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Aerospace
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Green Propellants for In-Space Propulsion
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Green Propellants for In-Space Propulsion

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

Deadline for manuscript submissions: 31 May 2025 | Viewed by 6400

Special Issue Editors

Dr. Christoph Kirchberger

E-Mail Website
Guest Editor
Satellite and Orbital Propulsion Department, Institute of Space Propulsion, German Aerospace Center (DLR), D-74239 Hardthausen, Germany
Interests: the research focus of Dr. Kirchberger and his team is the applicability, compatibility and handling of both known and new energetic materials for rocket propulsion. Their work covers topics regarding advanced and “green” propellants for satellite and orbital propulsion, i.e., the application of energetic ionic liquids, nitrous oxide-based mixtures (HyNOx, NOFBx), hypergolic fuels and high test peroxide (H2O2), as well as gelled propellants, hybrid rocket propulsion, and the challenges of high-speed air-breathing/scramjet propulsion. He is the author or co-author of more than 80 scientific publications so far
Dr. Angelo Pasini

E-Mail Website
Guest Editor
Department of Civil and Industrial Engineering, Aerospace Division, University of Pisa, 56122 Pisa, Italy
Interests: Prof. Angelo Pasini is an assistant professor of aerospace propulsion at the University of Pisa with strong industrial experience. His research activities have been mainly on testing green propellant rockets and on experimental campaigns on pumps for liquid propellant rockets. His fields of interest are turbomachinery, cavitation, rotor dynamics and flow instabilities in space rocket turbopumps, high-speed hybrid bearings, non-toxic propellants and catalytic beds for hydrogen peroxide decomposition. He is the author or co-author of about 100 scientific publications so far
Dr. Ulrich Gotzig

E-Mail Website
Guest Editor
ArianeGroup, Lampoldshausen, 74239 Hardthausen, Germany
Interests: in-space storable; propulsion system; space electric propulsion; development, qualification and industrial series manufacturing of propulsion systems and components; green propellants

Special Issue Information

Dear Colleagues,

The propulsion systems of spacecraft currently rely on high-performance yet storable propellants, which are traditionally hydrazine-based fuels and nitrogen oxides. These propellants, while well investigated and effective, are highly toxic, posing significant health and environmental risks. Handling these chemicals requires stringent safety protocols, leading to high operational costs and potential regulatory restrictions. To address these challenges, so-called green propellants—alternatives that are less toxic, non-carcinogenic, and easier to handle—have been investigated and developed in recent years. Green propellants aim to offer similar or improved performance compared to traditional propellants while minimizing health risks and environmental impact.

Key examples of green monopropellants include hydroxylammonium nitrate fuel/oxidizer mixtures like ASCENT (AF-M315E), which provides higher performance and lower toxicity than hydrazine, and LMP-103S, a blend featuring ammonium dinitramide (ADN) that is suitable for various space missions and which already features some flight heritage. Highly concentrated hydrogen peroxide—also known as high-test peroxide (HTP)—and nitrous oxide are also potential green monopropellants but are only as well relevant as oxidizers in green bipropellant systems. In combination with a catalyst or suitable additive in fuel, such systems may be hypergolic, i.e., they may operate without a dedicated ignition system, reducing complexity and increasing reliability.

The advancements in green propellant technology reflect a broader industry trend towards safer and more sustainable space exploration. This Special Issue seeks to provide an overview of green propellants and the latest propulsion technologies, highlighting significant developments and their potential benefits for future space missions.

Dr. Christoph Kirchberger
Dr. Angelo Pasini
Dr. Ulrich Gotzig
Guest Editors

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

  • rocket propulsion
  • space propulsion
  • in-space propulsion
  • satellite propulsion
  • lander propulsion
  • advanced propellants
  • green propellants
  • hydrazine replacement
  • non-toxic fuels
  • less-toxic fuels
  • hydrogen peroxide
  • high-test peroxide
  • H2O2
  • nitrous oxide
  • N2O
  • water propulsion
  • catalyst
  • ionic liquids
  • energetic ionic liquids
  • EIL
  • nitromethane
  • hypergolic
  • ammonium dinitramide
  • ADN
  • LMP-103S
  • ASCENT
  • AF-M315E
  • hydroxylammonium nitrate
  • HAN
  • SHP163
  • Life Cycle Analysis (LCA)
  • sustainability

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Published Papers (4 papers)

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Research

10 pages, 1865 KiB  
Article
Theoretical Research on the Combustion Characteristics of Ammonium Dinitramide-Based Non-Toxic Aerospace Propellant
by Jianhui Han, Ming Wen, Yanji Hong, Baosheng Du, Luyun Jiang, Haichao Cui, Gaoping Feng and Junling Song
Aerospace 2025, 12(4), 295; https://doi.org/10.3390/aerospace12040295 - 31 Mar 2025
Viewed by 234
Abstract
Propellants play a crucial role in the propulsion systems of aerospace vehicles, and their combustion characteristics are susceptible to external environmental conditions. This study systematically investigated the impact of various initial conditions on the combustion process of ADN-based propellant, including combustion products, equilibrium [...] Read more.
Propellants play a crucial role in the propulsion systems of aerospace vehicles, and their combustion characteristics are susceptible to external environmental conditions. This study systematically investigated the impact of various initial conditions on the combustion process of ADN-based propellant, including combustion products, equilibrium pressure, adiabatic temperature, and ignition delay time. The results indicate that the primary combustion products of ADN-based propellant include N2O, N2, CO2, OH, and others. ADN-based propellant exhibits a distinct two-stage combustion process under low pressure and temperature conditions (P0 = 2 atm, T0 = 586 K). Conversely, under high pressure and temperature conditions (P0 = 10 atm, T0 = 2930 K), the two stages of combustion occur almost simultaneously, making them difficult to distinguish. Furthermore, as the initial temperature increases, the ignition delay time decreases significantly, and the combustion rate accelerates. When the initial temperature rises from 400 K to 2800 K at a pressure of P0 =10 atm, the ignition delay time decreases from 3.5 ms to 0.6 μs. Interestingly, changes in initial pressure have a relatively minor impact on the ignition delay time compared to changes in temperature. Therefore, temperature has a more crucial influence on the combustion characteristics of ADN-based propellant than pressure. This study holds promise for providing new combustion optimization strategies for the aerospace industry and promoting the development of aircraft designs towards higher performance and sustainability. Full article
(This article belongs to the Special Issue Green Propellants for In-Space Propulsion)
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12 pages, 6085 KiB  
Article
Demonstration of Polyethylene Nitrous Oxide Catalytic Decomposition Hybrid Thruster with Dual-Catalyst Bed Preheated by Hydrogen Peroxide
by Seungho Lee, Vincent Mario Pierre Ugolini, Eunsang Jung and Sejin Kwon
Aerospace 2025, 12(2), 158; https://doi.org/10.3390/aerospace12020158 - 18 Feb 2025
Viewed by 448
Abstract
Although various studies on nitrous oxide as a prospective green propellant have been recently explored, a polyethylene nitrous oxide catalytic decomposition hybrid thruster was barely demonstrated due to an inordinately high catalyst preheating time of a heater, which led to the destruction of [...] Read more.
Although various studies on nitrous oxide as a prospective green propellant have been recently explored, a polyethylene nitrous oxide catalytic decomposition hybrid thruster was barely demonstrated due to an inordinately high catalyst preheating time of a heater, which led to the destruction of components. Therefore, hydrogen peroxide was used as a preheatant, a substance to preheat, with a dual-catalyst bed. The thruster with polyethylene (PE) as a fuel, N2O as an oxidizer, H2O2 as the preheatant, Ru/Al2O3 as a catalyst for the oxidizer, and Pt/Al2O3 as a catalyst for the preheatant was arranged. A preheatant supply time of 10 s with a maximum catalyst bed temperature of more than 500 °C and without combustion and an oxidizer supply time of 20 s with a burning time of approximately 15 s were decided. Because the catalyst bed upstream part for decomposing the preheatant was far from the post-combustion chamber, the post-combustion chamber pressure increased and the preheatant mass flow rate decreased after a hard start during the preheatant supply time. Moreover, because the catalyst bed upstream part primarily contributed to preheating, the maximum catalyst bed temperature was less than the decomposition temperature of the preheatant during the preheatant supply time. Additionally, because the catalyst bed downstream part for decomposing the oxidizer was far from the post-combustion chamber, the post-combustion chamber pressure decreased and then increased during a transient state in the oxidizer supply time. Full article
(This article belongs to the Special Issue Green Propellants for In-Space Propulsion)
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24 pages, 22757 KiB  
Article
Combustion Visualization and Liquid Jet in Crossflow Analysis of H2O2/Kerosene Bipropellant Thruster
by Suk Min Choi, Sangwoo Jung, Vincent Mario Pierre Ugolini and Sejin Kwon
Aerospace 2025, 12(2), 110; https://doi.org/10.3390/aerospace12020110 - 31 Jan 2025
Viewed by 668
Abstract
In the H2O2/Kerosene bipropellant thruster, a liquid fuel jet is transversely injected into a crossflow of hot oxygen and water vapor, catalytically decomposed from a liquid oxidizer. Due to the high temperature and oxygen-rich environment, kerosene is auto-ignited without [...] Read more.
In the H2O2/Kerosene bipropellant thruster, a liquid fuel jet is transversely injected into a crossflow of hot oxygen and water vapor, catalytically decomposed from a liquid oxidizer. Due to the high temperature and oxygen-rich environment, kerosene is auto-ignited without the need for an additional ignition source. Hence, fuel trajectory and breakup processes play a significant role in determining the performance of the rocket engine. However, little effort has been made to analyze these characteristics during actual rocket engine operation, mainly due to its harsh operating conditions of high temperature and pressure. In this study, an optically accessible combustion chamber was prepared to visualize the trajectory and breakup processes of the liquid jet during rocket engine operation. Physical and chemical processes inside the chamber were recorded using a high-speed camera utilizing a shadowgraph technique along with chemiluminescence suppression. Hot-fire tests were performed using 90 wt.% hydrogen peroxide and Jet A-1 in various jet-to-crossflow momentum flux ratios. Test cases with water injection replacing fuel were conducted with varying momentum flux ratios to identify the effect of the combustion process on the liquid jet. The study revealed that the existing correlations for the liquid jet trajectory commonly used for designing the H2O2/Kerosene bipropellant thruster in the past induced significant errors and suggested that the radiation heat transfer from the combustion flame downstream could affect the breakup processes upstream. A new correlation was suggested that accurately predicts the liquid fuel jet trajectory of the H2O2/Kerosene bipropellant thruster. Full article
(This article belongs to the Special Issue Green Propellants for In-Space Propulsion)
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16 pages, 690 KiB  
Article
Characterization of an Ignition System for Nitromethane-Based Monopropellants
by Maxim Kurilov, Christoph U. Kirchberger and Stefan Schlechtriem
Aerospace 2024, 11(12), 1001; https://doi.org/10.3390/aerospace11121001 - 3 Dec 2024
Viewed by 4244
Abstract
This paper presents the results of a hot-fire test campaign aimed at characterizing a newly developed ignition system for nitromethane-based green monopropellants. Nitromethane-based propellants are a cost-effective replacement for hydrazine and energetic ionic liquid hydrazine alternatives such as LMP-103S and ASCENT. The developed [...] Read more.
This paper presents the results of a hot-fire test campaign aimed at characterizing a newly developed ignition system for nitromethane-based green monopropellants. Nitromethane-based propellants are a cost-effective replacement for hydrazine and energetic ionic liquid hydrazine alternatives such as LMP-103S and ASCENT. The developed system uses a glow plug as the ignition source. Additionally, gaseous oxygen is injected simultaneously into the combustion chamber at the beginning of a firing. After closing the oxygen valve, a pure monopropellant operation follows. Three test series were carried out using NMP-001, a previously characterized nitromethane-based monopropellant. During the first test series, the required ROF for ignition was identified to be above 0.3. In the second test series, the low-pressure combustion limit was shown to be 13.9 bar, which is significantly lower than the 30 bar limit of heritage nitromethane-based monopropellants. In the third test series, the oxygen injection timing was optimized to minimize the required amount of oxygen for successful ignition to 1.5 g per ignition in this test setup. This approach to ignition is more cost effective than the catalytic initiation used for other monopropellants because neither costly precious-metal catalytic materials nor lengthy preheating procedures are required. Full article
(This article belongs to the Special Issue Green Propellants for In-Space Propulsion)
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