsion: Advances and Challenges (3rd Volume)

A special issue of Aerospace (ISSN 2226-4310). This special issue belongs to the section "Astronautics & Space Science".

Deadline for manuscript submissions: 31 August 2025 | Viewed by 16497

Special Issue Editor

Prof. Dr. Kyun Ho Lee

E-Mail Website
Guest Editor
Department of Aerospace Engineering, Sejong University, Seoul 143-741, Republic of Korea
Interests: space propulsion; satellite system; thermal engineering; CFD; inverse heat transfer analysis; rarefied flow
Special Issues, Collections and Topics in MDPI journals

Special Issue Information

Dear Colleagues,

A major function of space propulsion systems is to accelerate spacecraft by producing a propulsive force (thrust) or a change in velocity (delta-V) by ejecting propellant mass at a high speed into the air or space based on Newton’s laws of motion. This plays an important role in the acceleration, attitude control, drag make-up, and orbit transfer maneuvers of spacecraft. The various types of space propulsion systems can be defined depending on what kind of energy source is used and how the energy is generated to provide thrust. At present, chemical and electric propulsion systems are the preferred types of systems for various spacecrafts. Applications of space propulsion can be classified into three different categories: escape propulsion (from Earth’s surface to its orbit), in-space propulsion (in Earth’s orbit), and deep space propulsion (from Earth’s orbit to outer space).

Since Goddard‘s first successful flight of a liquid propellant rocket in 1926, the roles of space propulsion have become more important and complex for the successful completion of predefined mission goals, as recent demands on the function of space propulsion have diversified. Thus, various new and advanced concepts of space propulsion technologies are under investigation and development, especially for small-lift launch vehicles, reusable launch vehicles, Earth-orbiting satellites, deep space explorers, cubesats, and many other spacecraft applications.

This Special Issue invites contributions relating to recent advances and challenges for space propulsion technologies. Submissions are welcome from a whole range of space propulsion topics, including, but not limited to:

  • Concept, theory, and related science and engineering;
  • Design, modeling, simulation, and analysis;
  • Mission and application;
  • Launch and flight/orbit operation;
  • Experiment, test, and verification;
  • Propellant (solid, liquid, gas, non-toxic, gelled, etc.);
  • Thrust generation method and type (chemical, electric, hybrid, solar sail, nuclear, etc.);
  • Hardware (material, part, component, equipment, assembly, and system) and software;
  • Manufacturing, integration, and facility.

Prof. Dr. Kyun Ho Lee
Guest Editor

Manuscript Submission Information

Manuscripts should be submitted online at www.mdpi.com by registering and logging in to this website. Once you are registered, click here to go to the submission form. Manuscripts can be submitted until the deadline. All submissions that pass pre-check are peer-reviewed. Accepted papers will be published continuously in the journal (as soon as accepted) and will be listed together on the special issue website. Research articles, review articles as well as short communications are invited. For planned papers, a title and short abstract (about 100 words) can be sent to the Editorial Office for announcement on this website.

Submitted manuscripts should not have been published previously, nor be under consideration for publication elsewhere (except conference proceedings papers). All manuscripts are thoroughly refereed through a single-blind peer-review process. A guide for authors and other relevant information for submission of manuscripts is available on the Instructions for Authors page. Aerospace is an international peer-reviewed open access monthly journal published by MDPI.

Please visit the Instructions for Authors page before submitting a manuscript. The Article Processing Charge (APC) for publication in this open access journal is 2400 CHF (Swiss Francs). Submitted papers should be well formatted and use good English. Authors may use MDPI's English editing service prior to publication or during author revisions.

Keywords

  • space propulsion
  • chemical propulsion
  • electric propulsion
  • hybrid propulsion
  • solar sail propulsion
  • nuclear propulsion
  • spacecraft
  • rocket
  • launch vehicle
  • satellite
  • cubesat
  • deep space explorer

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Related Special Issues

Published Papers (8 papers)

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Research

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24 pages, 11695 KiB  
Article
Experimental Investigation of PWM Throttling in a 50-Newton-Class HTP Monopropellant Thruster: Analysis of Pressure Surges and Oscillations
by Suk Min Choi and Christian Bach
Aerospace 2025, 12(5), 418; https://doi.org/10.3390/aerospace12050418 - 8 May 2025
Viewed by 216
Abstract
High-test peroxide (HTP) monopropellant thrusters are being considered for spacecraft lander missions due to their simplicity and reduced toxicity compared to traditional propellants. Pulse-Width Modulation (PWM) throttling is a key technique for precise thrust control in such systems. However, PWM throttling can lead [...] Read more.
High-test peroxide (HTP) monopropellant thrusters are being considered for spacecraft lander missions due to their simplicity and reduced toxicity compared to traditional propellants. Pulse-Width Modulation (PWM) throttling is a key technique for precise thrust control in such systems. However, PWM throttling can lead to pressure surges and oscillations in the propellant feed system, potentially compromising system reliability. This study investigates the influence of PWM parameters, specifically duty cycle and frequency, on pressure surges and oscillations in a 50-newton-class HTP monopropellant thruster. The objective is to identify stable operating conditions that mitigate these effects, thereby enhancing the reliability of PWM throttling for lander applications. An experimental setup was developed, including a 50-newton-class thruster with a MnO2/La/Al2O3 catalyst and a solenoid valve for PWM control. Cold flow tests using water characterized the valve response and water hammer effects, while hot fire tests with 90 wt.% HTP were used to evaluate thruster performance under steady-state and PWM conditions. Analytical methods, including Joukowsky’s equation and power spectral density analysis, were used to interpret the data and understand the underlying mechanisms. The results showed that while surge pressures generally aligned with steady-state values, specific PWM conditions led to amplified surges, particularly at low duty cycles. Additionally, high duty cycles induced chugging instability. The natural frequencies of the feed system were found to play a crucial role in these phenomena. Stable operating conditions were identified by avoiding duty cycles that cause constructive interference of pressure waves. This research demonstrates that by carefully selecting PWM parameters based on the feed system’s dynamic characteristics, pressure surges and oscillations can be minimized, ensuring reliable operation of HTP monopropellant thrusters in PWM throttling mode. These findings contribute to the development of more efficient and safer propulsion systems for spacecraft landers. Full article
(This article belongs to the Special Issue Space Propulsion: Advances and Challenges (3rd Volume))
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29 pages, 6040 KiB  
Article
Properties and Behavior of 3D-Printed ABS Fuel in a 10 N Hybrid Rocket: Experimental and Numerical Insights
by Sergio Cassese, Veniero Marco Capone, Riccardo Guida, Stefano Mungiguerra and Raffaele Savino
Aerospace 2025, 12(4), 291; https://doi.org/10.3390/aerospace12040291 - 30 Mar 2025
Viewed by 316
Abstract
In a global landscape where the launch of satellites into space is growing exponentially, there is an increasing demand for propulsion solutions to perform various types of maneuvers. In this context, the present study aims to investigate a 3D-printed ABS (Acrylonitrile Butadiene Styrene)-based [...] Read more.
In a global landscape where the launch of satellites into space is growing exponentially, there is an increasing demand for propulsion solutions to perform various types of maneuvers. In this context, the present study aims to investigate a 3D-printed ABS (Acrylonitrile Butadiene Styrene)-based fuel for use in a 10 N-scale hybrid rocket in order to promote cost-effective and environmentally friendly access to space. As this material is currently unknown in this field and lacks a thermodynamic database, characterization of its pyrolysis process was carried out through a mixed approach combining experimental data and numerical simulations. The experiments show excellent performance of the H2O2-3D-printed ABS pair; despite the lack of information on its thermodynamically relevant quantities, it was possible to accurately reconstruct the fuel consumption profile as well as its regression rate and the spatial and temporal average values using the numerical model and Arrhenius parameters derived in this work. The methodology and results obtained herein represent tools that can be useful for the design of small-scale rockets using 3D-printed ABS-based fuels as well as a starting point for the development and analysis of the complex geometries made possible through additive manufacturing. Full article
(This article belongs to the Special Issue Space Propulsion: Advances and Challenges (3rd Volume))
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30 pages, 9515 KiB  
Article
RANS Simulations of Advanced Nozzle Performance and Retro-Flow Interactions for Vertical Landing of Reusable Launch Vehicles
by Giuseppe Scarlatella, Jan Sieder-Katzmann, Martin Propst, Theodor Heutling, Jannis Petersen, Felix Weber, Marco Portolani, Marco Garutti, Daniele Bianchi, Dario Pastrone, Andrea Ferrero, Martin Tajmar and Christian Bach
Aerospace 2025, 12(2), 124; https://doi.org/10.3390/aerospace12020124 - 6 Feb 2025
Viewed by 960
Abstract
In recent years, advanced nozzle concepts have attracted interest because of advancements in their technology readiness level and studies on applications to vertical take-off and landing reusable launch vehicles. This is ascribable to their intrinsic altitude compensation properties, which could mitigate the additional [...] Read more.
In recent years, advanced nozzle concepts have attracted interest because of advancements in their technology readiness level and studies on applications to vertical take-off and landing reusable launch vehicles. This is ascribable to their intrinsic altitude compensation properties, which could mitigate the additional propellant cost resulting from the vertical landing manoeuvres based on retro-propulsion. Experimental and numerical campaigns at the Technical University of Dresden test the performance of annular-aerospike, dual-bell, and expansion-deflection nozzles compared with conventional bell-shaped nozzles in various subsonic counter-flow regimes and atmospheric conditions. The methods of investigation and a detailed description of the experimental and numerical results are reported. More specifically, the study offers a comparison between advanced and conventional nozzles, with a focus on nozzle performance through experiments and aerodynamic performance and retro-flow interaction through simulations. The flow topology that is established within the area of interaction between nozzle jets and counter-flows is detailed, with the advantages and limitations of each advanced nozzle in terms of adaptive performance. The numerical simulations confirm that advanced nozzles achieve altitude compensation in retro-flow configurations. Moreover, the distance obtained from the models for jet penetration into subsonic counter-flows is compatible with empirical formulations available in the literature. Full article
(This article belongs to the Special Issue Space Propulsion: Advances and Challenges (3rd Volume))
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26 pages, 8455 KiB  
Article
Re-Entry Comparison of a Spacecraft in Low Earth Orbit: Propulsion-Assisted vs. Non-Propulsive Configurations
by Antonio Sannino, Dylan De Prisco, Sergio Cassese, Stefano Mungiguerra, Anselmo Cecere and Raffaele Savino
Aerospace 2025, 12(2), 79; https://doi.org/10.3390/aerospace12020079 - 23 Jan 2025
Viewed by 1185
Abstract
This paper presents a mission concept for a Low Earth Orbit (LEO) satellite equipped with a payload for space experiments, designed to be recovered on Earth post-mission. The focus of this study is on developing a mission concept with fast de-orbit and accurate [...] Read more.
This paper presents a mission concept for a Low Earth Orbit (LEO) satellite equipped with a payload for space experiments, designed to be recovered on Earth post-mission. The focus of this study is on developing a mission concept with fast de-orbit and accurate landing capability for a small satellite payload. Two re-entry configurations are analyzed: one employing a deployable aero-brake heat shield for aerodynamic descent and another integrating a propulsion system. Aerodynamic analysis of the capsule, including drag coefficient and stability at relevant altitudes, was conducted using the Direct Simulation Monte Carlo (DSMC) method. A trade-off analysis, accounting for uncertainties such as CD, atmospheric density, and ignition timing, revealed significant differences in mission profiles. A propulsion system providing a ΔV of approximately 100 m/s reduces descent time from 54 days (aerodynamic-only re-entry) to under 1 h, without altering trajectory. Drag-related uncertainties contribute to a landing dispersion of ~100 km, while a ±1% error in total impulse increases dispersion to 400 km. A monopropellant rocket engine was preliminarily designed, meeting constraints such as catalytic chamber pressure and performance targets. The resulting thruster, weighing under 4 kg and contained within a 250 mm-high, 350 mm-diameter cylinder, supports three potential component layouts. Full article
(This article belongs to the Special Issue Space Propulsion: Advances and Challenges (3rd Volume))
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22 pages, 6486 KiB  
Article
Fuel Ignition in HTP Hybrid Rockets at Very Low Mass Fluxes: Challenges and Pulsed Preheating Techniques Using Palladium-Coated Catalysts
by Sergio Cassese, Stefano Mungiguerra, Veniero Marco Capone, Riccardo Guida, Anselmo Cecere and Raffaele Savino
Aerospace 2024, 11(11), 884; https://doi.org/10.3390/aerospace11110884 - 26 Oct 2024
Cited by 2 | Viewed by 1359
Abstract
In a worldwide scenario which sees an increasing number of small satellite launches, novel mission concepts may be unlocked providing the spacecrafts with the very precise and rapid maneuvering capability that electric thrusters cannot guarantee. In this context, chemical thrusters appear to be [...] Read more.
In a worldwide scenario which sees an increasing number of small satellite launches, novel mission concepts may be unlocked providing the spacecrafts with the very precise and rapid maneuvering capability that electric thrusters cannot guarantee. In this context, chemical thrusters appear to be a possible solution. This work aimed to experimentally study and solve the problem of ignition for 10 N hybrid rockets based on hydrogen peroxide. Firstly, the study analyzed the performance of a monopropellant engine capable of functioning as a hybrid injection system. In particular, the effects of the liquid mass injected, the initial temperature, and the supply pressure on the pulsed engine performance were experimentally investigated. The injected mass showed a greater impact on the performance with respect to the starting chamber temperature and injection pressure. This thruster also showed a good potential for space applications. In the second part of the work, the objective was to find an ignition procedure that reduced propellant consumption and eliminated the need for a glow plug. This is important because the electrical power consumption in real applications significantly affects other subsystems and is undesirable for chemical engines. Different ignition procedures were tested to emphasize their respective advantages and disadvantages, and the findings indicated that the concept of pulsed preheating is feasible with only a small amount of propellant consumption, while substantially decreasing the ignition duration from approximately 45 min to a maximum of just 3 min. Finally, similar ignition procedures were adopted using different fuels. The results showed that PVC and ABS, under the same operating conditions, ignite more easily than HDPE, which requires an oxidizer consumption approximately double that of the other two fuels. Considerations about the effect of chamber pressure and oxidizer mass flow rate on engine ignition were also included. Full article
(This article belongs to the Special Issue Space Propulsion: Advances and Challenges (3rd Volume))
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20 pages, 4559 KiB  
Article
Turbopump Parametric Modelling and Reliability Assessment for Reusable Rocket Engine Applications
by Mateusz T. Gulczyński, Robson H. S. Hahn, Jan C. Deeken and Michael Oschwald
Aerospace 2024, 11(10), 808; https://doi.org/10.3390/aerospace11100808 - 2 Oct 2024
Cited by 1 | Viewed by 2554
Abstract
The development of modern reusable launchers, such as the Themis project with its LOX/LCH4 Prometheus engine, CALLISTO—a reusable VTVL-launcher first-stage demonstrator with a LOX/LH2 RSR2 engine, and SpaceX’s Falcon 9 with its Merlin 1D engine, underscores the need for advanced control algorithms to [...] Read more.
The development of modern reusable launchers, such as the Themis project with its LOX/LCH4 Prometheus engine, CALLISTO—a reusable VTVL-launcher first-stage demonstrator with a LOX/LH2 RSR2 engine, and SpaceX’s Falcon 9 with its Merlin 1D engine, underscores the need for advanced control algorithms to ensure reliable engine operation. The multi-restart capability of these engines imposes additional requirements for throttling, necessitating an extended controller-validity domain to safely achieve low thrust levels across various operating regimes. This capability also increases the risk of component failure, especially as engine parameters evolve with mission profiles. To address this, our study evaluates the dynamic reliability of reusable rocket engines (RREs) and their subcomponents under different failure modes using multi-physics system-level modelling and simulation, with a particular focus on turbopump components. Transient condition modelling and performance analysis, conducted using EcosimPro-ESPSS software (version 6.4.34), revealed that turbopump components maintain high reliability under nominal conditions, with turbine blades demonstrating significant fatigue life even under varying thermal and mechanical loads. Additionally, the proposed predictive model estimates the remaining useful life of critical components, offering valuable insights for improving the longevity and reliability of turbopumps in reusable rocket engines. This study employs deterministic, thermally dependent structural simulations, with key control objectives including end-state tracking of combustion chamber pressure and mixture ratios and the verification of operational constraints, exemplified by the LUMEN demonstrator engine and the LE-5B-2 engine class. Full article
(This article belongs to the Special Issue Space Propulsion: Advances and Challenges (3rd Volume))
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31 pages, 19796 KiB  
Article
Effect of Multicoaxial Injectors on Nitrogen Film Cooling in a GCH4/GO2 Thrust Chamber for Small-Scale Methane Rocket Engines: A CFD Study
by Kanmaniraja Radhakrishnan, Dong Hwi Ha and Hyoung Jin Lee
Aerospace 2024, 11(9), 744; https://doi.org/10.3390/aerospace11090744 - 11 Sep 2024
Viewed by 1003
Abstract
Improper film cooling design and positioning of an injector in the face plate cause thermal damage to the thrust chamber wall and lead to rocket engine failures. An experimental combustor with five shear coaxial injectors was damaged owing to inadequate film cooling distribution [...] Read more.
Improper film cooling design and positioning of an injector in the face plate cause thermal damage to the thrust chamber wall and lead to rocket engine failures. An experimental combustor with five shear coaxial injectors was damaged owing to inadequate film cooling distribution on the thrust chamber wall. The present study aimed to simulate the experimental test case and investigate the causes of the thermal damage. In the simulation, gaseous methane and oxygen were injected at the inner and outer inlets of the shear coaxial injectors and nitrogen, used as the coolant, was injected near the upstream of the chamber wall. The turbulent chemistry interaction was modeled using a reduced DRM-19 mechanism by incorporating the Eddy Dissipation Concept model. Numerical investigations were conducted to examine the cause of thermal damage. The temperature contours of the thrust chamber wall were compared with the experimental image of the damaged wall. Further, simulations of single-row (SR) and multi-row (MR) injector configurations were conducted to assess the effect on film cooling distribution. The adiabatic film cooling effectiveness and specific impulse were determined for all simulated cases. The results showed that MR simulations with narrow injector angles had poor film cooling performance, while wider angles led to lower specific impulse. The face plate with an angle of 15 degrees between the injector positions showed better performance in terms of considering both the film cooling and specific impulse. Full article
(This article belongs to the Special Issue Space Propulsion: Advances and Challenges (3rd Volume))
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Review

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50 pages, 10198 KiB  
Review
A Review of Recent Developments in Hybrid Rocket Propulsion and Its Applications
by Shih-Sin Wei, Meng-Che Li, Alfred Lai, Tzu-Hao Chou and Jong-Shinn Wu
Aerospace 2024, 11(9), 739; https://doi.org/10.3390/aerospace11090739 - 9 Sep 2024
Cited by 5 | Viewed by 8351
Abstract
This paper extensively reviews hybrid rocket propulsion-related activities from combustion engine designs to launch tests. Starting with a brief review of rocket propulsion development history, a comparison among the three bi-propellant rocket propulsion approaches, and hybrid rocket engine design guidelines, a very thorough [...] Read more.
This paper extensively reviews hybrid rocket propulsion-related activities from combustion engine designs to launch tests. Starting with a brief review of rocket propulsion development history, a comparison among the three bi-propellant rocket propulsion approaches, and hybrid rocket engine design guidelines, a very thorough review related to hybrid rocket propulsion and its applications is presented in this paper. In addition to propellant choice, engine design also affects the hybrid rocket performance and, therefore, a variety of engine designs, considering, e.g., fuel geometry, swirl injection, ignition designs, and some innovative flow-channel designs are also explored. Furthermore, many fundamental studies on increasing hybrid rocket engine performances, such as regression rate enhancement, mixing enhancement, and combustion optimization, are also reviewed. Many problems that will be encountered for practical applications are also reviewed and discussed, including the O/F ratio shift, low-frequency instability, and scale-up methods. For hybrid rocket engine applications in the future, advanced capabilities and lightweight design of the hybrid rocket engine, such as throttling capability, thrust vectoring control concept, insulation materials, 3D-printing manufacturing technologies, and flight demonstrations, are also included. Finally, some active hybrid rocket research teams and their plans for flight activities are briefly introduced. Full article
(This article belongs to the Special Issue Space Propulsion: Advances and Challenges (3rd Volume))
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