Development of Green Bipropellant Thrusters and Engines Using 98% Hydrogen Peroxide as Oxidizer

Abstract

The need for non-toxic chemical propulsion systems is growing stronger in today’s space sector. One of the possible solutions for next-generation bipropellant systems is using hydrogen peroxide as the oxidizer. However, there is limited knowledge about using 98% High-Test Peroxide (HTP), which can enable high mass and volumetric performance. Therefore, this paper presents an overview of the development of green bipropellant technology using 98% HTP. The goal is to cover nearly 15 years of experience with 98% HTP and over 10 years of the use of bipropellants containing 98% HTP. The development approach and methods, including component testing and hot-firing, are described. This paper provides test data for various types of bipropellant thrusters and engines producing between 20 and 7000 N of thrust in vacuum, which is the range typically utilized for in-space propulsion. Fuel ignition processes via utilization of a catalyst bed and via hypergolic ignition are analyzed. Successful demonstrations under different operating requirements (steady state, pulse-mode operations, throttleability, etc.) are discussed. The obtained results show that green bipropellants could compete with traditional storable bipropellant technologies. The challenges and opportunities associated with using HTP bipropellants in complete propulsion systems are listed. This paper concludes with recommendations for further research.

1. Introduction

The rapid development of the space sector and the ongoing newspace transformation [1] allow for a more dynamic implementation of new technologies, whose main advantage is potentially lower costs. This is an opportunity for new concepts providing more environmentally friendly solutions, including concepts using green propulsion. The growing costs of handling toxic propellants (and the growing awareness regarding the full lifecycle of materials and propellants via the Lifecycle Assessment approach [2]) have led to wider consideration of alternatives to hydrazine, its derivatives and mixed oxides of nitrogen for in-space propulsion, which have been used extensively for decades due to their performance, reliability and long-term storability. While advances in the use of monopropellants can be seen [3], it is bipropellants that are used in systems requiring high performance. The decrease in costs to launch payloads into orbit during recent years has led to increased attention towards new bipropellant combinations as alternatives to MON/MMH, even in the case where a lower specific impulse is present. However, with lower-performance fuels like ethanol, 98% HTP-based bipropellants have higher-density-specific impulse than toxic propellant combinations. This is due to the high density of HTP (1431 kg/m³ in the case of 98% HTP at 25 °C [4]), and its oxidizer-to-fuel ratio is a few times higher than that of bipropellants using NTO or MON as the oxidizer. This allows the propulsion system size and dry mass to be minimized. It is notable that when using 98% HTP, some propellant combinations can match the performance of MON/MMH, both in terms of density-specific impulse and specific impulse [5], as shown in

Table 1. Comparison of theoretical maximum performance of toxic storable bipropellants with green HTP-based solutions.

Oxidizer	Fuel	Vacuum Specific Impulse [s]	Vacuum Density-Specific Impulse [s kg/dm3]	Oxidizer-to-Fuel Ratio
MON-3	MMH	348	412	1.65 1
NTO	UDMH	344	423	2.00 2
98% HTP 3	96% C2H5OH 4	336	441	4.20
98% HTP	TMPDA	348	467	6.00

¹ The most commonly utilized oxidizer-to-fuel ratio allows for the use of the same oxidizer and fuel tank size; specific impulse maximization only occurs with a rarely used OFR of 2.5, which gives a vacuum specific impulse of 360 s and a density-specific impulse of 443 sˑkg/dm³. ² The oxidizer-to-fuel ratio commonly used in the case of NTO/UDMH systems; specific impulse maximization occurs for an OFR of 3.0, which gives a vacuum specific impulse of 359 s and a density-specific impulse of 461 sˑkg/dm³. ³ The mass-wise concentration of HTP in a hydrogen peroxide–water solution; all HTP concentrations in this paper are based on mass. ⁴ Commercially procured ethanol with a concentration of 96% is actually a concentration of 94% based on mass (the 96% label refers to the volumetric concentration); this was taken into consideration in the relevant thermochemical calculations.

. For a chamber pressure of 10 bar, typical for chemical thrusters for in-space applications, and a nozzle expansion ratio of 200, the maximum theoretical (equilibrium) performance (no losses assumed) was obtained using thermo-equilibrium software [6]. The theoretical performance of MON/MMH for the typically utilized 1.65 oxidizer-to-fuel ratio (OFR) could be matched using 98% HTP and TMPDA if an engine operating at an OFR of 6.0 were to be developed. Moreover, the density-specific impulse of such a formulation would exceed that of MON/MMH by over 13%, potentially leading to additional system-level gains. Using 98% HTP with ethanol, the specific impulse is approximately 3% lower than that of MON-3/MMH (at OFR of 1.65), but it exceeds MON-3/MMH’s volumetric specific impulse by 7%.

1.1. Non-Cryogenic Oxidizers: Hydrogen Peroxide vs. Nitrous Oxide

HTP is a strong oxidizer, second to liquid oxygen, which is cryogenic [7]; thus, it is the non-cryogenic oxidizer with the highest performance with flight heritage. Two non-cryogenic oxidizers are most commonly used as alternative bipropellant combinations for in-space applications: hydrogen peroxide and nitrous oxide. Both can also be used as monopropellants. Both have numerous advantages and disadvantages, which are described in detail in [8]. Several research studies have suggested that the years-long in-space storage of these propellants is possible, despite historical concerns.

A comparison of HTP to other oxidizers (both liquid and solid) is presented in [9]. For in-space bipropellant applications, HTP has been compared to N₂O—both technologies have strong advocates (i.e., [10] for nitrous oxide and [11] for HTP). Comparing bipropellants using HTP with propellants using nitrous oxide as the oxidizer is more complex due to the different storage and feeding pressures in the case of N₂O applications. The propulsion system architecture is different for self-pressurized nitrous oxide compared with classic pressure-fed systems. However, if one considers only specific impulse and density-specific impulse values under the same conditions, swift comparisons can be made. Ref. [12] compared 100% HTP and N₂O, both used with ethanol and assuming a chamber pressure of 10 bar, a nozzle area ratio of 50 and an initial propellant temperature of 293.15 K. HTP showed a 6% improvement in I_sp (specific impulse) and a 82% improvement in density I_sp. This is due to the low density of N₂O, which is 787 kg/m3 at 20 °C when kept in the liquid state [8]. Current development efforts most commonly consider 87.5–98% HTP concentrations. Detailed comparisons of the performance of 87.5% HTP, 98% HTP and nitrous oxide with various fuels is provided in [8]. The 87.5% HTP grade gives better performance than N₂O [8], and every 1% increase in HTP concentration results in roughly a 1% theoretical I_sp gain in monopropellant and a 0.5% gain in bipropellant applications [11]. Therefore, 98% HTP is preferred over the 87.5% grade for applications requiring high performance. Several recent projects funded by the European Space Agency (ESA) have 98% concentration HTP as baseline for new HTP-based bipropellant systems [13]. Modern HTP concentration and purification technologies allow for increased safety of high HTP grades, as it is known that the grade of HTP effects its storability. Ref. [14] present arguments supporting the fact that a high concentration can allow for good storability, often beyond the storability of lower grades. Despite some concerns regarding highly concentrated-HTP applications [15], 98% HTP has been included as “Type 98” in the relevant MIL standard [16].

1.2. Use of 98% HTP and Other HTP Concentrations for Rocket and Spacecraft Propulsion

Heritage HTP bipropellant components and systems (mostly in Germany, the United Kingdom, the United States and the Soviet Union) most commonly used 80–90% HTP [17,18], and the use of 98% HTP has been historically limited. However, already as of 1955, Becco in the United States produced several tons of 99.7% HTP per year [19]. As for use in rocket propulsion, ref. [20] mentions early tests in the 1960s of catalyst beds with 98% HTP. In 1965, a manual regarding HTP with extensive experimental data regarding 98% HTP was published by Food Machinery & Chemical Corporation (FMC) [21], where over 30 years of 98+% HTP production, handling and transportation took place [17]. Throughout the following decades, 98% HTP was tested for rocket propulsion applications by, e.g., General Kinetics and Rockedyne [17]. In 2000, 98% HTP bipropellant technology was considered by NASA to be one of the most promising for 21st-century non-cryogenic propulsion [22]. In Europe, preliminary tests of 98% HTP in bipropellant engines took place in 2014 in Poland with Jet-A as fuel [23]; and ref. [24] mentions tests on 98% HTP with propyne in 2016 in Italy. Work on the utilization of 98% HTP in hybrid rocket motors has also been carried out—mainly in Poland and China, as described in [25]. In recent years, 98% HTP with ethanol as fuel has been selected by Ariane Group for the engine of the Lunanova kick stage, which could be present in future versions of the Ariane 6 launch vehicle [26]. Various R&D efforts in the 21st century report considerations on using 98% HTP but provide experimental data only regarding lower HTP concentrations. Modern bipropellant thruster and engine R&D using lower HTP concentrations includes work in, e.g., Austria [27], Germany [28,29], the United States (including in detonative rocket engines [30]), China [31], South Korea (up to 95%) [32,33] and Norway (87.5%). A Norwegian company, Nammo, uses heritage technology originating from torpedoes, where robust silver catalyst bed technology is utilized [34]. Silver bed technology is, however, effective for HTP concentrations only up to 92% due to the melting point of silver [11]. Work with high grades of HTP has been, until recently, limited due to its availability and cost, as well as due to some concerns regarding the safety and handling of 98% HTP. Historically, several concerns regarding the use of HTP have been presented [15], which may have had the consequence of discouraging the attention of the international rocket propulsion community. However, after some years of limited interest, work with HTP has been growing since the beginning of the 21st century. Nowadays, HTP is becoming increasingly popular among newspace companies. Already as of 2022, Okninski [5] estimated a total of 41 entities developing HTP-based propulsion systems (not only bipropellant). This includes suborbital and orbital launch vehicles, in-orbit systems and exploration vehicles. Over 55% of them were for in-orbit propulsion, but some from the launch vehicle category could also be counted as in-space propulsion, in the case of kick stages, which are to be fired after reaching an initial orbit. Benchmark Space and Rocketlab have successfully fired their HTP-based bipropellant engines in orbit. In 2024 Italian Avio announced the first firing of its Multi-Purpose Green Engine for future Vega launch vehicle evolutions and in-space transportation [35].

1.3. Bipropellant Thrusters and Engines Using HTP—Classification by Ignition Method

The following types of liquid bipropellant thrusters and engines using HTP as oxidizer can be distinguished:

• Quasi-hypergolic, where a catalyst bed is used for HTP decomposition and the fuel is injected into the hot gaseous decomposition products of HTP, allowing for ignition without an additional heat source, as described in [23]. This method has been the one most widely utilized in heritage bipropellant systems using HTP. Some references claim that the implementation of a catalyst bed does not decrease system performance because it does not lead to larger engines or increased mass [36].
• Liquid–liquid injection, where HTP and fuel are injected into the combustion chamber as liquids and require an additional ignition system to enable engine firing. An example of this is provided in Ariane Group’s work within ESA FLPP [37,38]. No information on high-TRL solutions using this technology could be found in the literature.
• Hypergolic, where the fuel ignites in the combustion chamber upon contact with liquid HTP, allowing for reliable, repeatable ignition without an ignitor: thruster/engine firing is thus controlled via the valve opening sequence. The fuels can be made hypergolic with HTP through different methods (catalysts, energetic additives, etc.) [39].

Types 1 and 3 of bipropellant thrusters and engines are discussed within this paper.

1.4. The Origins of R&D with 98% HTP at Lukasiewicz ILOT

Close to 15 years of work with hydrogen peroxide and over 10 years of experimental research on bipropellant systems using 98% HTP as the oxidizer make Lukasiewicz ILOT one of the leaders in HTP-based propulsion research. Work on HTP started in a similar period as in Norway [40], but 98% HTP has been used from start, instead of 87.5%. In 2009 hydrogen peroxide was considered at Lukasiewicz ILOT for the first projects regarding hybrid rocket motors [41]. Experimental work with 98% HTP started in 2011, when the first hybrid rocket motors were tested under laboratory conditions [25]. Wide-ranging research on 98% HTP was possible only due to the indigenously developed HTP concentration and purification technology (in 2011) based on vacuum fractional distillation [42,43,44,45]. First steps in international cooperation regarding HTP-based propulsion were performed within the GRASP (“Green Advanced Space Propulsion”) project conducted within the EC FP7 program. This international project, after considering over 100 propellant candidates, concluded that HTP (next to ADN-based propellants) is the most promising for monopropellants and that HTP is the most promising oxidizer for bipropellant applications [46]. Work on catalysts for 98% HTP by Rarata and Surmacz followed, within the financial scheme for ESA Cooperating States (Plan for European Cooperating States, PECS) with a focus on decomposition of HTP using Al₂O₃ as a support and manganese oxides as the active phase [47,48,49,50]. Poland joined the ESA as a Member State in November 2012, and new opportunities followed. The first bipropellant activities at Lukasiewicz ILOT were funded internally, using quasi-hypergolic ignition via the utilization of a catalyst bed. They started in late 2013 with initial bipropellant firing tests in 2014 [23]. Meanwhile, first considerations of hypergolic rocket engines using HTP took place within the PulCheR (“Pulsed Chemical Rocket with Green High Performance Propellants”) project within EC FP7. Further work continued within ESA-funded activities, with the GRACE (“Green Bi-propellant Apogee Rocket Engine for Future Spacecraft”) project initiated in 2015 being one of the very first Polish ESA projects to be launched. It focused on a bipropellant design utilizing a catalyst bed. Since then, several activities with 98% HTP have followed. An overview of projects that have been carried out with a focus on use of 98% HTP can be found in Refs. [13,51]. Key advancements are presented in this paper.

1.5. Potential of 98% HTP as Oxidizer

The importance of the topic presented in this paper is a result of the ongoing search for non-toxic propellants for in-space applications. Due to miniaturization (somewhat slowed down by the recent decrease in launch costs; this, however, allows for less stringent performance requirements), small satellites are capable of carrying out more complex missions; in addition, bipropellant systems are now also considered for smaller platforms. Despite the growth of electric spacecraft propulsion [52], bipropellant technology remains important due to of the high number of satellite missions where orbital maneuver duration and thrust level play a critical role. Moreover, space debris mitigation requirements result in the need for end-of-life disposal. Short-duration deorbiting maneuvers require the use of chemical propulsion, and bipropellant technology is one of the options that can be considered. The main aim of this paper is to present to the space propulsion community recent advances in projects, mainly funded via the European Space Agency, regarding the utilization of 98% HTP as the oxidizer in bipropellant thrusters and engines. This paper demonstrates the transition from early proof-of-concept tests to maturation and product development for flight applications. It is shown that 98% HTP bipropellants are promising for spacecraft, launch vehicles and in-orbit space transportation in general.

2. Materials and Methods

2.1. Development Approach and Methods

Materials and methods relevant to green bipropellant technology development are presented in

Table 2. Shortlist of materials and methods dedicated to the R&D of bipropellant thrusters and engines using 98% HTP as the oxidizer.

Development Task	Approach and Way Forward Defined for Further Development
98% HTP manufacturing	Use of vacuum fractional distillation to increase concentration of commercially available hydrogen peroxide grades; in-house capability of obtaining 98% HTP and up to 99.99% HTP for small-scale research [13]. Since early successful technology development [42], several advancements have been made, allowing for improvement in the purity of HTP, not published for proprietary reasons. In case where larger volumes of oxidizer are required, there is the possibility to procure HTP from an industrial partner who uses Lukasiewicz ILOT’s concentration and purification technology.
98% HTP stabilization	HTP stabilization is performed using additives, i.e., sodium stannate and sodium nitrate. The resulting amounts of tin and the NO3 anion are in line with the MIL-PRF-16005F standard regarding HTP. Ref. [42] explains that sodium stannate hydrolyzes in HTP, which leads to the formation of colloidal-hydrated stannic oxide. The adsorption of catalytic metallic cations by the colloid ensures stabilization. The stabilization approach used does not hinder reaching near-adiabatic temperatures for 98% HTP decomposition (over 1200 K for 98% HTP), and no issues with catalyst poisoning due to the stabilizers used are experienced.
98% HTP handling	Work with HTP in line with its MSDS (Material Safety Data Sheet), using similar measures to those for lower HTP grades—keeping it away from fuels and organics and ensuring adequate cleanliness, as well as regular training for engineering and technical workers handling HTP. Several guidelines can be found in [44,53], which are in line with state-of-the art HTP practices [54,55,56].
98% HTP storability	Work on purity of 98% HTP (beyond the MIL standard [16]), with parallel efforts on optimization of material selection and surface preparation, passivation techniques and possible advances in HTP stabilization.
98% HTP decomposition	Catalyst technology selection depends on detailed application-driven requirements (lifetime, bed loading, pulse-mode operation performance, cold-stating capability, etc.) and is linked to HTP grade (concentration, purity and stabilization). Low-cost manganese oxide on ceramic pellets has been successfully employed [48,49,57,58,59,60,61], and more advanced solutions using metallic foam [61,62,63] and AM metallic support have been tested [51]. Some have successfully demonstrated cold-start capability, as shown in [61,62], without failing despite over 30 repetitions of firing initiated under cold (non-pre-heated) conditions. Pulse-mode operations were successfully demonstrated for monopropellant thrusters [61,62]. Novel catalysts using modified silver capable of operation with 98% HTP were also introduced [61]. Details of the highest-performing catalyst beds have not been published, as it is an important element covered by Intellectual Property rights. For the newest solution, all manufacturing steps (both support material and active material) are performed in-house. Early work on other methods of HTP decomposition has also been performed; this includes thermal 98% HTP decomposition [64] and UV-induced decomposition [65], both studies were in cooperation with the Warsaw University of Technology.
Fuel for quasi-hypergolic thrusters and engines	A down-selection of fuel candidates for use with 98% HTP has been carried out in several European projects. Current development efforts on non-hypergolic engines include the use of TMPDA and ethanol as fuel, where the first one is selected in case of performance-driven requirements. Earlier work included tests with, i.e., Jet-A [66] (also during the very first proof-of-concept tests [23]), turpentine [66] and butanol [67]. For very-high-performance applications (in both quasi- and hypergolic rocket engines), DMAZ [68] was also considered.
Fuel for hypergolic thrusters and engines	Extensive trade-off analyses to identify fuels for hypergolic applications with 98% HTP have been performed [68]. A few fuels have been successfully employed in thruster and engine testing. After multiple literature reviews, an early screening was performed using fuel ignition drop tests. First, simple atmospheric drop tests were performed, which later involved more advanced instrumentation and test setups. These included drop tests under controlled conditions (pressurization of the drop test section with neutral gas within a range of pressures relevant for fuel thruster/engine conditions and control of the temperature of the fuel droplet). Further tests included single-element injector testing with no combustion chamber (and, in some campaigns, using a closed volume for pressurization [69]). Moreover, full injector head testing without a combustion chamber and final validation in breadboard thrusters and engines and ultimately in flight-like thrusters have been performed. Apart from the fuel ignition delay, combustion roughness and performance have been considered. Long-duration in-space storability is a critical driver regarding detailed fuel compositions for orbital missions. The most promising fuel research [70] has not been published due to the criticality of Intellectual Property, but early work in this field including some experimental data can be found in [68,69,71,72].
Design and modeling	The design of the engine mechanical is performed in NX software, with the most advanced projects being conducted in Teamcenter, allowing for Product Lifecycle Management, crucial to larger R&D teams working on higher-TRL projects. Simulations include all state-of-the-art efforts linked to structural and thermal design, essential to flight-hardware development. Several in-house tools are used for advanced heat load estimation, film-cooling modeling and structural optimization for regeneratively cooled combustion chambers. Work has been also performed in the field of acoustics [73]. System-level trade-offs have been supported by tools allowing for propulsion system multidisciplinary optimization, implementing simplified models [74].
Combustion chamber technology	Breadboard tests including battleship combustion chambers that facilitate instrumentation enable short-duration firing (including heat sink chambers made of electrolytic copper). Different solutions are used for flight-like designs: radiation-cooled chambers with film cooling, as well as regeneratively cooled engines. In the case of the former type of chamber, coated titanium zirconium molybdenum (TZM) alloy has been successfully used. Work on coatings for niobium combustion chambers has also been carried out [75,76]. An innovative method of combustion chamber surface temperature distribution investigation via the use of special paint has been studied [77,78]. Some combustion chambers are additively manufactured as a single element, including the regeneratively cooled AM combustion chambers for the throttleable engine described in [79], where copper chromium zirconium (CuCrZr) alloy is used. Further AM technology development with high-temperature metal powders is being considered.
Cooling	Various cooling methods have been used during R&D projects related to green rocket propulsion [13]. Non-cooled chambers, accumulative chamber cooling and regenerative water cooling were employed in several early tests. In further development, for reaching the thermal steady state during the firing of 20 N and 420 N thrusters, radiative combustion chamber cooling with internal film cooling was used (in parallel), which is in line with solutions for state-of-the-art bipropellant engines and thrusters of size up to that of LAEs. For engines developing several kN of thrust, the ultimate cooling technology should be regenerative chamber cooling using the propellant utilized. In this case, initial testing has been performed with instrumented water-cooled chambers allowing for detailed thermal load characterization.
Injector technology	Impinging injectors are used for hypergolic thrusters and engines. As for quasi-hypergolic engines, where liquid fuel ignition occurs in the high-temperature gaseous decomposition products of 98% HTP, early designs considered perpendicular injection of the fuel, as shown in [80]. The results of some trade-offs regarding the number of orifices, injection angle and the effective combustion chamber length are shown in [81]. The evolution of detailed modeling methods supported by experimental data regarding injection of liquid jet into gaseous crossflow can be found in [79,82]. Both hypergolic and quasi-hypergolic engines and thrusters successfully use AM injectors manufactured in-house [79,83]. The first proof-of-concept AM injector design work started in 2014.
Hot-firing testing	Full thruster and engine testing is carried out at several test benches. Most importantly, both sea-level and high-altitude test benches are available for green bipropellants using 98% HTP as the oxidizer. Extensive instrumentation, including fast cameras, infrared cameras, hundreds of control and data acquisition channels, etc., is available [84]. Steady-state vacuum testing is possible for thrusters up to 500 N thanks to the new electric-pump-based facility [85,86] and for those up to 5000 N in case of the sea-level test cell. In addition, a mobile hot-firing facility allowing for testing engines up to 60 kN of thrust is under development [87].
Building flight heritage of 98% HTP	No information about 98% HTP use in orbit could be found as of 2025. In 2017, 98% HTP was tested in suborbital flight for the first time during the ILR-33 AMBER rocket maiden flight [25]. In total, 5 successful launches of AMBER and AMBER 2K variants have been carried out, including reaching the Von Karman line in 2024 [25,88,89,90]. However, these vehicles use 98% HTP in a hybrid rocket motor, not a liquid bipropellant system. As of 2025, there is no information in the literature regarding in-flight use of 98% HTP in bipropellant mode, not even in suborbital missions.
System development	Work on development of full propulsion systems using 98% HTP has been performed for monopropellant systems, including orbital missions [91,92], and is ongoing, currently at a lower TRL, for ISTVs, including both monopropellant and bipropellant thrusters [93]. Fluidic components for 98% HTP for large satellite platforms are also under development, but this does not consider the full system (Liquid Apogee Engine (LAE), valves, etc.); however, analyses concerning system-level aspects have been performed for such platforms [74]. It is well understood that regardless of the platform size and mission, the development of all necessary fluidic components for use of 98% HTP (and the respective fuel) is necessary. Most importantly this includes compatibility and storability issues, which are especially important for multi-year-long in-space operations. Lessons that could aid development of orbital propulsion systems utilizing 98% HTP can also be learned from work on non-spacecraft HTP systems, including suborbital systems, and from launch vehicle development.
Product development, commercialization and industrialization	Successful commercialization of 98% HTP bipropellant thrusters and engines is expected to be possible due to the pressure to replace MON/MMH and other bipropellants based on NTO and hydrazine derivatives, as recommended in the REACH regulation [94]. It is considered very likely that commercial clients, especially within the newspace ecosystem, will be the first to implement bipropellant systems using 98% HTP as the oxidizer. Space propulsion development efforts at the Lukasiewicz Institute of Aviation have identified potential future customer requirements and key topics still to be addressed by remaining development work. To ensure that the technologies under development have commercial potential, the involvement of industry, including potential end-users, is a widely adopted practice. This happens via involvement of end-users in consortia or as requirement-generating partners. In this way, i.e., European Large Satellite Integrators or STS primes are identified for each potential product, and only market-driven products enter development. Electronic Engine/Thruster Control Units are also under development to be able to provide a complete system for customers [95,96]. Several commercialization approaches are possible, including spin-offs, such as Thaliana Space, who will industrialize the GRACE 98% HTP bipropellant engine [97]. Other forms of licensing, joint ventures, etc., are also possible. The strategy regarding technology industrialization assumes that the majority of the value chain must be located in Poland and rely on European technologies, while end-users can be international partners, including from outside Europe.

. Each engineering task is shown with comments regarding the R&D approach and the selected way forward. There is obviously some change in the planned methods and approaches as new knowledge and experience are acquired; however, so far, such updates have not had a large effect on the overall development roadmaps.

2.2. Thrusters and Engines

The following paragraphs focus on the three thrusters and engines under development, which were specified in Table 3. A wider range of products under development, including other thrusters, valves and non-HTP propulsion solutions, can be found in [13,51].

2.2.1. GRACE (Green Liquid Apogee Engine for Future Spacecraft)

Based on promising results of early bipropellant engine testing and significant progress with catalysts, work under an ESA contractual frame was initiated. Lukasiewicz ILOT started a new activity to develop an LAE-class green bipropellant thruster. The baseline concept involved catalytically decomposed 98% HTP as the oxidizer, while the fuel remained the subject of a two-step down-selection process. An extensive trade-off analysis was followed by the hot-firing test campaign, encompassing the three most promising candidates. TMPDA (N,N,N′,N′-Tetramethyl-1,3-propanediamine), assay 99, used by the chemical industry as a building block for synthesis and an additive for polymer production [98], was finally selected as the fuel for the engine. This project, called GRACE, eventually transformed into a wider development program. Initially planned as an LAE, the GRACE engine has evolved into a main propulsion subsystem, designed for kick stages of launch vehicles and servicing spacecraft. While the first phase of the GRACE program was focused on the initial selection of the engine configuration and the demonstration of the concept, the second phase proved the maturity of critical technologies. The performed hot-firing testing supported all the design and analytical activities in terms of the selection of the fuel injector and combustion chamber length. The present configuration of the engine can be seen in Figure 1.

The major technological challenge underlying the selected catalytic concept was an efficient and robust catalyst decomposing 98% hydrogen peroxide. The resulting product, internally developed by Lukasiewicz ILOT during Phase 2 of the program, has proven to meet all the given requirements for the catalyst. The developed active phase combines high catalytic activity with mechanical strength, resistance to thermal parameters (stress and shocks) and tolerance to stabilizers present in the propellant. This active phase can be deposited on carriers in the form of pellets, metallic or ceramic foams, and meshes. Selection methodologies for hydrogen peroxide catalysts were established based on the surface loading and volumetric loading of the catalytic bed. A dedicated packing procedure was developed for the catalyst in the engine. These methods ensure that the catalyst remains in position during operation, preventing bed movement, pressure instabilities and abrasion of the active phase over the lifetime of the catalyst. Beyond the load parameters, an equally important factor is the uniform distribution of propellant across the front surface of the catalyst, which is considered during the design of the hydrogen peroxide injection method.

The other technological gap—a high-temperature- and oxidation-resistant material for the combustion chamber—was also focused on in parallel. This ceramic-coated molybdenum alloy served for the manufacturing of radiatively and film-cooled combustion chambers. The injector of liquid jet into crossflow (gas–liquid type) provides efficient and stable combustion, completing the list of critical components. Liquid fuel is injected into the crossflow of high-temperature catalyst bed decomposition products, consisting of gaseous oxygen and gaseous water.

The identification of the optimum design point of a radiatively and film-cooled thruster required solving a coupled dependency among the given constants and variables. The chamber material service temperature and the chamber pressure acted as constants. Variables included the engine (global) mixture ratio affecting the combustion gas temperature and the amount of the fuel injected into the chamber internal wall (fuel film cooling). The impact of both the global mixture ratio and the percentage of the fuel film flow on combustion efficiency could not be neglected in this activity. The preliminary approach to this challenge involved a separately fed fuel film-cooling injector. The thermal steady-state tests performed in this configuration helped to identify the flow rate of the fuel corresponding to the given throat temperature of the chamber. Moreover, testing different quasi-design points and a test-by-test analysis of the results helped identify the specific operating conditions corresponding to the maximum propulsive performance. The ultimate experimental setup involved one injector manifold allowing for the injection of fuel both for film cooling and into the main combustion zone.

2.2.2. Throttleable Rocket Engines

Throttling has been identified by the ESA as a key technology required both for the reusability of LVs’ first stages and for future lunar and planetary descent elements. The relevant applications require what can be described as “deep-throttling”. Decreasing thrust down to 10% of its nominal value has been identified in Ref. [99] as necessary for the most demanding lunar missions. Moreover, apart from exploration, small LVs with a single engine in their first stage would require deep-throttleability to land safely, enabling reusability (Falcon 9 during its descent maneuvers turns off its LREs, leaving just one out of nine LREs firing at touchdown, thus not requiring deep-throttleability, although throttling to some degree).

Lukasiewicz ILOT developed a throttleable monopropellant system for CNES for its FROG-H project [100]. This system can throttle down to below 10% of nominal thrust by controlling the HTP mass flow. However, to allow for the efficient deep-throttling of bipropellant rocket engines, oxidizer and fuel mass flow regulation is not enough [101]. Apart from varying propellant flows, a variable-geometry injector was proposed. Pintle injectors were used both during a pathfinder ESA activity (launched as of 2018/2019), where stable operation at 20% of nominal thrust was demonstrated (but without dynamic throttling) [67], and during consecutive larger ESA projects with pintle actuation and dynamic throttling demonstrations. Whilst during the pathfinder project butanol was used as fuel with 98% HTP, the follow-up activities focus on ethanol with 98% HTP. The Throttleable Liquid Propulsion Demonstrator (TLPD), realized within the ESA Future Launchers Preparatory Programme (FLPP), enables deep-throttling down to 10% of the design thrust (5 kN at sea level and ~7 kN in vacuum). It uses ethanol at 96% concentration (volumetric), which corresponds to the most commonly offered maximum concentration obtained via distillation. Throttling is realized through two independently actuated cavitating mass flow regulation valves and a variable-geometry fuel injector, as shown in Figure 2. Combustion performance and stability are ensured thanks to the variable-geometry fuel pintle injector, which allows the injection pressure drop to be controlled. Thanks to the catalytic decomposition of the oxidizer, which provides sufficient heat to ignite the mixture smoothly, there is no need for an additional source of heat for ignition.

2.2.3. Hypergolic Thrusters and Rocket Engines

Various applications require a reliable multi-reignition capability and high-performance pulse-mode operations. This includes both Reaction Control Systems (for spacecraft, launch vehicles and missile systems), as well as various other satellite propulsion systems. Moreover, reducing the mass of the propulsion system while increasing its reliability is a key factor in rocket launcher design. Eliminating the ignition system and its associated equipment can significantly decrease the total mass of the propulsion modules.

The most common approach in performance-driven systems is to apply hypergolic propellant technology. Until now, highly toxic and corrosive hydrazine and NTO compounds or their derivatives have been used in the field of rocket engine technology. Nonetheless, over the last two decades, a new approach to the development of hypergolic propellants—commonly referred to as green propellants—has been initiated to reduce their toxicity, corrosiveness and environmental impact. First preliminary considerations in this field [102] were followed by the involvement of the Lukasiewicz Institute of Aviation in the “Pulsed Chemical Rocket with Green High Performance Propellants” (PUCHER) European project [103]. Work using statutory funds followed, allowing for small-scale laboratory demonstrations. The national HIPERGOL project, run thanks to co-funding by the National Research and Development Centre from the budget of the Polish Ministry of National Defense [104], performed the validation of various hypergolic fuels during extensive screenings and ultimately performed a demonstration within a 5000 N (SL) hypergolic rocket engine. Despite successful finalization of the R&D project [105], key project outcomes have not been published as of 2025. Throughout several internal projects, further testing continued using a breadboard 500 N (SL) thruster with a regeneratively water-cooled combustion chamber, shown in Figure 3.

Hypergolic fuel development activities using internal budget continued even after obtaining an ESA contract dedicated to green hypergolic thrust technology development, due to the criticality of the relevant IP and increased interest of commercial players. The most promising results were obtained using the NNP-103B fuel composition. The key ingredients of the proprietary NNP fuel family have not been publicly disclosed and are considered commercially sensitive. All ingredients are not subject to REACH, nor are candidates for the REACH list and have low vapor pressure. All fuels of the NNP family have densities exceeding 850 kg/m³ at 20 °C (enabling both high-density and -volumetric performance). While the NNP fuel family is not presented in Table 1, its most basic composition gives a theoretical 469 sˑkg/dm³ of vacuum density-specific impulse (for assumptions in line with the calculation results presented in Table 1). Some of the more hydrogen-rich compositions allow for an increase in the specific impulse beyond the values listed in Table 1. The fuel’s IDT and viscosity allow for efficient use in hypergolic thrusters. NNP fuel boiling and freezing points are acceptable considering use on board spacecraft. Key ongoing work is focused on tuning the NNP family fuel compositions to allow for long-duration in-space applications. Ultimate fuel characterization must include extensive data on storability of fuel (via accelerated aging tests) and its material compatibility. The hypergolic thruster development presented within this paper concerns the ESA “10–20 N green bipropellant thruster” (NGBT) project, where the NNP-103B fuel was used. The thruster utilized a TZM combustion chamber with coating as in the GRACE program. A nominal inlet pressure of 13 bar was used. The thruster mounted for vertical firing can be seen in Figure 4.

Table 3. Bipropellant rocket thrusters and engines under development at Lukasiewicz ILOT.

Designation	Oxidizer	Fuel	Vacuum Thrust [N] 1	Projected Vacuum Specific Impulse [s] 2	Key Features	Application	References
NGBT	98% HTP	NNP	20	319	Hypergolic	Satellite propulsion, RCS for EDL	[13,63]
GRACE	98% HTP	TMPDA	420	299	Quasi-hypergolic	LAEs for telecom spacecraft for transfer to GEO, In-Space Transportation Vehicle (ISTV) propulsion and kick stage for micro- and small LVs	[57,66,81,83]
TLPD	98% HTP	C2H5OH	7000	300	Deep-throttleable and quasi-hypergolic	Lunar/planetary lander propulsion and kick stage for large LVs or upper-stage for micro-LVs 3	[63,79,106]

¹ Ultimate thrust level may be adapted to some extent to fill market needs. ² Vacuum specific impulse calculated using c* efficiencies obtained during SL hot-firing. NGBT performance assumes a nozzle expansion ratio of 300, while the presented GRACE and TLPD performance assumes a nozzle expansion ratio of 220. ³ In case of most LV applications, the engine does not need to be deep-throttleable (it can use a fixed injector geometry and no cavitating venturi valves).

Table 3. Bipropellant rocket thrusters and engines under development at Lukasiewicz ILOT.

Designation	Oxidizer	Fuel	Vacuum Thrust [N] 1	Projected Vacuum Specific Impulse [s] 2	Key Features	Application	References
NGBT	98% HTP	NNP	20	319	Hypergolic	Satellite propulsion, RCS for EDL	[13,63]
GRACE	98% HTP	TMPDA	420	299	Quasi-hypergolic	LAEs for telecom spacecraft for transfer to GEO, In-Space Transportation Vehicle (ISTV) propulsion and kick stage for micro- and small LVs	[57,66,81,83]
TLPD	98% HTP	C2H5OH	7000	300	Deep-throttleable and quasi-hypergolic	Lunar/planetary lander propulsion and kick stage for large LVs or upper-stage for micro-LVs 3	[63,79,106]

¹ Ultimate thrust level may be adapted to some extent to fill market needs. ² Vacuum specific impulse calculated using c* efficiencies obtained during SL hot-firing. NGBT performance assumes a nozzle expansion ratio of 300, while the presented GRACE and TLPD performance assumes a nozzle expansion ratio of 220. ³ In case of most LV applications, the engine does not need to be deep-throttleable (it can use a fixed injector geometry and no cavitating venturi valves).

3. Results of Work on Bipropellants Using 98% HTP as Oxidizer

3.1. GRACE Rocket Engine Test Results

3.1.1. Pre-Qualification of the Catalyst Bed

Catalyst technology maturation has been recognized as a critical step in the engine development process. The successfully completed monopropellant hot-firing test campaign, involving over 330 kg of 98% HTP, delivered promising results. The catalyst operated correctly throughout this significant propellant mass flow, which was later increased to 480 kg by its re-use in the subsequent, bipropellant test campaign. No indication of the approaching end of life could be observed. The catalyst performance, by means of its efficiency of 98% HTP full decomposition, was maintained within the entire run time. Both pre-heated and cold-start (294 K) [61] characteristics were identified as acceptable in terms of fast and reliable bipropellant ignition. The maximum demonstrated continuous monopropellant test (see Figure 5), limited by the capacity of the feed system (50 kg of 98% HTP), lasted 450 s.

3.1.2. Experimental Bipropellant Investigation Towards the Optimum Design Point

Both monopropellant and bipropellant hot-firing tests were performed under atmospheric conditions in the internal test facility at Lukasiewicz ILOT.

Hot-firing test results confirmed the design estimation of the efficiency of characteristic velocity (η_c* = 93%) at the nominal operating point of the engine. Based on the atmospheric test results and the estimation of nozzle efficiency, the vacuum impulse of the GRACE engine, expected to be reached with the nozzle area ratio of 220 at the mixture ratio of 3.3, is 299 s. This value is far from the maximum theoretical performance cited in Table 1 due to the OFR selected.

The latest test campaign, employing two engines tested sequentially, proved that the thrust and total impulse repeatability (both test-to-test and engine-to-engine repeatability) for the given operating conditions remained at acceptable levels. Discrepancies in both thrust and total impulse remained below 1% of expected values. The roughness of the chamber pressure, characterized by 3σ, remained between 1% and 2% of its mean value, with no evidence of evolution throughout the entire test campaign and with just a minor impact of the operating point. Images from GRACE testing are provided in Figure 6.

Apart from the nominal operating point, the engine experienced a wide qualification pressure envelope, via modification of the inlet pressure. Testing at extreme operating points translated into a chamber pressure in the range of 8 bar to 11 bar and a mixture ratio from 2.75 to 4. The impact on combustion stability resulting from testing at these extremes appeared to be insignificant.

The longest single firing demonstrated so far lasted 180 s. The chamber pressure plot for this test is shown in Figure 7. The maximum duration was defined by the feed system capacity and budgetary constraints of the project. The total single-engine accumulated firing duration exceeded 20 min, with no issues identified after test article investigation.

3.2. Throttleable Rocket Engine Test Results

The TLPD technology enables thrust modulation down to 10% of the nominal operating point (5 kN at sea level). In the recent hot-firing campaign (2024), the engine managed to achieve 10% of thrust with high combustion efficiency, encountering efficiency decrease in the high-thrust region. Thrust up to 100% was demonstrated [79]. Tests prove that the c* efficiency remains high down to 10% of the design point.

Table 4. Combustion chamber pressure roughness during rocket engine throttling [79].

Thrust Level [% of Nominal Level]	Chamber Pressure [barA]	Pressure Roughness 3 Sigma/Avg. [%]	Pressure Roughness Peak to Peak [%]
100	17.0	0.8	5.5
80	14.1	1.1	7.6
50	9.6	0.7	5.4
30	6.6	1.1	7.2
20	5.0	1.4	11.3
17	4.6	2.6	17.0
15	4.2	3.9	19.2
10	3.3	13.1	45.4

presents thrust stability in terms of the pressure roughness versus throttling point. Since the roughness is the highest for the lowest-thrust operating points, more data for operating points at 10% and 15% of nominal thrust are provided. Figure 8 presents dynamic chamber pressure measurements at 10% of nominal thrust, and Figure 9 provides results of the Fast Fourier Transform (FFT) of chamber pressure data at 10% of nominal thrust. Similarly, Figure 10 presents dynamic chamber pressure measurements at 15% of nominal thrust, and Figure 11 gives results of the FFT of chamber pressure at 15% of nominal thrust. While the chamber pressure roughness is larger in the case of the 10% thrust point than in case of the 15% thrust point, a more uniform roughness distribution can be observed for the 10% thrust point. It can be seen that the 40 Hz frequency dominates in both FFT analyses.

An image from a bipropellant firing test is shown in Figure 7. Notably, the outflow is significantly cooled, due to the oversized regenerative cooling, with water condensation visible along the nozzle edge in Figure 12.

The engine was tested with open-loop control as well in closed-loop mode with thrust/pressure setpoint varying during the test. Example results are presented in [79] along with the model prediction, using a simulation tool developed in-house. The ignition robustness, operation stability, throttling in both modes and reignition capabilities were successfully demonstrated. Figure 13 presents the results of selected bipropellant tests, with a comparison to simulations, where the test results along with the command signal are visible in the plot. A wider description regarding the engine control logic is available in [107]. The system enabled thrust variation at a rate of 2–4 kN/s depending on the defined thrust profile. Further results can be found in [79].

3.3. Hypergolic Thruster and Engine Test Results

As part of the HIPERGOL project, new technologies were developed and implemented to support the design and testing of hypergolic propulsion systems under development at the Lukasiewicz Institute of Aviation. Work on the 5 kN regeneratively cooled combustion chamber initiated a new approach to designing and manufacturing using 3D copper printing at Lukasiewicz ILOT. A novel hypergolic propellant formulation, consisting of 98% HTP and a pyridine-based fuel blend, was evaluated among other propellant solutions. The reliable and rapid ignition of the developed hypergolic propellant based on 98% HTP and the pyridine blend was demonstrated, with an ignition delay time (IDT) below 3 ms under drop test conditions and below 10 ms during full rocket engine injector operation. Video stills from tests of the ignition sequence, without the mounting of the combustion chamber, are visible in Figure 14. Finally, the 5000 N demonstrator was successfully tested. The instrumented regeneratively water-cooled combustion chamber and a still from engine hot-firing are shown in Figure 15. A cumulative hot-firing time of 57 s was achieved with no failure. The longest single test lasted 10 s, allowing the combustion chamber to reach steady-state thermal conditions (Figure 16).

As for the 20 N hypergolic thruster (NGBT), tested with the NNP fuel in a dedicated ESA project, a wide hot-firing campaign of the test item was executed. The IDT of the hypergolic fuel was around 5 ms during drop tests, injector firing tests (Figure 17) and full thruster hot-firing tests (Figure 18). Both pulse-mode and steady-state operation were verified. Single-burn durations between 20 ms and 40 s were performed. The successfully delivered minimum impulse bit was 2 Ns. Pulse-mode performance characteristics are given in Figure 19. In total, over 500 thruster firing tests were carried out with a cumulative duration of above 10 min. The thruster delivered a high c* efficiency of 0.98, which corresponds to a projected vacuum specific impulse of 319 s when assuming realistic efficiencies of the high-altitude nozzle. A pressure roughness of 1.1% was measured, and an example pressure plot from one of the firing tests is provided in Figure 20 [51].

4. Discussion

All three of the presented bipropellant thrusters and engines using 98% HTP as oxidizer provided promising results regarding further development and potential flight applications. Since each of them has different requirements due to very different applications (steady-state operation vs. pulse-mode capability vs. throttleability), the results presented are discussed separately for each of the development efforts.

4.1. GRACE

Results of the catalyst bed pre-qualification and bipropellant steady-state operation have been shown. While the 480 kg 98% HTP test of the catalyst bed and verifying the bipropellant performance throughout its operation envelope were critical milestones, the total bipropellant burn duration did not exceed durations ultimately required for most applications. The most demanding application in terms of total operation time is serving as an LAE for GEO telecom platforms, where the required burn duration would be substantially lower for use in a kick stage or an ISTV. However, it is expected that the longer engine operation will not be an issue, which is in line with post-test hardware inspection results.

Future applications include use of the engine onboard of the IOSHEX platform, which is to serve as an advanced satellite dispenser and ISTV allowing for spacecraft servicing, etc. More data on IOSHEX with a focus on trade-offs regarding its propulsion system can be found in [93], where a vehicle configuration of circa 1 ton of wet mass, including 220 kg of propellants, is discussed. GRACE was also preliminarily considered for upper-stage propulsion for micro-launchers during feasibility studies carried out at the Lukasiewicz Institute of Aviation. In 2025, MaiaSpace selected the 98% bipropellant rocket engine technology based on GRACE for the kick stage of its small, partly reusable launch vehicle [97], giving the opportunity for flight application within a major European Space Transportation System, currently under development.

4.2. Throttleable Rocket Engine

Dynamic throttling, including preliminary tests with closed-loop control throttling, has been performed. Decreasing thrust down to 10% of its nominal level was demonstrated, and decreases down to 20% were shown to be very good in terms of acceptable levels of combustion chamber pressure roughness. High combustion efficiency for deep-throttling was obtained. Subsequent optimization shall focus on increasing the combustion efficiency near the nominal operating point (94.5% efficiency at 30% of nominal thrust was assumed satisfactory).

The completed throttleable engine development phases were performed for engines with a thrust level corresponding to the European ARGONAUT mission’s lander rocket engine [108] and to future European lunar lander architectures. However, unlike in ARGONAUT, deep-throttleability was a key requirement. Beyond serving as a building-block for VTVL demonstration and for lunar and planetary landers, the robust 98% HTP + ethanol technology is interesting for launch vehicle upper stages and kick stages. Naturally, a non-throttleable configuration with a fixed pintle geometry, no actuators and cavitating venturi valves would be used in such a case.

Further work is needed on regenerative cooling with 98% HTP, since historical European heritage and literature data are limited to 90% HTP applications. A promising but currently unexplored topic is throttling using hypergolic propellants.

4.3. Hypergolic Thrusters and Engines

The presented results show a novel perspective for green bipropellant technology in high-performance spacecraft propulsion, including ISTVs. Due to its capability regarding pulse-mode operation, it can serve as the RCS of entry-descent elements and launch vehicles. The estimated vacuum specific impulse of 319 s makes it attractive for upper and kick stages of launch vehicles and for exploration missions if multi-year-long in-space storability is enabled.

However, complete fuel characterization is needed. The NNP fuel compositions must be optimized for in-space storability. Further work, apart from propellant maturation, will focus on the qualification of the specific processes for thruster manufacture.

4.4. General Discussion

Numerous current development efforts, also for commercial application, still focus on lower concentration grades of HTP, but the presented results may give further arguments that work on 98% HTP is promising and may be, in many cases, the best solution for in-space environmentally friendly propulsion. In particular it is promising for LEO but may also be advantageous even for more demanding destinations including GEO and lunar missions [2]. HTP may also be of benefit in new mission scenarios (satellite servicing, orbital propellant loading, Active Debris Removal, etc.).

Mastering cooling technology for HTP-based bipropellant thrusters and engines is essential to their further performance optimization. In case of thrust levels below 500 N, film cooling with a radiation-cooled combustion chamber are natural technologies to be applied. In case of heritage thrusters and engines using NTO or MON as the oxidizer and hydrazine derivatives as fuel, the typical OFR is between 1.6 and 2.7. For HTP-based bipropellants, the OFR, depending on the fuel, may be between 3.0 and 8.0. Thus, little fuel is available for film cooling. This may lead to suboptimal performance when using fuel film cooling. Therefore, the oxidizer should also be considered for film cooling. This requires significant know-how regarding construction material interaction with HTP and on oxidizer thermal decomposition. Regardless of whether it is the fuel or the oxidizer performing the film cooling, the fraction of mass flow directed for film cooling must be minimized to allow for propulsive performance maximization. As for challenges of the cooling of engines of several kN of thrust, regenerative cooling is to be employed. The high OFR of HTP-based bipropellants leaves a low mass flow of fuel to be used for regenerative cooling; thus, cooling with HTP should be considered. While this has been successfully demonstrated historically, new research on cooling with modern HTP grades (including 98% HTP) is needed.

Whilst there has been much scientific debate on the maximum length of missions using HTP, new technologies may lead to new capabilities. Concerns regarding the use of high grades of HTP may become partly obsolete due to new advances not only with the oxidizer itself but also on feeding systems and fluidic components optimized for use with HTP. Of course, necessary precautions and adequate handling procedures must always be implemented, as with any liquid rocket propulsion technology. While a leak of HTP [109] has increased SpaceX requirement’s for hosting spacecraft using HTP on its Falcon 9 launch vehicle, heavy investments in Italy use of lower grades of HTP in orbital missions by Benchmark Space and RocketLab, and major interest of Newspace players give promise for wider HTP bipropellant application. The market outlook shows potential for bipropellant rocket engines using 98% HTP producing between 1 N and 30 000 N of thrust. Some companies consider HTP to be an oxidizer for the main stages of their orbital launch vehicles (as bi- [110] and hybrid [111] propellants), which has been historically shown to be possible (with the British Gamma rocket engines from the 1950s to the early 1970s [18] and attempted on a very large scale in the case of the 3.6 MN Beal Aerospace engine tested in the United States in 2000 [112]).

Further work must now allow for gradual TRL increase and ultimately qualification for orbital use of 98% HTP bipropellants. So far, use of 98% HTP was implemented in the hybrid rocket motors of the ILR-33 AMBER 2K suborbital rocket, which reached the Von Karman Line [88], and in several ongoing spacecraft projects at flight qualification level but limited to monopropellant systems (e.g., in mid-2025, Arkadia Space announced the successful use of an HTP monopropellant system in its orbital mission [113]).

Several of the potential challenges regarding wider incorporation of bipropellant thrusters and engines using 98% HTP as the oxidizer are being tackled thanks to the ESA Innovative Propulsion Cross Cutting Initiative. Specific key topics are to be addressed within new ESA projects which form the hydrogen peroxide roadmap, which has been prepared by the Agency taking into account inputs from key industry and R&D entities working on green propulsion [114].

Specific needs for 98% HTP bipropellant advancement are linked to system-level aspects and availability of commercially off-the-shelf fluidic components. Further work includes wider implementation of hypergolic engines, which allow for reliable high-performance pulse-mode operation.

5. Conclusions

Increasing interest regarding the use of HTP and in particular 98% HTP can be seen. This pertains to research and technology implementation for near-term and next-generation missions. The presented advancements show several promising technologies linked to 98% HTP. Taking into account the discussed development efforts, the following conclusions can be made:

• Some “world’s first” achievements (assuming that data are available in the literature) have been made, including

  ○

  Successful tests of an LAE thrust-class bipropellant rocket engine using 98% HTP as the oxidizer (GRACE).

  ○

  Successful tests on a larger (multi-kN) rocket engine using non-toxic liquid hypergolic propellants with 98% HTP as the oxidizer (HIPERGOL).

  ○

  Successful use of a non-toxic hypergolic propellant combination with expected vacuum performance beyond 310 s (based on measured c* efficiencies during SL firing) (NGBT).

  ○

  The successful deep-throttling of a 98% HTP bipropellant engine (TLPD).

• No showstoppers regarding further engine development have been identified so far, despite further research and development work being needed. The obtained combustion stability and pressure roughness are in line with requirements for bipropellants using catalyst beds as well as hypergolic thrusters, although some optimization may be required regarding pressure roughness at deep-throttling operating points (low thrust). Both types of HTP-based bipropulsion systems (using catalyst beds and hypergolics) have been proven to allow for reliable reignition.
• Some non-typical solutions have been verified under laboratory conditions, e.g., the adaptable pintle injector, the TZM combustion chamber with coating, additive-manufactured injector heads and combustion chambers, and novel fuels and catalyst beds (modified silver and AM support, metallic foams, etc.).
• The obtained and expected vacuum performance for 98% HTP with TMPDA and ethanol was satisfactory but may not be enough for the most demanding missions. However, apart from engine c* efficiency optimization and operation at closer-to-optimal OFRs, other solutions (such as the high-performance NNP hypergolic fuels or, e.g., DMAZ) have been also identified.
• Naturally some potential limitations must be mentioned. They are not considered showstoppers but require continued research efforts. They include making sure that the catalyst lifetime matches the requirements for, e.g., telecom platforms, where the propellant throughput is of a few tons, beyond the 480 kg of 98% HTP tested with the GRACE engine. This may pose challenges for the catalyst bed, although no signs of failure modes have been identified in the tests undertaken so far. For hypergolic thrusters and engines, fuel storability must be ensured, and more research in this field is of critical importance. Lastly, while HTP storability has been historically shown to be possible, more work is needed, as mentioned in conclusion number 6. The vast majority of the challenges mentioned will be addressed by the European space sector within activities listed in the ESA Innovative Propulsion Cross Cutting Initiative roadmap regarding HTP utilization. Further development is also needed regarding the thrusters and engines themselves. Since SL testing was successful, nozzle skirt development allowing for efficient in-space applications is required. Vacuum performance confirmation must follow. Qualification of engine manufacturing processes and engine qualification for space applications may strongly differ depending on whether institutional missions or newspace clients are considered.
• While the presented results are promising, they consider mostly advancements in thruster and engine development. Wide technology application may only be possible with advancements regarding system-level 98% HTP applications. Propulsion system components compatible with the oxidizer must become more widely available. This includes not only fittings, valves, seals but also propellant tanks with Propellant Management Devices, which allow for enhanced storability. Current fluidic components and systems include stainless steel components, which are not acceptable for some long-duration missions. Material compatibility issues are of vital importance since numerous aerospace alloys are incompatible with 98% HTP. The space propulsion community must build a database and have wider access to data on the behavior of long-term HTP wetted materials. However, high grades of HTP (concentration- and purity-wise) and modern innovations regarding HTP stabilization, along with new propellant tank technologies (including surface finishing and passivation), should allow for HTP use in missions beyond a duration of a few years. Newspace application-focused HTP grades (beyond the 98 grade from MIL-PRF-16005F) and their standardization should also be considered for future research.
• This paper, whilst focusing on hydrogen peroxide, presented hydrogen peroxide and nitrous oxide as two major candidates for modern non-cryogenic in-space propulsion systems. It is clear that both may be successfully used. The authors believe that the growing space technology market leaves plenty of room for successful market application of both types of technology.

With the goal of wide implementation of 98% HTP in flight systems, this paper’s aim is to turn the attention of the space community to new opportunities provided by modern application of HTP as the oxidizer in liquid rocket engines. New projects and market needs may allow for successful wide-reaching applications of the presented technologies.