A Review of Recent Developments in Hybrid Rocket Propulsion and Its Applications

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 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.

1. Introduction of Hybrid Rocket Propulsion

1.1. Historical Perspective

Recently, space technological development has entered a new era of potential applications, including satellites, space travel, space station development, asteroid mining, and even the colonization of the moon and Mars, among others. Compared with the glorious Apollo Space Program in the 1960s and 1970s, advanced space propulsion technology has made significant progress, making rockets and satellites no longer exclusive to the country’s institutions. Private enterprises also play a pivotal role in the burgeoning launch business. For example, the well-known private company called SpaceX, which was founded in 2002, had successfully recovered the first-stage rocket booster of Falcon 9 for the first time in human history in 2015. In addition, NASA plans to use the SLS (space-launch system) rocket to land on the moon again in 2025 and formally establish human territory on the land outside the Earth for the first time in history.

Regarding rocket propulsion, humans are currently developing various types of chemical engines for different missions. Chemical rockets generally include three types, namely solid rocket motor (SRM, hereafter), liquid rocket engine (LRE, hereafter), and hybrid rocket engine (HRE, hereafter). Surmacz and Rarata (2009) [1] mentioned that “To make Space widely accessible, it is necessary to reduce the development, production, and utilization cost of space transportation systems”. In addition, should the advantages and disadvantages of the three types of rockets be reconsidered, the main benefits of the HRE are proposed, including simplicity, safety, stop and restart ability, and throttling ability, to name a few. And these advantages make HRE an important up-and-coming option, which was also proposed by Ordahl et al. (1965) [2] at a very early time.

The development of the HRE concept originated in 1933 when Soviet scientists Tikhonarov and Korolev used liquid oxygen (L_ox) and gelled gasoline as rocket propellants. The primary HRE development from 1933 to 2010 was described by Okninski et al. (2021) [3] in detail, as well as elaborated by Altman and Holzman (2007) [4] and Surmacz and Rarata (2009) [1]. Regarding the development of HRE in the past decade, researchers focused on dealing with HRE’s unresolved problems, such as rocket scalability issues, large motor combustion instabilities, and low combustion efficiency, for the purpose of applying HRE technology to the regime of space travel and transportation of low-earth-orbit (LEO) satellites. For example, Richard Branson founded Virgin Galactic in 2004, planning to provide suborbital flight services, and completed its first flight to 85 km above sea level in July 2021, which was elaborated on by Amos (2021) [5]. And spaceship two was applying hybrid rocket propulsion technology because of its inherent safety during operation. South Korea’s Innospace [6], Germany’s HyImpulse [7], and Australia’s Gilmour [8] all plan to launch their satellite launch vehicles using hybrid rocket propulsion. More details of the progress of the recent hybrid rocket will be explained at the end of the article.

1.2. Fundamental Principles of Chemical Bi-Propellant Rocket Thrusters

The bi-propellant rocket uses two or more compounds, which are normally known as the fuel and the oxidizer for chemical reactions through energetic combustion to produce thrust. It can be classified into three major types, including the SRM, the LRE, and the HRE, respectively. The schematic diagrams of the three types of rockets, illustrating the basic operating principles, are shown in Figure 1, the details of which are explained by Sutton and Biblarz (2019) [9] and Kuo and Chiaverini (2007) [10]. SRM is known for its simple configuration and large thrust-to-weight ratio because of an inherent high-volumetric I_sp (or density-specific impulse). From Figure 1a, the SRM mainly consists of a large outer casing that houses the well-mixed propellants in a solid state, in which the nearly enclosed space accommodates the combustion process with high temperature and high pressure. For the LRE, the major characteristics include its high performance, throttling capability, and re-ignition possibility, using a relatively small combustion chamber, as compared with the other two types of propulsion. The latter makes the thrust vector control easier, as compared with the other two types. As shown in Figure 1b, the LRE consists of two propellant storage tanks, independent and/or combined plumbing and feeding systems, and combustor(s). Compared with SRM, the LRE design and implementation become much more complicated. HRE is somewhere in between its two counterparts. Concepts of HRE can be traced back to the 1930s of the last century, the details of which are explained by Altman (1991) [11]. Figure 1c illustrates the basic system breakdown of the HRE. A typical HRE consists of a tank that stores the liquid (or gaseous) oxidizer, a single propellant feeding system, and a combustor that houses the fuel grain in a solid form. There is no perfect HRE configuration for all applications. Therefore, understanding all the characteristics and making the most of their advantages while minimizing the disadvantages is the major goal of a rocket scientist in the design of a good rocket propulsion system that fits the mission requirements requested by the stakeholders.

1.3. Characteristics, Advantages, and Disadvantages of HREs

To compare the above-mentioned three types of rocket thrusters, some basic parameters and characteristics are considered as described next.

1.3.1. Specific Impulse as Propulsion Performance Indicator

Specific impulse (I_sp) is the parameter describing the impulse obtained by consuming a certain amount of propellant (Equation (1)). It is defined as the thrust obtained per unit mass flow rate of the propellants (Equation (2)), where F and $\dot{m}$ are the thrust and the instantaneous mass flow rate, respectively. In rocket science, I_sp is also the product of two other parameters, C_f and C* (Equation (3)). Note C_f is the thrust coefficient produced by optimizing the geometry of the nozzle through the operating condition of the thruster, especially the pressure ratio. C* is the characteristic velocity that can be produced by a certain combination of propellants and is affected by the chemical compound and the composition of the oxidizer-to-fuel mass ratio (O/F ratio).

$$I_{sp}={\displaystyle \frac{\int Fdt}{g\int \dot{m}dt}}$$

(1)

$$I_{sp}={\displaystyle \frac{F}{\dot{m}g}}$$

(2)

$$I_{sp}={\displaystyle \frac{c^{*}\times c_f}{g}}$$

(3)

1.3.2. Combustion Characteristics

The combustion mechanism of a typical HRE is shown in Figure 2. The fuel grain placed in the combustor is consumed and burnt as the HRE operates continuously. This means the shape and effective surface area of the fuel grain varies as a function of time. Furthermore, the combustion of the HRE is a typical diffusion flame, while its two other counterparts may be considered as premixed flames in general. This leads to two major concerns for HRE performance. The first concern is the low rate of thermal decomposition, pyrolysis, and melting of the fuel surface. It is known that the fuel regression rate limits the mass flow rate of the fuel provided to the flame position. In general, the fuel regression rate of the solid fuel of the HRE is relatively low, which imposes a challenge for designing an HRE with a large thrust. This leads to a low thrust with a potentially shifting O/F ratio during operation because of grain geometry change during operation. The second concern is the low mixing efficiency of the fuel and oxidizer resulting in a low combustion efficiency and, thus, a low I_sp in practice. Since the oxidizer is injected into the combustor and flows over the fuel grain surface, there is an interface where the gaseous oxidizer meets and reacts with the decomposed fuel vapor. In other words, the combustion efficiency would be generally low in the regions away from the interface, since the composition of the fuel and oxidizer is far from stoichiometric. These may affect the HRE performance if not designed properly. In order to increase the combustion efficiency, most HRE combustors have a larger length-to-diameter (L/D) ratio. This may be properly resolved by introducing some enhancement mechanisms for the fuel regression rate, which will be introduced later in this paper.

1.3.3. Propellant Storage and Handling Safety

As stated in the previous section, HRE contains a fluid oxidizer and a solid-fuel grain. Although the concept of an inverse HRE has also been introduced by Bennett Jr. (1971) [12], no further application has been found in recent rocket research or designs. The separated storage of the propellant of HRE makes it relatively safe from an unexpected reaction between propellants that may cause an explosion, for example, in the SRM. The fuel grain of the HRE uses either thermoset (i.e., rubber) or thermoplastic (i.e., plastic) material, such as hydroxyl-terminated polybutadiene (HTPB), polyethylene (PE), and paraffin, to name a few. These fuels are non-volatile materials, and therefore, no fuel-leak-related explosion risk needs to be considered in the storage and handling procedures. With the two properties stated above, HRE is non-explosive compared to the two other types of rocket thrusters. Nevertheless, liquid oxidizers are still energetic substances generally. A handle-with-care strategy is still a must in the exercise of this field.

1.3.4. Throttling and Re-Ignition Capability

Assuming that the I_sp is nearly constant for a certain rocket, the instantaneous thrust level of a thruster is proportional to the mass flow rate of the propellants (Equation (2)). For the SRM, the mass flow rate is a function of the instantaneous fuel surface area, which is designed in advance and cannot be changed during operation. Therefore, the throttling capability of SRM is nearly impossible. As for the LRE, the throttling valves in the propellant feeding system control the mass flow rates of both the fuel and the oxidizer injected into the combustor. Therefore, the thrust level of the liquid engine can be readily controlled precisely by changing the mass flow rates of the propellants, should the combustion be stable. For the HRE, the oxidizer mass flow rate is controllable via a throttling valve of the oxidizer feeding system. Therefore, the thrust of the HRE can be precisely controlled in theory. Furthermore, unlike the SRM, which consumes all its propellant in one single burn, the shutdown and second or more ignitions of an HRE during operation become possible by controlling the valve(s) and other measures. In this regard, Saraniero et al. (1973) [13] presented and discussed the restart transients of hybrid rockets. Karabeyoglu et al. [14] presented a theoretical fuel regression investigation of the transient combustion of hybrid rockets, including ignition and throttling. In general, the potential shutdown and reignition of HRE is promising for a rocket with many applications.

1.3.5. Cost-Effectiveness

The cost of a rocket thruster can be classified into two parts, which include the production cost and the storage and handling cost. As stated in the previous section, the SRM system is relatively simple, and therefore, the production cost of an SRM is relatively cheap, compared to its counterparts. On the other hand, the complexity of the LRE is the highest among the three, so the production cost is generally the highest and is much more expensive than the SRM. As for the handling cost, the SRM has explosion handling precautions due to its solid-state premixed propellants that are similar to trinitrotoluene (TNT). In the case of the HRE, the production cost is somewhere between the SRM and the LRE, while the handling is highly safe and, thus, low in cost as compared with the two other types. Therefore, the overall cost of the HRE is generally lower than the SRM and the LRE if properly designed.

In general, the HRE is a safe and cost-effective type of rocket thruster, with both throttling and reignition capabilities. There are a variety of propellant options, which will be discussed in the following section, and most of these combinations can lead to satisfactory propulsion performance. The combustion and flame type of HRE may not be as ideal for combustion, with an optimal efficiency such as SRM or LRE. However, this could be well resolved by a proper engine design.

1.4. Theoretical Propulsion Performances for Different Fuel–Oxidizer Combinations

1.4.1. Typical Fuels and Oxidizers

For the propellant selection, there are several requirements that we need to take into account. Many criteria are briefly summarized as follows (aerospace notes) [15] for the readers’ reference.

• Affordability (Cost)

The fuel combination we choose must be low-cost. The overall costs, including research and design costs, fuel–oxidizer production costs, production facility costs, operating costs, transportation costs, and decommissioning costs, must be affordable;

b.

System performance

The design objective of the propulsion system is to optimize the fuel–oxidizer utilization efficiency in order to meet the requirements of the mission. And therefore, the combustion performance of the fuel–oxidizer combination and its storage and handling in the overall propulsion system must also be taken into account properly. It is generally better with a high I_sp, less propellant residual, good shutdown and re-ignition characteristics, ease of storage, and simple handling procedures;

c.

Survivability (Safety)

All risks and hazards generated during handling must be well understood and prevented in advance. In case of failure, the risk of personnel injury and damage to equipment, facilities, or the environment must be minimized;

d.

Reliability

The impact of reliability on the overall system is critical. All the risks, including technical, manufacturing, and failure risks, shall be minimized and well understood;

e.

Controllability

A stable combustion process is a requirement for the control process. The thrust level shall be based on the system requirements maintained within a reasonable range. The responding time of control or command signals shall also be based on the system requirements under acceptable tolerances;

f.

Maintainability

System simplicity, reliable diagnosis, easy parts access, and replacement are characteristics of maintainability. Maintaining processes shall cause a minimal hazard to service personnel;

g.

Geometric constraints

All the propulsion systems need to meet the geometric constraints of the vehicle, namely, they need to fit into the vehicle;

h.

Operability

Operation simplicity and a verified operation manual/instructions are the requirements. Procedures for loading propellants and preparing power, launch, and igniter checks must be simple, robust, and reliable;

i.

Producibility

The propulsion system shall be easily manufactured, inspected, and assembled. All critical manufacturing processes shall be well understood. All components must have well-known characteristics and be easily inspected during and after the production;

j.

Environmental acceptability

The propellant and its exhaust plume shall be non-toxic and not cause unacceptable damage to personnel, equipment, and surrounding rural areas.

Okninski et al. (2021) [3] have reviewed the space transportation developments using HRE propulsion worldwide from 2005 to 2020. Combined with the consideration of the propellant selection criteria, we can find that most of the traditional HREs use a combination of solid fuel and liquid oxidizer. Nevertheless, we might use the combination of a solid oxidizer, such as ammonium perchlorate, with a liquid fuel, such as kerosene, hydrazine, or liquified hydrogen. This kind of combination is called a reverse HRE. But due to the limited number of options for solid oxidizers and the difficulty in fabrication (Kuo and Chiaverini, 2007) [10], a reverse HRE configuration is much less common and will not be considered in this paper.

Thermoset and thermoplastic forms of polymers are currently used in most solid fuels for HRE. In addition to natural rubber, the most significant class of solid fuels is polymer-based synthetic rubber based on a polybutadiene monomer (PB). Typical PB-based polymers include PB-acrylonitrile (PBAN), PB-acrylic acid (PBAA), carbon-terminated PB (CTPB), and hydroxyl-terminated polybutadiene (HTPB), among others. Among them, HTPB is a very classic fuel for HRE because it is low-cost, commercially available, and very safe to handle. Plexiglass (polymethyl methacrylate or PMMA), polyethylene (PE), paraffin waxes, and other polymers are also commonly found in research due to their low cost and availability (Humble et al., 1995) [16].

The liquid oxidizer used in the HRE can be in the form of a liquid or a gas, and it is basically the same as those used in the LRE. These typical examples are O₂, N₂O, N₂O₄, H₂O₂, HNO₃, FLOX, inhibited red fuming nitric acid (IRFNA), chlorine trifluoride (ClF₃), and hydroxyl amine nitrate (HAN). Kuo and Chiaverini (2007) [10] have summarized the performances of the common combinations of HRE propellants, which are included in

Table 1. Performance of hybrid rocket propellants, Pc = 34.45 bar, Pe = 1 atm (Data from Kuo and Chiaverini, 2007 [10]).

Propellant Combination (Fuel + Oxidizer)	Optimum O/F	Sea Level Isp (s)	C* (m/s)	Remarks
Carbon, Air	11.3	184	1224
Carbon, Lox	1.9	249	1599	Cryogenic
Carbon, N2O	6.3	236	1522
Cellulose, Gox	1.0	247	1572
CH4(s), Lox	3	291	1871	Cryogenic
CH4(s)/Be(36%), Lox	1.3	306	1918	Cryogenic
HTPB, F2 + Lox	3.3	314	2045	Cryogenic
HTPB, IRFNA	4.3	247	1591
HTPB, Lox	1.9	280	1820	Cryogenic
HTPB, N2O	7.1	247	1604
HTPB, N2O4	3.5	258	1663
HTPB/Al(40%), Lox	1.1	274	1757	Cryogenic
HTPB/Al(40%), N2O	3.5	252	1637
HTPB/Al(40%), N2O4	1.7	261	1679
HTPB/Al(60%), F2 + Lox	2.5	312	2006	Cryogenic
Li/LiH/HTPB, F2 + Lox	2.8	326	2118	Cryogenic
NH3(s)/Be(26%), Lox	0.47	307	1967	Cryogenic
Paraffin, Lox	2.5	281	1804	Cryogenic
Paraffin, N2O	8.0	248	1606
Paraffin, N2O4	4.0	259	1667
Pentance(s), Lox	2.7	279	1789	Cryogenic
PE, N2O	8.0	247	1600
PE, Lox	2.5	279	1791	Cryogenic
PMMA, Lox	1.5	259	1661	Cryogenic
JP-4, AN	17.0	216	1418	Reverse HRE
JP-4, AP	9.1	235	1526	Reverse HRE
JP-4, NP	3.6	259	1669	Reverse HRE

.

1.4.2. Density-Specific Impulse

I_sp is a measure of the amount of thrust that can be produced by a given amount of propellants and is independent of the propellant density. On the other hand, density-specific impulse (Density I_sp) is a performance indicator that combines specific impulse with fuel density. It is the product of the specific impulse and the fuel density (Equation (4)) (Sutton and Biblarz, 2019) [9].

$${Density I}_{sp}={\rho }_{av}\times I_{sp}$$

(4)

where Density I_sp is the density-specific impulse in g-s/cm³, ρ_av is the average density of propellants (fuel and oxidizer) in g/cm³, and I_sp is the specific impulse in seconds. Density I_sp is a measure of how much thrust can be produced per unit volume of propellants. Density I_sp allows us to compare the performance of different propellant combinations under the same volumetric conditions. In addition, I_sp is a measure of the amount of thrust produced per unit mass flow rate of the propellants, while the density-specific impulse is a measure of the amount of thrust produced per unit volume flow rate of the propellants. Specific impulse is useful for comparing different propulsion systems, while Density I_sp is useful for comparing different propellant formulations for the same propulsion system.

Figure 3 shows the I_sp and Density I_sp performances of common LREs and HREs at different O/F ratios. The results used NASA CEA [17] and ProPEP3 [18] for the calculations, which are based on a chamber pressure of 40 bar and a nozzle expansion ratio of 20 with shifting equilibrium. Although the L_ox/RP-1 combination has the highest theoretical vacuum I_sp, using 90% hydrogen peroxide can achieve a Density I_sp of nearly 400 g-s/cm³, which is much higher than the L_ox/RP-1 liquid rocket combination (~350 g-s/cm³), and much higher than N₂O as an oxidizer for the HRE (~230 g-s/cm³). The latter also explains why the high-power hybrid rocket using N₂O is generally very bulky. This represents that the HRE system can achieve higher fuel efficiency in a smaller volume if the propellants are properly chosen. The smaller volume also means a lighter structural weight, which is quite significant for the overall propulsion system’s weight reduction for a launch-vehicle design.

1.5. The Challenge for Hybrid Rocket Space Transportation

Recently, due to the rise of the new space industry, numerous satellite launch companies were founded. More than 100 newly designed small satellite launchers around the world are currently under development and are expected to be operative within five years (Tugnoli et al., 2019) [19] (Niederstrasser, 2021) [20] (Kulu, 2023) [21]. Low cost effectiveness, simplicity, reliability, low complexity, and safety are the primary criteria for massive access to space nowadays. And this also provides a good motivation and opportunity for the development of HRE-based launch vehicles (Casalino et al., 2022) [22]. Furthermore, if the propellant is chosen only to contain hydrogen, carbon, nitrogen, and oxygen atoms, the combustion products would be only nitrogen, water, and carbon dioxide, which is indeed environmentally friendly. As shown in Figure 4, if the target is to have a cost-effective rocket with fair performance, the HRE propulsion system does have the opportunity to achieve the goal by properly designing various key components. Although the current maturity of HRE development is much lower than solid and liquid rockets, the barriers to entering this field are comparatively lower as well. Therefore, many private companies in the new space industry have set the goal of using the HRE, such as Virgin Galactic from America [23], Blushift from America [24], Vaya Space from America [25], Firehawk from America [26], Gilmour from Australia [8], ATspace from Australia [27], Innospace from South Korea [6], SpaceRyde from Canada [28], Reaction Dynamics from Canada [29], Letara from Japan [30], HyImpulse from Germany [31], Pulsar Fusion from the United Kingdom [31], T4i from Italy [32], SpaceForest from Poland [32], DeltaV from Türkiye [32], and HyPrSpace from France [33], among others. Okninski et al. (2021) [3] have summarized and listed the teams in more detail that have developed the HRE propulsion system for space transportation since 2005.

Nevertheless, some technical challenges still exist in the development of the HRE for space transportation. Karabeyoglu (2017) [34] and Mazetti et al. (2016) [35] mentioned that the main disadvantages include difficulty in reaching high regression rate values using standard fuels and lower combustion efficiency due to complex physical–chemical diffusion flame characteristics, which causes lower volumetric loading and challenges to the design of high-thrust engines. Okninski et al. (2021) [3] also mentioned that various factors, including propellant selection, combustion instabilities, O/F ratio shifts, throttling effect, fuel residual, thermal insulation for long burning time, and how to design an efficient thrust vector control system for an HRE propulsion system, are also needed to be considered in space transportation. These challenges are described below and some potential methods for mitigating these issues are summarized. Since the main purpose of developing a hybrid rocket is to reduce the cost of space transportation with sufficient efficiency in many aspects, the most important guideline to bear in mind is that the solution should not jeopardize the advantages of classical hybrid rockets, which is otherwise unnecessary in essence.

2. Oxidizers and Fuels for Hybrid Rocket Propulsion

2.1. Oxygen (O₂)

The liquid-phase oxygen (L_ox) and gas-phase oxygen (G_ox) are widely applied for the HRE system because of their convenience of handling and high combustion efficiency (McFarlane et al., 1993) [36] (Greiner et al., 1994) [37]. Though oxygen provides these benefits, however, the L_ox or G_ox requires the storage method to prevent storage tank damage from cryogenic (L_ox for example) or high pressure (G_ox for example). The high pressure indicates the requirement for a thicker and, thus, heavier storage tank in a flight mission (Ventura and Heister, 1995) [38].

To improve the poor combustion properties of HRE, several studies proposed a new simple technique that applied swirl to an oxidizer flow. Yuasa et al. (1999) [39] indicated that, due to the high fuel surface pressure caused by the swirling injector, the regression rate of the polymethyl methacrylate (PMMA) fuel tends to increase. Quadros and Lacava (2019) [40] indicated the swirling injection method was also applied to liquefiable fuels, and the test results pointed out that the pre-chamber length and the angle of an entry slope on the grain do not contribute significantly to the performance or stability of the motor with swirl injection. Bianchi et al. (2015) [41] elaborated on the comparative results of numerical simulations and the experiments of the lab-scale flow field in a G_ox/HTPB HRE. The results show that the simulation methods can predict the main ballistic features of the HRE. Story (2006) [42] from NASA, on the other hand, completed the large-scale L_ox-based HRE tests and demonstrated the throttling capability, simplicity, and restarting characteristics.

2.2. Hydrogen Peroxide (H₂O₂)

Hydrogen peroxide (HP or HTP hereafter) is also an oxidizer that has been widely used for propulsion (Walter, 1954) [43]. During WWII, the Germans operated the V2 rocket using 90% hydrogen peroxide reacting with the catalyst (potassium permanganate) to make the oxidizer pump work. According to the literature (Walter, 1954) [43], there was never a reported failure that was directly caused by the HP system. Moreover, HP was also can be used as an oxidizer for propulsion systems (Wernimont and Heister, 1995) [44]. Due to the fact HP is in the liquid phase at room temperature, and it is of high density, these advantages make the propellant feeding system of HP-based rockets simpler to design and the onboard storage tank structure lighter. But the highly reactive and corrosive property of HP has its operating concern. The potential HP operating hazard includes storage tank liner corrosion, auto-ignition, and detonation under certain conditions (Hannum, 1972) [45]. The handbook mentions the risk of detonation when the 26% mole fraction of HP vapor operates under high pressure (Constantine, 1967) [46], but it also indicates that, although the detonation occurs during specific conditions, the storage of the HP is easier and safe under an uncontaminated and ambient environment because of low vapor pressure.

The application of the HP-based HRE has been investigated by several teams around the world. Norway’s Nammo completed the first launch of the sounding rocket Nucleus in 2018, reaching 107.4 km (Faenza et al., 2019) [47]. Taiwan’s Advanced Rocket Research Center (ARRC) completed the hovering flight test in 2020, showing the throttling capability and thrust vectoring control of the HP-based HRE system (Wei et al., 2022) [48]. Later on, the same team performed the first flight test of the HP-based single-stage quad-HRE with autonomous guidance and control capability in 2022. The team from Poland also developed the ILR-33 Amber sounding rocket designated for microgravity experiments, which is under development at the Institute of Aviation in Warsaw (Marciniak et al., 2018) [49], with a 98% HP-based main engine and a solid-fuel booster (Okninski et al., 2021) [50]. Later on, the ILR-33 Amber 2K sounding rocket reached an altitude of 101 km in 2024 (Parsonson, 2024) [51]. The Amber team will perform extensive research in microgravity in the near future. Gilmour [8] in Australia is also developing a hydrogen peroxide-based hybrid rocket satellite launch vehicle.

2.3. Nitrous Oxide (N₂O)

Under atmospheric pressure conditions, nitrous oxide (N₂O) is a transparent, odorless, and non-toxic gas. The vapor pressure of N₂O is about 30~60 bar, depending on the temperature. In practical applications, N₂O is stored in a pressure vessel at vapor pressure, which means liquid and gas co-exist and are in an equilibrium state. Many inherent advantages and features were introduced by Robert Goddard in 1914 [9]. Its characteristics of stability, ease of storage in low ambient temperatures, and self-pressurization have made it very popular as a selected oxidizer in rocket programs. Virgin Galactic’s SpaceShipOne spacecraft was powered by an HRE using N₂O and HTPB (Kelly et al., 2017) [52] (Thicksten et al., 2008) [53], as well as the Hybrid Engine Development (HyEnD) hybrid sounding rocket project by the Institute of Space Systems at the University of Stuttgart (Kobald et al., 2019) [54] and the design and development of a 100 km N₂O/paraffin HRE vehicle by Stanford University (Dyer et al., 2007) [55].

The main benefit of using N₂O as a rocket oxidizer is its self-pressurizing ability. The vapor pressure at room temperature is about 52 bar. Therefore, an extra pressurizing system, such as a pressurant tank or pump in the plumbing system, is not required to make the oxidizer flow into the combustion chamber. Except for the benefit explained above, Zakirov et al. (2001) [56] introduced more fundamental properties of nitrous oxide, making it attractive as a rocket propellant, including decomposition exothermically with the adiabatic decomposition temperature reaching ~1640 °C, and the free oxygen available by decomposition can be combusted with a wide variety of fuels.

In actual operation, Arves et al. (1997) [57] elaborated on the effect of the saturated vapor pressure of N₂O at different temperatures. Liquid vaporization is an endothermic process that causes its temperature to drop sharply, further leading to a decrease in saturated vapor pressure. This feature also makes nitrous oxide have the negative aspect of not affording a long burning time. On the other hand, lower oxidizer density causes a bigger pressure tank, which makes the system very bulky, and it is difficult to reduce the weight, especially when the rocket scales up to a large size.

In addition to the shortcomings mentioned above, N₂O has other significant problems. Among them, N₂O needs an extremely high activation energy to be decomposed into oxygen and nitrogen for combustion with extremely high temperatures, namely 520 to 850 °C (Hennemann et al., 2014) [58], even higher than 1200 K (Saripalli and Sedwick, 2020) [59], which is much higher than the hydrogen peroxide mentioned previously. In addition, N₂O is a two-phase flow during operation, which makes its application in the deep-throttle complex. Whitmore et al. (2014) [60] proposed that the N₂O has the capacity to throttle under some specific conditions; one of the conditions is that the upstream must remain in the state of supersaturated liquid. However, during the throttle-down procedure, the fluid’s temperature will drop to the saturation temperature, resulting in the vaporization of the liquid. When the opening angle was further reduced, nitrous oxide passing through the throttle valve would have a cavitation effect, which increases the possibility of combustion instability in the combustion chamber, which may result in the loss of the flow rate control. Furthermore, Whitmore et al. (2014) [61] also pointed out that N₂O had another significant problem, which is that, once the nozzle started eroding, the throttling characteristic would become unpredictable.

2.4. Nytrox

Among all the oxidizers used in HRE propulsion, O₂ and N₂O are the common propellant due to their wide availability, broad base of use, cost effectiveness, relatively low hazard level, and environmental friendliness. However, each of them has its own shortcomings, such as G_ox having low density and high storage pressure, L_ox having to be stored at deep cryogenic temperatures, and N₂O being very sensitive to temperature. In order to maximize the benefits of the pure components while retaining their practical advantages and reducing their disadvantages, Karabeyoglu and SPG engineers (2009) [62] (2014) [63] propose a new class of oxidizers based on mixtures of N₂O and O₂ referred to as Nytrox.

A comparison of the L_ox, N₂O, and Nytrox as oxidizers is summarized in

Table 2. Comparison of pure O₂ (L_ox), N₂O, and Nytrox as oxidizers (Data from Karabeyoglu, 2009 [62]).

Feature	Lox	N2O	Nytrox
Density	4	2	4
Isp Performance	5	3	4
Impulse Density	4	1	3
Self-Pressurization Capability	1	5	3
Performance Tuning Capability	1	2	5
Chemical Compatibility	5	5	5
Chemical Stability	5	4	4
Gas-Phase Combustion	1 *	3	5
Hypergolicity	1	1	1
Motor Stability/Efficiency	2	5	4
Performance Tuning Capability	1	2	5
Storability	1	5	3
Toxicity	5	4	4
Ease of handling	3	5	4
Material Cost	5	4	4
Overall Safety	3	2	5

5: best performance. 1: worst performance. * He pressurization.

. The table shows the clear advantage of Nytrox over the pure substances in many key aspects, which would allow the designer to formulate an oxidizer ideal for the particular application of interest. The basic idea is to combine the high vapor pressure of dissolved oxygen (as the pressurizing agent) with the high density of refrigerated N₂O (as the densifying component) to produce a safe, non-toxic, and self-pressurizing oxidizer with high density and good I_sp performance. The main idea is summarized in Figure 5. Figure 5a shows how the vapor pressure of a Nytrox mixture depends on the mole fraction of oxygen in the liquid and vapor phase at a given temperature. As increasing amounts of oxygen are dissolved in the N₂O liquid, the vapor pressure varies from 20 to 120 atm. This is in the useful range for a rocket propulsion system. Note that, at mixture equilibrium, the vapor phase is mostly oxygen, greatly reducing the possibility of direct N₂O decomposition in the oxidizer tank. Figure 5b shows the relation between liquid density and vapor pressure for L_ox, N₂O, and Nytrox at various temperatures. Note that the densities and vapor pressures of the pure substances are very sensitive to temperature. Adding oxygen to refrigerated N₂O increases the vapor pressure of the mixture, with only a relatively small decrease in liquid density at a fixed temperature, as indicated in Figure 5b.

In practical application, Whitmore and Stoddard (2019–2020) [64,65,66,67] replace G_ox with Nytrox for a small spacecraft HRE thruster with acrylonitrile butadiene styrene (ABS) fuel. They created Nytrox by bubbling gaseous oxygen under high pressure into ice-bathed N₂O until the solution reached saturation level and conducted tests for performance comparison. It was shown to work effectively as a “drop in” replacement for G_ox, exhibiting a slightly reduced I_sp and regression rate, but with the trade of significantly higher volumetric efficiency. The further research about the cold-start ignition latency of Nytrox was conducted by Whitmore and Frischkorn (2020) [68], and solutions for reducing or eliminating the Nytrox-caused ignition latency were presented.

Other applications of Nytrox in recent years include the following. Sella et al. (2020) [69] developed a Nytrox/paraffin HRE, which uses Nytrox96 (liquid-phase mass composition 96% N₂O + 4% O₂ @ °C, 45bar), and Kumar and Thamizarasan (2022) [70] studied on different combinations of the gaseous N₂O and oxygen, i.e., 50:50, 70:30, 80:20 ratios with wax-based fuel. The study of Kumar uses gaseous N₂O and O₂, which were injected separately without any premixing. The test result shows that the 50:50 ratio was the most promising in all respects of performance improvement.

Compared to conventional oxidizers, Nytrox has better self-pressurization characteristics than N₂O. Nytrox provides further safety and cost improvements without sacrificing performance. Since the mixture allows for two independent control variables (temperature and pressure), different compositions of Nytrox were applied in several studies, and could easily be fine-tuned to optimize for a particular application in the future. The disadvantage is that it is more difficult to ignite and lower the operation temperature than N₂O, and the impact of the change of dissolved oxygen during the self-pressurized operation on engine performance deserves further investigation in the near future. A short review of the experimental activities based on H₂O₂, N₂O, and Nytrox for hybrid rocket propulsion applications was conducted by Paravan et al. (2023) [71].

2.5. Thermoplastic and Thermoset Fuel

Typical HRE solid-fuel grains are mostly polymers, including thermoplastics and thermosets. For thermoplastics, high-density polyethylene (HDPE), low-density polyethylene (LDPE), polymethyl methacrylate (PMMA), and polypropylene (PP) are widely used for their maturity of processing, storage, chemical compatibility, and commercial availability. For thermosets, hydroxyl-terminated polybutadiene (HTPB) has been used extensively in solid rockets and is a popular choice in hybrid rockets.

Table 3. Comparison of the physical properties of various solid fuels.

	Density (g/cm3)	Melting Temperature (°C)	Tensile Strength (MPa)	Structural Formula
LDPE	0.919	110	10
HDPE	0.952	130	26
PP	0.904	165	34
PMMA	1.185	160	75
HTPB	0.9~1.5 [72]	-	3.0 [73]
Paraffin wax	0.9	68	0.885~0.994 [74]	(C31H64 paraffin)

summarizes the main physical properties of common solid fuels. Okninski et al. (2021) [3] summarized the hybrid propellants used in various vehicles since 2005. The distinct material properties of different solid-fuel grains are described, in turn, next.

PMMA has a high density of 1.18 g/cm³, great mechanical strength, and excellent visible light transmittance. Most PMMA fuel grains are used in lab-scale HRE; the transparent nature makes it easier to visualize fuel regression from a transverse direction throughout the axial direction (Yuasa et al., 1999) [39].

For polyethylene (C2H4)_n, both LDPE and HDPE can be considered as saturated hydrocarbon, with LDPE having a branched structure lowering the regularity, thus reducing the density, melting point, and crystallinity compared to that of HDPE. Polypropylene (C3H6)_n (PP) is mostly liner isotactic in molecular structure, with a similar density to polyethylene and a higher melting temperature.

HTPB is a translucent liquid that can be reacted with di- of polyfunctional isocyanates (NCO) to produce polyurethanes. The cross-linking of terminal hydroxyl groups (OH) in the HTPB prepolymer and the NCO results in an elastomeric network capable of withstanding harsh stress loadings in flight (Davenas, 2003) [75]. Villar et al. (2011) [76] pointed out that the NCO/OH ratio can affect the mechanical properties, as well as thermal aging, regarding the service life of HTPB.

In search of fuel grains with a higher regression rate, n-alkanes petroleum waxes, such as paraffin wax and microcrystalline wax, are also used as solid fuels. Owing to the low melt viscosity, wax-based fuels exhibit an over 3-fold regression rate increase compared to that of HTPB fuel (Karabeyoglu et al., 2004) [77]. However, due to concerns about wax structural integrity, most research considered blending polymers into the fuel to increase fuel mechanical strength, even being enhanced with 3D-printing structures, which will be explained in the next section.

If it is for environmental friendliness, sustainability, and, ideally, carbon neutrality, Barato (2023) [78] considers polyethylene, polypropylene, and paraffin wax as the best candidates because of their performance and ease of reuse.

2.6. Fuel Regression Rate

The combustion process of HRE is considered a typical turbulent diffusion flame. A typical HRE configuration is illustrated in Figure 6, where a liquid oxidizer is injected axially into the core of solid fuel (Marquardt and Majdalani, 2019) [79]. Upon ignition, the heated solid fuel is vaporized and moves radically from the fuel surface to the core, where it mixes with the gaseous oxidizer flow. Once combustion occurs, the energy released from chemical reactions provides sufficient heat to sustain the entire combustion process. A flame sheet is formed and located where the mass ratio of the fuel to oxidizer is stoichiometric.

The fuel regression rate plays a vital role in HRE design because a designer can specify a desired thrust level by determining the operating O/F ratio. However, the complicated combustion physics makes fuel regression rate prediction very challenging. Models that could predict the fuel regression rate are often referred to as the “regression rate law”. Marxman and Gilbert (1963) [80] developed a turbulent boundary-layer regression rate model based on the heat-transfer mechanism to predict the local instantaneous fuel regression rate. This model furthers the development of Lees’s (1958) [81] analytical model of turbulent convective heat transfer within a reactive blowing environment. Because of the high stream velocity and the blowing effect (fuel evaporation), which reduces the transition Reynolds number, the boundary layer is expected to be turbulent over most of its length. Although several experiments have proven that this model is valid, Marxman’s regression rate law somehow overpredicts the fuel regression rate. Thus, it is inaccurate for rocket design purposes. The overpredicted value is because the model is based on a slab fuel configuration, while most HRE fuel today is cylindrical. In addition, Marxman assumed that the flow is developing along the entire combustion chamber, making his model fail to predict the entrance region dominated by injector effects, and the HRE flow becomes fully developed. Nevertheless, Marxman’s regression rate law identifies and relates the factors that influence the fuel regression rate. In 2006, Karabeyoglu and Zilliac (2006) [82] proposed a new model that improved Marxman’s regression rate law, making it applicable to a cylindrical fuel grain and a pre-combustion chamber configuration. However, their regression rate law suffered a high oxidizer mass flux exponent and was sensitive to fuel surface temperature and port gas viscosity. The noteworthy achievements of their work were that they further identified factors that influenced the regression rate and improved the regression rate model by considering the entrance effect and a cylindrical fuel configuration.

Although many researchers have tried to propose a generalized model for determining the regression rate based on the fundamental properties of propellants, none of these theories can accurately estimate the regression rate value due to the lack of an accurate combustion model for HRE. The prediction models yield significant errors when applied to different engine configurations, such as scales and propellant combinations, and the design of injectors. For the above reasons, most research uses an empirical regression rate law for rocket design purposes. With several ground test data, a curve-fitting technique is used to compute the coefficient and exponents of regression rate law. The typical regression rate law is written as follows:

$$\dot{r}=a{G}_{ox}^n{x}^m$$

(5)

where x is the axial distance of the fuel, $G_{ox}^n$ is the oxidizer mass flux, prefactor a, and exponents n and m are determined experimentally and will vary with different fuel–oxidizer combinations. In addition, mass flux is expressed in terms of oxidizer mass flux rather than total mass flux, since oxidizer mass flux is relatively easy to determine from experiments. The drawbacks of this approach are the high cost and time consumption because extensive static-fire tests with different test conditions are required to construct the regression rate law.

Figure 7 (Sutton and Biblarz, 2019) [9] illustrates the overall regression rate behavior at different oxidizer mass fluxes. The intermediate range of mass flux is turbulent diffusion dominated, inferring that the combustion process is entirely governed by turbulent heat and mass transfer. In the turbulent diffusion-limited regime, the regression rate is pressure-independent, and its dependence on mass flux is consistent. It should be noted that most of the practical operation conditions of HRE are in this regime. In addition, throttling HRE becomes possible because of the consistent relationship between the regression rate and the mass flux.

On one hand, in the low mass flux regime, radiative heat transfer becomes important due to the weaker turbulence intensity and the radiating particles. In this regime, the regression rate is pressure-sensitive. Chiaverini et al. (2000) [83] have found a pressure-dependent phenomenon in their experimental study of HTPB fuel grains. They observed a strong difference from Marxman’s convection-driven model in the lower oxidizer mass flux regime, while the test data agree well with Marxman’s model in the higher oxidizer mass flux regime. Further, the regression rate increases at higher chamber pressure with a low mass flux, implying that radiation effects become significant in HTPB HRE.

On the other hand, in the high mass flux regime, the chemical kinetics becomes dominant and replaces the diffusion mechanism at low chamber pressures, since the pressure-sensitive reaction rate is slow enough to be a rate-limiting mechanism. A group of researchers has devised models to depict combustion behaviors where chemical kinetics plays a significant role. Miller (1966) [84] has developed a regression rate model by incorporating turbulent mass transfer and chemical kinetics. His model has successfully correlated experimental fuel regression rate data, including gaseous fluorine and oxygen. Smoot and Price (1967) [85] used numerous experiments to confirm the pressure dependency on the fuel regression rate. They also found that the regression rate became nearly mass flux-independent and pressure-sensitive once a threshold value of mass flux was reached. Moreover, Ramohalli and Stickler (1971) [86] proposed a polymer-degradation theory to explain pressure-dependent combustion. They believed that the unreacted gaseous oxidizer flux to the wall would cause surface oxidative depolymerization, thereby aiding the thermal degradation for producing vaporizable fuel.

Despite recent efforts to develop a comprehensive regression rate theory for hybrid rockets, Marxman’s turbulent diffusion-limited model is still the most used model that provides a reasonably adequate prediction at moderate pressures and mass fluxes. Some correction terms are required to increase the accuracy at low pressures and low mass fluxes conditions. In addition, Marxman’s model’s high sensitivity to the given parameters hinders designers from predicting performance for different configurations. Consequently, designers prefer using the empirical regression rate law for hybrid rocket development in practice. But the scaling issues remain one of the major challenges in HRE technology development.

3. Specific Issues for Hybrid Rocket Propulsion

The design options for hybrid rockets are pretty diverse, and there may be different design options according to different application requirements. For example, the applications of upper stage, kick stage, or related in space propulsion tend to have smaller thrust and longer combustion time, so a high fuel regression rate may not be necessary conditions. However, if the application is a high-thrust propulsion system, how to increase the fuel regression rate is an important research direction. Therefore, in many applications, typical downsides of an HRE are low regression rate and poor mixing, which leads to low thrust and low combustion efficiency, respectively. Increasing heat-transfer efficiency from the reaction zone into the fuel surface may speed up the regression rate of the fuel. Sometimes, it may also boost the mixing and combustion efficiency. Applying swirl injection of the oxidizer, which would enhance the convective heat transfer to the fuel surface, is another way to increase the regression rate and mixing (Pal et al., 2021) [87] (Okninski et al., 2021) [3]. In the following sections, a variety of engine enhancements proposed by worldwide HRE engineers and scientists to improve the two stated problems will be discussed next.

3.1. Solid-Fuel Grain Design for Enhancing Fuel Mass Flow Rate

3.1.1. Star-Port Design

An intuitive method to increase the fuel mass flow rate of HRE, which is similar to SRM, is by increasing the perimeter-to-area ratio of the solid fuel, which naturally leads to a star-shaped port (Figure 8). Heeg et al. (2020) [88], students from Germany, designed and tested a star-port HRE named HYDRA 4X that was operated using HTPB and N₂O and produced a thrust level of 100~230 kgf. Tian et al. (2022) [89] from China published an article about the numerical investigation of different HRE designs, including a base case, a segmented star, and a staggered segmented star (Figure 9). The propellants chosen for this work are HTPB and 90% H₂O₂. A variety of cases were studied, and one of them was verified by a hot-fire test. They claimed that flow vortices are found in the segment design and a staggered setup may further boost the vortex intensity. While investigating the temperature distributions from the numerical results, they also found an increase in temperature, from 3030 K to 3113 K, in the latter design. The overall combustion efficiency from their study using star-port designs may exceed 91%, which is more than a 4% enhancement of their base case.

3.1.2. Multi-Port Design

Splitting a single large port of the HRE fuel into several smaller ones is another approach to increase the perimeter-to-area ratio of a cross-section of the fuel grain. Multi-port HRE studies can be found in the articles published by Kim et al. (2013) [90], Tian et al. (2014) [91], Ahn et al. (2018) [92], Yun et al. (2021) [93], and AMROC (Boardman et al., 1996) [94] (Story et al., 2003) [95], among others. Implementing the multi-port design for the solid-fuel grain may directly increase the fuel mass flow rate due to the increase of the fuel pyrolysis surface area. This design may also significantly increase the L/D ratio in each port, which may enhance the mixing efficiency as a result.

Kim et al. (2013) [90] from South Korea investigated the fuel-burning behavior by applying G_ox on one to five equal-diameter ports of PE and PMMA fuel grains and two different scales of PE fuel grains. They noticed that the overall regression rate of three, four, and five ports is larger than a single or two ports. The test results also show that the performance of the HRE differs in the different stages of fuel evolution (the initial stage, partially merged stage, and fully merged stage). Furthermore, as the HRE fuel grain scales up, the end-burning effect of the aft part of the fuel grain starts to significantly affect the test results. They also stated that, due to the scale of their study, their conclusion was based on the test data of their work, which were all on the scale of a 10 s burn, and more studies deserved to be carried out for further applications.

Ahn et al. (2018) [92], also from South Korea, investigated the engine performance of HDPE and 90% H₂O₂ using four hot-fire tests. The cases tested were with 1 port, 2 ports, 3 ports, and 14 ports, and each was with its own port size and fuel length. The port size and length of the fuel grain were each optimized by the theoretical O/F ratio. In Ahn’s test configuration, the pressures of the pre- and post-combustion locations were both measured. They found that the initial pressure difference between the pre- and post-combustion locations is nearly identical for all four cases, and the difference decreases and disappears as time goes on during the firing. Furthermore, the time from ignition to merging decreases as the number of ports increases, as expected. Yun et al. (2021) [93] believed that it is a result of the increment of the J ratio (the ratio of port area to nozzle area) as the engine operates. They investigated the combustion behavior in a multi-port fuel grain as a function of the J ratio. They developed an in-house code to predict the HRE performance running with HDPE and H₂O₂. They found that the regression rate of each port decreases as the number of ports increases, but the overall fuel consumption increases. The port studied is in a circular shape, probably because of easy fabrication. In general, this type of multi-port fuel grain may induce serious port merging or fuel consumption residual issues. In addition, different types of multi-port designs can optimize the engine performance, which are described next.

Figure 10 illustrates several typical wagon-wheel-type multi-port designs of the fuel grains. The simplest one is that stated in (Kuo and Chiaverini, 2007) [10] by AMROC’s early tests, called the double-D. Tian et al. (2014) [91] proposed a three-port design similar to the double-D. They numerically predicted the contour behavior of the fuel and verified it with two hot-fire tests. AMROC designed an HRE with a thrust of 250 klbs using the wagon-wheel concept (Boardman et al., 1996) [94] (Story et al., 2003) [95]. The success of the test made this rocket engine the HRE with the largest thrust in their time. A lot of AMROC’s research articles could be found, and many of them are not cited in this article. For non-cited articles, it does not mean that they lack significance, and readers are still encouraged to search and read them.

3.1.3. Helical-Port Design and 3D-Printed Fuel Grains

A screw-like helical configuration in the fuel port could also induce the rotating flow effect to enhance the fuel regression. Spiral flow, which revolves around the port axis, is a similar idea to a swirl injection design. The spiral flow may thin the flow boundary layer, enhance heat transfer, increase the regression rate, promote mixing efficiency, and, thus, increase the thrust level. Lee et al. (2007) [96] fired a demonstrative PMMA HRE that operates at about 20 bars. They compared the pitches of 6 and 100 with the base case and found that the helical port fuel supported by swirl injection may have a good performance. But these types of fuel grains were quite difficult to produce in the early days. Recently, by taking the advantage of additive 3D-printing technology to manufacture a complicated port, this is no longer an issue in practice.

Due to the additive nature of additive manufacturing (AM) in 3D printing, the technique costs less in material and time compared to the conventional subtractive methods. The novel process is versatile, highly free-form, and costs less, which has been attracting attention as a replacement for the conventional subtractive methods. Generally, 3D printing using plastic materials as an ingredient can be divided into four types, including material extrusion (ME), powder bed fusion (PBF), direct ink writing (DIW), and vat photopolymerization (VP) (Oztan and Coverstone, 2021) [97]. Among these, material extrusion is the most commonly used in composite HRE fuel because it can use thermoplastics such as polylactic acid (PLA), ABS, and PMMA, which are great materials as HRE fuel grains. Yu et al. (2023) [98] investigated the combustion performance of various packing densities of ABS with grid-like structures, and (2024) [99] investigated the combustion performance of various common polymer materials that can be additively manufactured by fused deposition. Whitmore et al. (2017) [100], Marshall et al. (2019) [101], and Dinisman et al. (2024) [102] used various 3D-printed helical fuel ports to investigate their corresponding combustion characteristics.

Fuller et al. (2011) [103] demonstrated a proof of concept utilizing 3D-printing technology to manufacture PMMA-made port fuel grains with complex geometry. Follow-up studies by Armold et al. (2013) [104] were conducted by producing complex molds for paraffin fuel grain casting, and additives may also be applied. They claimed that an over 270% regression rate increase could be observed compared with some results published previously. Whitmore et al. (2015) [105] also demonstrated 3D-printed ABS fuel and tested it with G_ox. Wang et al. (2020) [106] utilized 3D-printed helical ABS as a base structure to host paraffin-based fuel grain, and oxygen is used as an oxidizer for the engine hot-fire tests. The study demonstrated that the ABS part of the fuel has a low regression rate compared with paraffin and, therefore, creates vortices in between the blades that enhance the mixing efficiency. Bisin and Paravan (2023) [107] conducted experimental research on ballistic performance for the reinforcement of paraffin-based fuels by 3D-printed cellular structures and proposed the gyroid as a suitable open-cell reinforcing structure that can effectively increase the fuel regression rate to more than 48%, with limited impact on the melt layer viscosity. In addition to 3D-printed helical port fuel grain, Saito et al. (2018) [108], Okuda et al. (2022) [109], and Hirai et al. (2023) [110] also conducted experimental research on an axial-injection end-burning HRE through 3D-printed fuel grains. Oztan and Coverstone (2021) [97] reviewed hybrid rocket fuel grain designs by using additive manufacturing techniques and stated that more extended performance tests are required to ensure their reliability and applicability, even though the 3D-printed fuel grains have the aforementioned advantages compared to the traditional manufacturing methods.

3.1.4. Special Fuel Grain Geometry Design

Flow turbulence and increasing shear intensity would break the boundary layer of a flow. Therefore, creating intensive turbulence or shear on the fuel surface might enhance the HRE performance. Gibbon and Haag (2001) [111] proposed a special design and named it the vortex flow pancake HRE (VFP hereafter). The fuel grain configuration of this HRE is extremely different from the traditional ones. The VFP consists of two disk-shaped fuels and one with a center hole (Figure 11). The oxidizer is injected in a tangential direction from the wall of the engine. The fluid flows by the surface of the disk fuel and then through the central hole of the lower disk. This type of engine is claimed to have a very high mixing efficiency. Another advantage of this type of HRE is the dramatic decrease in the L/D ratio. The L/D ratio of a conventional HRE is 10~15, and in some recent articles, the L/D may decrease to 2~5 with the implementation of some performance-enhancing methods. But for the VFP, the L/D is less than one and may be even smaller as the thrust scales-up.

Paravan et al. (2016) [112] further investigated the VFP engine design and tested using the G_ox/HTPB system. They stated that increasing the number of injectors on the circumference and increasing the flow speed may thin the boundary layer on the fuel grain surface and enhance the performance. Lai et al. (2017) [113] (2018) [114] proposed a similar design to the VFP, which is the dual-vortical flow (DVF) (Figure 12). They claimed that, by introducing another set of inverted vortical flow, the thrust may be twice as large, and the induced swirling torque of the engine, due to tangential injection, may be canceled if designed properly. Nagata et al. (2006) [115] (2022) [116] introduced another special type of fuel configuration, which is the cascaded multistage impinging jet (CAMUI). In the work of Nagata et al., several sections of fuel grains with two holes are placed one after another in a staggered manner (Figure 13). The flow from a fuel port will impinge on the next and cause intense turbulence, which enhances mixing and heat transfer.

3.1.5. Other Mixing Enhancement Mechanisms

Other means of producing swirling/vortex flows may also be found in HRE applications. Chen et al. (2013) [117] implemented tilted four pairs of blade-like structures that stand in the port of the fuel grain. They claimed that these pairs of mixing enhancers may create counter-rotating vortices in pairs that will complement each other in the flow field. The numerical results showed that the I_sp performance of a 300 kgf-class engine increased from 187 s to 222 s. They also predicted that, by adding a second set, the performance would be enhanced further. However, it would have a serious problem of erosion of the blades for long-time burning. Using diaphragms that disturb the main flow may also enhance HRE efficiencies. Zilliac et al. (2020) [118] placed a diaphragm and “efficiency-enhancing device” in the post-combustion chamber. With more than 100 ground tests performed, they observed an 11% efficiency increase on average. Tian et al. (2013) [119] employed a four-hole diaphragm at the post-combustion chamber of an eight-fin star-port HRE. Their hot-fire test comparison showed that combustion and I_sp efficiency increased from 93.9% to 97.34% and 80.77% to 87.28%, respectively. Furthermore, they also found a star shape on the nozzle erosion in the case without the diaphragm and a relatively uniform circular one in the test with the diaphragm. The diaphragm may also be placed in the front or at the center portion of the fuel grain. Kim et al. (2011) [120] placed a liquefying combustible diaphragm at the center of the fuel grain. The material of the diaphragm is 10 wt% LDPE and 90 wt% paraffin, in which the regression rate is about half of the pure paraffin. The difference in the regression rate creates a step on the fuel grain surface that serves as the diaphragm. They stated that the use of the exposed diaphragm showed a clear enhancement in the regression rate and the C* efficiency. On the downside of diaphragm implementation, pressure oscillations may occur and were investigated by Lee et al. (2022) [121]. They claimed that the vortices created by the diaphragm are not the main cause of high-pressure oscillations. Instead, “hole-tone” is the main cause due to the flow kept impinging on the front face of the diaphragm. They suggested a “stepped fuel grain” design would be in favor of oscillations concerns, which definitely requires further investigation.

3.1.6. Mixture Fuel

As HRE combustion is diffusion-limited; the solid-fuel regression rate is an order smaller than that of the SRM. However, for paraffin fuels, the regression rates are 3–4 times or even much higher as compared to those of conventional fuels. Karabeyoglu et al. (2001) [122] in the early 2000s derived an expression that shows an additional mass-transfer mechanism that accounts for the increase in the paraffin fuel regression rate. Such fuels produce a thin, unstable melt layer on the fuel grain surface. Liquid-fuel droplets are entrained and then injected into the boundary layer through the blowing of the oxidizer. The pyrolysis of liquid-fuel droplets happens between the melt layer and the flame front, which essentially bypasses the blocking effect. This entrained mass flow rate is found proportional to dynamic pressure and melt layer thickness and inversely proportional to the melt layer’s surface tension and viscosity.

Leccese et al. (2019) [123] presented several combustion visualization experiments from the last decade, in which they demonstrated droplet entrainment events through high-speed and Schlieren imaging. These experiments showed that the instability of the fuel melt layer leads to the formation of droplet entrainment, thus resulting in the growth in the regression rate. The wave-like structure over the fuel surface could be associated with the Kelvin–Helmholtz instability theory. Moreover, the droplet sphericity decreased and further transformed into a filamentous shape as the fuel viscosity increased, and eventually, the entrainment disappeared. For instance, the melt layer of the HDPE fuel would not form droplet entrainment while paraffin waxes would.

Since paraffin waxes are mechanically weak, concerns about structural failure from the high loading during motor operation were mentioned in various works. Numerous studies introduced a broad range of additives into paraffin-wax fuels, aiming to improve the mechanical properties of the paraffin-wax fuel. Most researchers employed polymeric additives, such as LDPE, EVA, and SEBS (Kobald et al., 2014) [124] (Maruyama et al., 2011) [125] (Kim et al., 2015) [126] (Pal et al., 2019) [127] (Mengu and Kumar, 2018) [128], mixing the thermosetting HTPB (Thomas et al., 2021) [129], and some utilized heterogeneous additives, such as aluminum and carbon black particles (Zilwa et al., 2003) [130] (Galfetti et al., 2011) [131]. The effect of additives improved the ultimate tensile strength, stiffness, and elongation of the fuel grain. However, for polymeric addition, since fuel viscosity increases, the regression rate would suffer a minor loss compared to pure paraffin-wax fuel. The selection of paraffin wax, percentages of one or more additives, material testing standards, and the fuel manufacturing process may vary in different studies, Veale et al. (2017) [132] suggested that, when an additive is found to meet the performance requirements, material testing should be conducted to characterize such a fuel grain, in addition to the hot-fire tests.

3.1.7. Composite Fuel with Additives

Mixed additives, such as expandable graphite, metal powder, hydride, organic salt, hypergolic additive, and even an oxidizing agent in the fuel, are among the well-known approaches to enhance the overall engine performance, fuel regression rate, and fuel density (Cantwell et al., 2010) [133] (Carmicino et al., 2014) [134] (Leccese et al., 2019) [123] (Elanjickal and Gany, 2020) [135]. In general, paraffin wax and HTPB have been found to be easy to mix with the additives, but it was found that it is difficult to disperse the mixtures uniformly. Moreover, if the nozzle is manufactured using graphite, the metal additives easily cause nozzle erosion by condensed combustion products (CCPs) that are generated during the combustion (Karakas et al., 2022) [136], thus seriously affecting the performance of the HRE.

With an HTPB-based fuel grain, Farbar et al. (2007) [137] investigated the addition of either aluminum or aluminum–magnesium alloy that reveals a correlation between the regression rate data for the metalized propellants deviated from that which was derived for turbulent convective heat-transfer-dominated behavior of pure propellants. C. Paravan (2019) [138] and Thomas et al. (2021) [139] also conducted a detailed discussion on the impact of different metallic additives on hybrid rocket combustion through experimental investigation, such as different metal compositions, sizes, and concentrations. The organic salt was also applied to the HRE fuel grain application. Wright et al. (2005) [140] point out that, under the lower oxygen flow rate condition (fuel rich), the thrust and total impulse significantly increased with the mixing of HTPB with the guanidinium azo-tetrazolat (GAT). Shark et al. (2013) [141] proposed the visualization method to observe the combustion phenomena generated by a variety of binders and energetic additives. The flame zone structure was observed by the method of optical cylindrical combustor (OCC). Shark et al. (2016) [142] explored that different flame zones were observed between aluminum, aluminum–polytetrafluoroethylene, and NaBH₄ as fuel additives, respectively. Different from the mechanism that aluminum and aluminum–polytetrafluoroethylene increased the regression rate due to enhanced heat release, the NaBH₄ additive fuel grain was observed to increase the regression rate and flame height because of hydrogen release into the flame zone.

Paraffin wax is also the famous binder to mix the additive to increase the performance of the HRE. A comprehensive assessment of performance additives for paraffin-based HRE fuel grain and preliminary test firings was presented by Karakas et al. (2019) [143]. Several metals and metal hydrons were selected to assess the thermochemical calculation results and material availability. The study showed that Al and LiAlH₄ are promising for high-performance fuel grain due to an increase in the C*, which in turn, increases delivered I_sp. Also, the additive NaBH₄ was also considered the potential additive to enhance motor performance, which can even be used as a hypergolic solid fuel if high-concentration hydrogen peroxide is used as the oxidant (Hawass et al., 2024) [144], but the storage problem of solid fuel still requires further investigation. The fuel-additive technology was also applied to the solid-fuel ramjet (SFRJ) (Karabeyoglu and Arkin, 2014) [145]. In the study, the aluminum or other metals did not effectively improve the I_sp reaction with G_ox. On the contrary, the AlH₃ additive delivered the best performance benefit due to the hydrogen release during the combustion process. However, it is not easy to store them in the ambient environment. On the other hand, although paraffin wax provides a low cost, good availability, and hydrophobic nature, the low tensile strength causes and promotes cracks when the fuel grain solidifies. To solve the problem, Matsumoto et al. (2018) [146] blended the PMMA powder to improve its tensile strength. The results demonstrated that the tensile strength increased as the added amount was increased.

In brief summary, a variety of HRE designs that may enhance the propulsion performance of the HRE have been presented. The regression rate of the fuel grain is an index that describes how much of the heat flux is transferred from the combustion zone to the fuel grain surface, which is directly related to the pyrolysis phenomena. Thus, increasing the regression rate is one of the most important goals of HRE scientists. Mixing of the fuel vapor and oxidizer is also a concern in investigating HRE performance. Swirling implementation and unique fuel geometry designs may help achieve the goals. Each of the techniques stated above has its own merits and may be used for consideration in practical applications in the future.

3.2. Swirl Oxidizer Injection

3.2.1. Experimental Studies on Swirl Injection

Oxidizer injection characteristics have significant effects on the hybrid rocket combustion process due to the different incoming oxidizer flow patterns (Carmicino and Sorge, 2005) [147] (Bouziane et al., 2019) [148]. Among these, swirl injection is a method of increasing the total momentum magnitude without increasing the mass flow rate of the oxidizer. Given a specific oxidizer mass flow rate, which means the axial flow speed is fixed, adding the tangential flow component can increase the total flow speed. Swirl injection of the oxidizer into the combustor can effectively increase the flow shear stress with the fuel grain surface and, thus, enhance the heat transfer. The higher the tangential flow speed increases, the stronger the shear stress becomes. The geometric swirl number, defined as Equation (6) below, is often used as one of the design parameters of a swirl injector.

$${SN}_g={\displaystyle \frac{(R_x-R_{\varnothing })R_x}{n{R}_{\varnothing }^2}}$$

(6)

where R_x is the injector exit radius, R_∅ is the inlet hole radius, and n is the number of the inlet hole, as shown in Figure 14.

Yuasa et al. (1999) [39] preliminarily demonstrated a visualization test on PMMA and G_ox. They claimed that implementing swirl injection has a promising impact on the fuel regression rate. Jones et al. (2009) [149] made a numerical prediction and performed six firing tests using axial and swirl injection flows, respectively. They found that the overall regression rates of swirl injection flows were much larger than axial ones and, furthermore, that the regression rates of the swirl injection were much more uniformly distributed throughout the port. The swirling injection method also increased the chamber pressure during combustion. Bellomo et al. (2013) [150] tested the swirl-injecting effect on a N₂O/paraffin system. They found that the regression rate of the fuel grain increased by 51% and the C* efficiency increased from less than 80% to more than 90%. Furthermore, they stated that the geometric swirl number should be corrected by the ratio of liquid and gas density of N₂O (Equation (7)) due to the phase change from liquid to gas after injection into the chamber.

$${SN}_{g, liquid}={SN}_{g,gas}{\displaystyle \frac{{\rho }_{gas}}{{\rho }_{liquid}}}$$

(7)

Franco et al. (2020) [151] experimentally investigated the swirl injection effect on a 300 N class H₂O₂/HDPE system. They studied the effect of the post-combustion chamber with a swirl injection HRE and found that the O/F ratio merely shifted during operation, and the post-combustion chamber length did not affect the combustion efficiency. Stella et al. (2024) [152] also experimentally investigated the H₂O₂ hybrid rocket with different swirl numbers and paraffin base fuels; it is proposed that these two design variables can be used to optimize the performance and envelope of hybrid propulsion systems according to different applications. Another type of swirl flow implementation was that which was proposed by ORBITEC™ [153,154,155]. Knuth et al. (1998) [153,154] (2002) [155] designed an HRE that tangentially injects the oxidizer from the rear end of the engine. The flow was directed to the fore-end with a spiral through the fuel grain surface. Then, the flow bounces back toward the nozzle at the center of the port, as shown in Figure 15. They described the design as the co-axial, co-swirling, and counter-flowing vortex HRE. The results showed that their HRE could have a regression rate of seven times more than the conventional ones, which is yet to be the upper bound. The proposed HRE can also dramatically reduce the L/D ratio. Lestrade et al. (2019) [156] and Lee et al. (2024) [157] proposed the design of end-burning swirling flow hybrid rocket combustion with a similar concept, and through experimental investigation, the combustion efficiency can reach about 90% and achieve a uniform O/F ratio during the burn time.

3.2.2. Numerical Investigations of Swirl Oxidizer Injection

With the progress of computational fluid dynamics (CFD), more and more robust numerical solvers are built to help researchers resolve the complex non-reactive and reactive flows, especially for HREs that use swirling injections. Cai et al. (2020) [158] numerically studied various combinations of swirl and head-end injectors. From Cai’s perspective, the regression rate may improve to 4.93 times under certain combinations. Kumar et al. (2013) [159] presented the numerical prediction of the swirling-type HRE using fuel grains with various length-to-diameter ratios. The computed temperature and concentrations of reactant and the production of CO₂ were used to show the effect of swirling flow. The results showed that the enhanced convective heat transfer was more effective for shorter fuel grains.

Paccagnella et al. (2015) [160] (2016) [161] (2017) [162] consecutively presented their numerical results that were focused on the effect of swirling injection in the HRE. A numerical model was built with a single-step reaction as the chemical mechanism, and the k-ω shear stress transport (SST) model was applied to consider the turbulence effect. Variations of the test conditions, such as different geometric swirl numbers and the oxidizer mass flow rates, were investigated systematically. And both the local swirl number and the swirl angle along the HRE axis were calculated to illustrate the characteristics of the HRE combustion field. An exponential relation between the oxidizer mass flux and the fuel regression rate was observed in the test cases. Li et al. (2016) [163] used 95% wt H₂O₂ and PE as the oxidizer and fuel, respectively, in their numerical model with a concept of multi-section swirl injection. The effect of different injection positions on the combustion field was illustrated, and the results showed that the additional injection sections increased the combustion efficiency than those without.

Barato et al. (2019) [164] numerically investigated the comparison of the swirling effect between centrally injected and sided injected. Both the accuracy of the predicted chamber pressure and the regression rate were validated in their study. Cai et al. (2020) [158] numerically modeled the combustion efficiency using high geometric swirl numbers. The results showed that the higher the swirl number, the higher the fuel regression rate and the higher the combustion efficiency. Recently, Li et al. (2022) [165] proposed the effect of regenerative cooling on the HRE by both experimental and numerical investigations. The results showed that the regenerative cooling made the fuel regression more uniform throughout the port. The reduced deviation of the regression rate along the port also enhanced the propulsive performance. For the numerical results based on the experimental tests, there were two kinds of effect that dominated the upstream and downstream burned fuel’s topology. For the upstream fuel grain surface, the regression was determined by both swirling and flow expansion effects. In the downstream fuel grain, the regression was determined by the residual swirling intensity. The numerical results were found to compare well with the experimental observations in their study, as illustrated in Figure 16.

The use of the swirling technology for the injection of the oxidizer into the combustion chamber increases the performance of HRE. However, there are, indeed, some obvious downsides. One immediate disadvantage is the inherent inevitably induced torque of the engine. Fortunately, this effect can be well-managed through the robust attitude control of modern rocket technology (Li et al., 2023) [166].

3.3. O/F Ratio Shift and Propellant Residue

O/F ratio shift is generally a commonly observed phenomenon in a conventional HRE. The main reason is that the fuel mass flow rate in the solid fuel of an HRE cannot be controlled independently from the oxidizer flow rate. According to the averaged regression rate law, the O/F ratio of a solid fuel with a single circular port can be written as:

$${\displaystyle \frac{O}{F}}={\displaystyle \frac{{\dot{m}}_{ox}}{{\dot{m}}_f}}={\displaystyle \frac{{\dot{m}}_{ox}}{{\rho }_f{A}_p\dot{r}}}={\displaystyle \frac{{\dot{m}}_{ox}}{{\rho }_{f}L\pi Da{\left(\frac{4{\dot{m}}_{ox}}{\pi D^2}\right)}^n}}={\displaystyle \frac{{\dot{m}}_{ox}^{1-n}{D}^{2n-1}}{4^n{\pi }^{1-n}a{\rho }_{f}L}}$$

(8)

where ${\rho }_f$ is the solid-fuel grain density, A_p is the circumferential surface area of the port, L is the port length, and D is the port diameter. We can see that the O/F ratio is a function of the oxidizer mass flow rate, the density of the solid fuel grain, the port length, the instantaneous port diameter, the coefficient a, and the exponent n. For a fixed oxidizer mass flow rate and port length, for n > 0.5, the O/F ratio would increase because of the increasing port diameter during burn time. In contrast, for n = 0.5, the O/F ratio would be fixed during the burn time, but the O/F ratio would shift by throttling the oxidizer mass flow rate instead. On the other hand, the nozzle throat erosion would lead to a reduction of the chamber pressure, which may change the fuel regression rate and, thus, the I_sp performance, which is also one of the factors that make the O/F ratio shift.

Figure 17 shows the theoretical vacuum I_sp as a function of the O/F ratio for the HTPB fuel reacting with various kinds of oxidizers at a chamber pressure of 40 barA with a nozzle expansion ratio of 20 and a shifting equilibrium. These were obtained using the chemical equilibrium online calculator “NASA-CEA” [17]. The results show that the theoretical I_sp cannot be maintained at the best value when the O/F ratio shifts at a constant oxidizer mass flow rate. This point is generally considered to be one of the most critical disadvantages of HRE. Kamps et al. (2021) [167] have proposed detailed mission analysis results on this point to evaluate the effect of the O/F shift and the nozzle erosion on the change in velocity (ΔV). Especially when using oxygen as the oxidizer, in which the change in propulsion performance would be highly sensitive to the O/F ratio shift near the optimal ratio (2.0–2.5). Especially if the O/F ratio shift is changed in the direction of increasing ratio, the nozzle throat erosion would be even worse. Therefore, the O/F ratio shift has been considered to be one of the major research topics in the HRE research community.

Karabeyoglu et al. (2014) [169] proposed the decrease in propulsion efficiency, due to the “O/F ratio shift” effect inherent to single circular port HREs, is not as large as it is believed in the field. A similar point of view was also put forward in the study of Wei et al. (2019) [168]. A slight O/F ratio shift or even erosion of the nozzle throat does not significantly affect the engine performance. Figure 18 shows the vacuum Oxidizer I_sp, defined as the thrust divided by the oxidizer mass flow rate instead of the propellant mass flow rate, as a function of O/F ratios for various kinds of oxidizers with fuel HTPB, in which the Oxidizer I_sp stays roughly the same as the O/F ratio, which is less than ~7, except for the combination of O₂ and HTPB. Through the definition of Oxidizer I_sp, it can be understood that the performance of oxidizer control thrust is almost the same in the lower O/F ratio range (e.g., <~7), which means that, even if the nozzle throat is eroded or the O/F ratio shifts, the engine thrust can still be controlled by throttling the oxidizer mass flow rate. It is noted that the Oxidizer I_sp decreases rapidly once the O/F ratio is larger than 1~2 when oxygen is used as the oxidizer, even though the Oxidizer I_sp is higher than the other cases that use high-concentration hydrogen peroxide. It will make the thrust control easy by simply throttling the oxidizer mass flow rate like the hydrogen peroxide, as shown in Figure 18.

The propellant residue is another problem of HRE. Ozawa et al. (2017) [170] proposed the propellant residue that is caused by the O/F ratio shift is another major factor that reduces flight performance, in addition to engine performance deterioration. They also evaluated the flight performances of the O/F ratio controlled and uncontrolled hybrid sounding rockets by flight simulations. The research clarified that the O/F ratio shift elimination technologies in HRE propulsion could improve flight performance at most by 10% in the presence of systematic error and by 5.2% in the presence of random error.

For the O/F ratio shift issue, Barato et al. (2021) [171] summarize two major techniques that have been proposed to directly control the O/F ratio in an HRE. They included the aft injection of the oxidizer (AOIM) and the altering-intensity swirled injection (A-SOFT). Ozawa et al. (2023) [172] further proposed a real-time fuel regression measurement method of multi-material-additive-manufactured solid fuel with an integrated sensor probe structure. And through this method, combined with the fuel mass flow rate control technique, there is an opportunity to achieve more precise closed-loop control of thrust and a mixture ratio of hybrid rockets. In actual space propulsion missions, real-time fuel regression measurement technology can also be used for real-time residual fuel monitoring. The same concept of measuring the instantaneous regression rate also includes an X-ray translucent casing (XTC) hybrid rocket engine (Evans et al., 2005) [173]. On the other hand, in the O/F ratio of an uncontrolled HRE, Figure 19 (Li et al., 2023) [166] shows the linear correlation of the total mass of the consumed oxidizer with the total consumption of the fuel grain, considering the different throttling steps. They proposed that the engine’s total impulse and fuel usage are both proportional to the total oxidizer mass consumption in the throttling range of 50–100% for all the test cases presented in their study. This major finding would have the opportunity to solve the propellant residue problem and, indeed, broaden the potential application of HRE to future space technology.

3.4. Low-Frequency Instability

HRE combustion frequently shows low-frequency instability (LFI hereafter) with a peak frequency of 10–20 Hz. It is currently speculated that the cause is related to multiple interactions among many complex physical processes, such as vortex shedding, boundary-layer oscillation, thermal lag characteristics of solid fuel, heat-transfer oscillations, and the design of the post-combustion chamber. And it is often accompanied by the phenomenon of pressure resonance (Karabeyoglu et al., 1999) [174] (Cantwell et al., 2010) [133] (Park and Lee, 2015) [175] (Messineo et al., 2019) [176]. Combustion instabilities lead to acceleration fluctuations in HRE propulsion systems, which are generally not catastrophic, but the LFI may cause resonance problems in the structural or control systems. This may limit the use of HRE as the choice of launch vehicles. This phenomenon would be more pronounced in liquefying-fuel HRE, such as the use of pure paraffin as the solid-fuel grain. Because of the combustion gas flow over the fuel surface, which induces the Kelvin–Helmholtz instability (KHI), liquid-fuel droplets would be affected by the periodic entrainment effect, which is combined with pressure oscillations by the thermal lag of solid fuel and could be suddenly amplified (Kobald et al., 2015) [177]. Therefore, many studies have explored on this topic (Petrarolo et al., 2018) [178]. For HP-based HREs that often use the catalyst bed, it is also necessary to consider the instability caused by the improper design of the catalytic bed (Jo et al., 2011) [179].

The proper orthogonal decomposition (POD) is a technique using visual image analysis, which has been used in diverse areas of research. The application of POD enables the recognition of the main structures of the flow field and the combustion flame appearing in the video data, which includes spatial and temporal analysis of the structures. Therefore, it has also been used to analyze the LFI of HRE combustion recently (Kobald et al., 2015) [177]. Through this technique, many research results have shown that the use of swirling injection, different fuel insert designs, and stepped grain design can greatly resolve the LFI issue (Lee et al., 2022) [121] (Bellomo et al., 2013) [150] (Kim, and Lee, 2019) [180,181] (Kim, and Lee, 2020) [182] (Hyun et al., 2021) [183].

3.5. Ignitor Design

The ignitor is one of the important parts of the HRE, which is used to start the engine. Based on the different oxidizers, various ignitors were designed. For instance, the catalyst bed is widely applied for HP-based HREs (Wei et al., 2022) [48]. The benefit of a catalytic ignitor provides the ignition process with less sensitivity to the environmental temperature. Thus, the catalyst-based ignitor could be considered as the upper-stage rocket engine ignitor. Morlan et al. (1999) [184] also investigated different catalyst materials, in which the catalyst bed showed an ignition delay, with an order of 10 ms. Moreover, the results pointed out that the high decomposition rate was still active, even after 900 s of H₂O₂ operation, and that indicates the re-ignition capability of the catalyst-based ignitor design.

Ignition using hypergolic propellants is also a widely used method, which generally combines two different chemicals that create a fast reaction with a very low activation energy. Traditionally, toxic and potentially carcinogenic elements were applied for early applications, such as hydrazine reacting with nitrogen tetroxide (NTO) or rocket-grade H₂O₂ (RGHP hereafter). In recent studies, the non-toxic hypergolic miscible fuels (NHMFs) class (developed by the US Navy) offers green hypergolic propellant solutions (Dennis and Clubb, 2016) [185]. In the study, several types of hypergolic formulas were studied with a series of hot-fire tests with a thrust level of 25 lbs. The numerical analyses of the hypergolic reactions were also investigated in the early years. For example, Lawver et al. (1967) [186] studied the pressure spike due to the engines’ starting translates and random pressure pulses during operation, and the results served as the guide to the engine design. The numerical analysis of the hypergolic (tetroxide-UDMH/N₂H₄) was developed by Chuen et al. (1966) [187] to study the hypergolic reaction working at low-pressure and in a low-temperature environment.

Two major types of hypergolic tests are studied nowadays, namely, type A liquid oxidizer with liquid-phase hypergolic fuel, and type B, the combination of a liquid oxidizer and a solid-phase hypergolic fuel. The first type of hypergolic combination is widely applied in rocketry, such as SpaceX’s Dragon 2 Spaceship. Also, the fundamental studies of the type A hypergolic reaction were investigated by many academics around the world, which are described next.

Bhosale et al. (2020) [188] developed a series of type A studies of, e.g., reactions between rocket-grade H₂O₂ and sodium iodide-added fuel. Fundamental tests were performed to find out the ignition delay time in different test conditions. The bipropellant rocket engine was also developed and tested for verification. Kang et al. (2019) [189] indicated the design of the engine that has dual-mode operation from the bipropellant mode to the monopropellant mode, which has the potential of deep throttling. On the other hand, the hard-start phenomenon of the hypergolic propellant during the hot-fire test was also studied to prevent the structural damage found in the other research of their group (Kang et al., 2017) [190].

Recently, type B hypergolics attracted much attention due to their safety and for nontoxic reasons. Kwon et al. (2022) [191] developed the ammonia borane-based (BH₆N) hypergolic ignitor to ignite the H₂O₂-based hybrid rocket engine. Furthermore, the fundamental fuel properties were studied experimentally by several teams (Castaneda et al., 2019) [192]. The Israel Institute of Technology teams set up the experiments to figure out the performance of metal-hydride-based hypergolic solid fuel. Nath et al. (2022) [193] observed the ignition delay of the solid-fuel additive with the metal-hydride by a high-speed camera, showing that the droplet contact velocity strongly affected the ignition performance.

The plasma-type ignitor and laser ignitor were recently developed by many teams. Utah State University has recently developed an arc-ignition ignitor (Whitmore et al., 2017) [194], creating the arc by high voltage and pyrolysis of the ABS fuel grain to provide the combustion conditions. The plasma torch was also applied to the ignitor technique (Lee et al., 2015) [195], and a new steam plasma igniter was designed and tested in the study. Dyrda et al. (2020) [196] have developed a diode laser-based ignition technique and presented the theory of the ignition mechanism, as well as several observations regarding the effects of oxidizer velocity and incident laser power on ignition delay.

3.6. Scaling-Up Methods

To develop large-scale sounding rockets or space-launch vehicle systems by using HRE propulsion, a practical and proper scale-up method plays an important role in designing the engine configuration, as well as the propellant recipe for a given mission. Most important of all is that small-scale tests can be used to infer the performance of a large-scale engine. Much research has been performed to investigate these effects on the fuel regression rate at different thrust scales, such as fuel surface pyrolysis, gas-phase combustion and mixing, temperature gradient, and heat flux on the fuel surface. The most direct and practical theory so far for explaining regression rate behavior was developed by Marxman et al. (1965) [197] (1967) [198], which indicates that oxidizer mass flux is the fundamental factor governing the rate of fuel consumption in a typical HRE. The averaged regression rate can be simplified as:

$$\dot{r}=a{G}_{ox}^n$$

(9)

where ṙ, a, Gox, and n are the fuel regression rate, the constant pre-factor, the oxidizer mass flux, and the constant exponent, respectively.

Cantwell et al. (2010) [133] have further verified the theory through more than 40 hot-fire tests, from small-scale to large-scale, that the regression rate behaviors are relatively consistent within a range of chamber pressures and oxidizer mass fluxes. In addition, Yun et al. (2021) [199] verified the scale-up procedure from single-port solid fuel to multiport solid fuel based on this theory. Cai et al. (2013) [200] further pointed out that the solid-fuel port size is a critical scale factor that needs to be considered in the scale-up process, even though it may not be as important as the oxidizer mass flux. The reason was that the temperature gradient would decrease when increasing the fuel port and then affect the fuel regression rate. The averaged empirical regression rate law can be modified as:

$$\dot{r}=a{G}_{ox}^n{d}^{-0.2}$$

(10)

where d is the port diameter.

In summary, HRE combustion involves many complex physical and chemical phenomena. There are preliminary basic guidelines for scale-up, but the current research only focuses on the design of axial oxidizer injection. Different oxidizer injection methods, such as swirling injection, deserve further investigation to ensure that the same scale-up methodology can be applied.

4. Advanced Studies of Hybrid Rocket Engines and Applications

4.1. Throttling Capacity of Hybrid Rocket Engine

The throttling capacity plays an important role in extending the applications of HRE in practical launch vehicles. For LREs, the difficulty of throttling originates from induced combustion instability that would influence its performance, especially during deep throttling (Kim et al., 2020) [201] (Casiano et al., 2010) [202]. Compared with the typical LRE, HREs have a simpler system and more potential for throttling abilities (Vergez, 1998) [203]. Thus, it is crucial for HREs to demonstrate their throttling capacity before further space transportation applications in the new space age (Xiao et al., 2021) [204]. According to a recent study based on the theoretical derivation of thrust generation, the throttling capacity of HRE could be derived from Newton’s third law of motion and the calculation of the NASA CEA [166]. The sea-level thrust was derived not only as a linear function of the chamber pressure but also as the oxidizer mass flow rate while using a nozzle with erosion. These observations also are in agreement with the findings in a previous study (Zhao et al., 2018) [205]. Even though there is importance in throttling, there have not been too many studies that demonstrated the throttling of an HRE, which are described next.

As one of the few teams dedicated to HRE throttling research, Whitmore et al. (2014) [60] used an industrial ball valve with a flow coefficient of 0.7 and a servo motor to control the mass flow rate of N₂O to throttle a laboratory-scale HRE. They first measured and correlated the relationship between the flow area and the opening percentage of the ball valve. Afterward, the valve’s opening percentage was used as an open-loop controller to throttle the N₂O mass flow rate. The thrust was successfully throttled and was stably adjusted from 800 to approximately 12 N, and the resulting highly deep throttling achieved 1.5%. Compared with the throttling capability of LREs (Casiano et al., 2010) [202], an HRE is easier to deep throttle than a nominal LRE for stable combustion. Later, they used the thrust data and chamber pressure data as feedback to make a closed-loop PI controller for N₂O mass flow rate control (Whitmore et al., 2014) [61]. The mean thrust accuracy was reduced to a value of as surprisingly low as 3.9% during the tests. However, the mass flow rate of N₂O is too hard to instantly estimate due to its supercritical fluid properties, which the flow temperature would decrease and lead to a density change during feeding. These would also decrease the throttling accuracy.

Recently, H₂O₂, unlike the supercritical flow properties mentioned above, has attracted more attention to throttling because of the advantages of its incompressible flow properties. Zhao et al. (2015) [206] used 90% wt H₂O₂/PE to demonstrate the capability of throttling, even deeply down to 20%. They used a calibrated cavitating venture with a controllable area to control the oxidizer mass flow rate. The position of the pintle stroke was pre-tested to obtain the transfer function of the mass flow rate and considered as the feedback during hot-fire tests. The chamber pressure and the thrust were both found to be highly linear to the oxidizer mass flow rate. Ruffin et al. (2018) [207] presented their results of throttling an HRE using H₂O₂ and HDPE, which used a pintle stroke driven by a step motor with an encoder to control the oxidizer mass flow rate. And it also obtained the various throttling scenarios under different feeding-tank pressures to extend its throttling capacity. Tests using both sinusoidal and impulsive as command profiles had also been demonstrated in their research. Later, Ruffin et al. (2022) [208] further developed a real-time controller with a specially manufactured car-gas-pedal-like mechanism to instantaneously control the oxidizer mass flow rate to simultaneously throttle the thrust. However, the proposed throttling method based on the pintle stroke or cavitating venturi would cause a considerable pressure loss in the propellant feed system when a large oxidizer mass flow rate is needed.

Recently, Li et al. (2023) [166] proposed a throttle valve system consisting of an industrial ball valve, a DC motor, and a calibrated in-house differential flowmeter as feedback to control the instant mass flow rate of H₂O₂ precisely. The thrust was found to be highly linear with respect to the H₂O₂ mass flow rate. The thrust was stably throttled, with a thrust uncertainty of less than 5%. The control bandwidth of this throttle valve system was found to be higher than 5 Hz. And the same team applied this throttle valve system in a first-ever hovering flight test using an HRE (Wei et al., 2022) [48], which further demonstrated its potential in future complex space applications.

4.2. Thrust Vector Control for Hybrid Rocket Engine

The ability of thrust vector control (TVC hereafter) is another one of the most critical technologies for any rocket propulsion system as a launch vehicle. It decides the rocket’s attitude and trajectory during the flight. Figure 20 summarizes the current common TVC systems. Traditionally, LRE uses a movable platform design equipped with a gimbal mechanism and a corresponding actuator to drive the thruster for vectoring because of the small size of the engine (Wang et al., 2016) [209] (Sackheim, 2013) [210]. Many advanced designs of TVC have also been proposed since the early days (Penchuk and Croopnick, 1983) [211] (Redmill et al., 1994) [212]. However, this kind of traditional TVC mechanism is not appropriate for SRM and HRE due to both having a relatively large combustion chamber, which would require a tremendous torque for moving the engine responsively and a heavier system weight for the vehicle. Thus, some alternative TVC designs should be otherwise used.

To overcome the above-mentioned obstacles for HRE in TVC, one approach to reduce the influence is to attach the actuator directly to the approximate mass center of the chamber, named direct-drive (DD) TVC. Wei et al. (2022) [48] not only successfully precisely drove the thrust vector of their HRE, but also accomplished a hovering flight test using the HRE, the HTTP-3AT. However, the proposed mechanism still has mass and size issues when it scales-up to larger space vehicles, which deserves further investigation.

The secondary injection TVC (SITVC hereafter) is another alternative approach to TVC (Marshall et al., 2019) [101] (Zmijanovic et al., 2014) [213]. The SITVC injects a fluid stream, whether gas or liquid, at the divergent part of a nozzle. This injection would induce shock waves along the nozzle wall and result in an asymmetric pressure distribution inside the supersonic fluid field (Green and JR., 1963) [214] (Zeamer, 1977) [215]. The resulting momentum distribution due to the sided injection would generate a side force that could change the overall vector of the thrust plume. One of the advantages is that the SITVC is simpler than the mechanical actuator-driven TVC, in which there is no need to apply the torque required for the moving parts related to the engine for the former. The only thing needed is to decide on a proper position for the injection, as well as the mass flow rate control. Many experimental and numerical investigations, normally for LREs, focusing on this SITVC have been proposed for many decades (Waithe and Deere, 2003) [216] (Chen and Liao, 2020) [217], but there seemed to be no real flight test so far.

Lee et al. (2019) [218] used 98 wt-% H₂O₂ to demonstrate the liquid-injection TVC (LITVC) in an H₂O₂/HDPE HRE. Variations of injected locations had been tested with a specific fabricated graphic nozzle. The corresponding thrust vector changes were not only visualized but also measured quantitatively. However, the development of LITVC for HREs is still challenging due to its changing combustion characteristics found in the hot-fire experiments.

Other TVC designs that can be considered employed similar methods of SRM and LRE, such as the Stewart platform (Miloš et al., 2015) [219], flexible nozzle joint (Berdoyes, 2006) [220] (Yağmur et al., 2022) [221], ball-and-socket joint (Larkin and Singh, 2011) [222], jet vanes (Yağmur et al., 2022) [221], rotatable offset nozzle, supersonic split line (Orbekk, 2006) [223], mechanical deflection design (Yağmur et al., 2022) [221], vernier thruster (Soyuz rocket), and differentially throttling (the Soviet N-1 rocket), to name a few. Among these, except for the design of flexible nozzle joints and jet vanes and the vernier thruster type, the relevant technology has not been tested for a long burning time or is just a theoretical derivation of the design, and further research is needed.

4.3. Insulation Material for Chamber Casing and Nozzle Material

The HRE combines the advantages and disadvantages of both the SRM and the LRE. Combustion chambers of HRE share similar mechanical designs to the SRM. Insulation materials developed for the solid motor can mostly be adopted for the HRE. Although they both share a similar combustion chamber design, there is one fundamental difference between HREs and SRMs. HREs use an oxidizer injection system, and SRMs premix the oxidizer and the fuel, resulting in different types of combustion. In the combustion chamber of an HRE, the solid fuel breaks down into a gaseous state and mixes with the gasified oxidizer, resulting in a typical diffusion flame.

4.3.1. Temperature Distribution

Reactive gases generated by a pre-mixed flame in a solid motor can be treated as homogeneous temperature distribution. However, a diffusion flame has relatively poor combustion efficiency, since the fuel and oxidizer diffuse into the reaction zone from opposite directions through a complex convective flow in the combustion port. Consequently, the regression rates at different locations of the combustion chamber may become different because of the different local heat fluxes.

4.3.2. Efficiency Enhancement Structure

Different internal structures, rather than a single and straight port in an HRE combustion chamber, were developed to enhance the combustion efficiency. Designs such as the additional pre- and post-combustion chambers were presented for better combustion efficiency (Mechentel, and Cantwell, 2013) [224]. This resulted in direct exposure to hot gas when the engine started burning. Other designs, like star-port fuel, diaphragm (Tian et al., 2013) [119], and the innovative injection method (Gomes et al., 2015) [225], were also developed for better mixing of the fuel and oxidizer in the combustion chamber. However, these designs cause a strongly non-homogeneous flow and, thus, non-uniform regression acting at certain places of thermal insulation material.

4.3.3. Oxidizer and Fuel Distribution

The CFD results showed that the distribution of concentrations of oxidizer and fuel could be strongly non-uniform in the combustion chamber of an HRE (Lazzarin et al., 2015) [226]. For example, it is often observed that the chamber could be extremely fuel lean in the pre-combustion chamber, while it is highly fuel rich in the post-combust chamber. Mechentel et al. (2020) [227] used a single-port PMMA fuel grain and a high-resolution camera to measure the space–time-dependent data of the fuel regression profile. This new method further improves our fundamental understanding of hybrid combustion.

4.3.4. Residual Fuel

Due to different fuel regression rates along the fuel grain surface during combustion, residual fuel is a common problem for HREs. Thus, HREs usually carry more fuel than needed, preventing fuel grain burn by causing serious O/F ratio shifts and other problems. Nevertheless, the remaining fuel grain may be used as flame retardant material and provide some insulation effect.

The materials for internal insulation of the HRE combustion chamber are different than those for the nozzle. Insulation materials do not need to endure high shear stress from high-speed gas flow, and high temperature and high pressure are the challenges instead. Natli et al. (2016) [228] reviewed fifty years of research on different thermal protective systems (TPS), including heat shields for hypersonic flight. In the review, non-ablative and ablative materials were discussed in detail. Among all the ablative materials, polymeric ablatives (PAs) were specifically emphasized. As compared to non-polymeric insulation, such as high melting point metal and ceramic, PAs have some distinct advantages, such as tunable density, lower cost, and higher heat shock resistance, among others.

Due to the oxidizer-rich environment that often occurs in the HRE combustion chamber, oxidation-corrosion-resisting ablative materials should be developed, according to Kuo and Chiaverini (2007) [10]. Milhomem et al. (2017) [229] evaluated some substrates of lab-scale HREs by many small-scale and short-duration tests. Thus, the materials used for large engines and long burn duration can be determined. To increase HRE’s compatibility against LRE and SRM, commercial products, like epoxies, polyesters, polyurethanes, and silicones, can be applied. They also found that silicone resin was more thermally stable than phenolic resin. Ablative silicone rubber, with fillers such as anti-oxidation inorganic materials and carbon fiber (Kim et al., 2021) [230], may become a good choice for HRE insulation purposes.

Vaka et al. (2021) [231] proposed an HRE as a realistic test bench for insulation materials screening. Due to the engine’s repeatability of output and easy operation, HRE hot fires can produce high pressure and high shear force flow with high heat flux, which the oxy-acetylene torch test cannot achieve as a comparison.

Barato et al. (2020) [232] and Barato (2021) [233] developed a one-dimensional code that can solve the transient heat equation with or without regression of the surface proposed and raises some key challenges for insulation materials, including the design of the TPS for a restart and more difficulty in managing the heat soak back of solid fuel for a long shutdown, especially the requirement of Hohmann transfer for the upper stage.

Over the years, there have been many studies on HRE nozzle throat materials. Kahraman et al. (2020) [234] showed that operating pressure, O/F ratio, fuel formulation, and nozzle material were very significant parameters in the nozzle erosion rate, and it was a key parameter to predict and stabilize the engine thrust. Whitmore et al. (2021) [235] also conducted in-depth experimental research on special nozzle materials, which used a pyrolytic graphite throat insert with a carbon-reinforced epoxy composite insulation material design to pull heat away from the throat into a high heat-capacity insulating layer and exhibited a fivefold decrease in erosion rate than the monolithic graphite nozzle. Gallo et al. (2024) [236] investigated the reliability and feasibility of a regenerative cooling system by using a liquid oxidizer in the carbon-based nozzle throat of the hybrid rocket engine.

To explore and accurately predict the complex influencing mechanisms and key influencing factors of hybrid rocket combustion, O/F shifts, and throttling on nozzle erosion, Rotondi et al. (2024) [237] established an empirical predictive model for the erosion of carbon-based nozzle throats.

However, there are very few studies on the insulation materials in HRE propulsion, and many other aspects of HRE have yet to be fully explored, such as, for example, different combinations of propellant, O/F ratio, chamber pressure, burn duration, and oxidizer flux, to name a few.

With the advancement of HRE technology, the uneven erosion of insulation materials within large-scale engines with prolonged burn duration has become an increasingly important issue to be addressed. This topic should be prioritized as a key area of research in the near future.

4.4. 3D Metal Printing for Complicated Chamber Design

Metal additive manufacturing (AM hereafter), also known as 3D metal printing, is a process of creating three-dimensional parts by building them up layer by layer using a digital file as a guide. This contrasts with traditional subtractive manufacturing methods, such as turning, boring, milling, and drilling, which involve removing material to create a final product. Additive manufacturing allows for higher design freedom and the ability to create complex geometries and internal structures that would be difficult or impossible otherwise to produce using traditional manufacturing methods. Normally, annealing after the AM process is often required to relieve the residual stress in the solid part. In addition, surface machining/grinding may be also required, depending on the application requirements. There are several different types of metal additive manufacturing technologies, which are described next.

4.4.1. Powder Bed Fusion

This process uses a laser or electron beam to melt and fuse small particles of metal powder with sizes on the order of tens of micrometers together to create a solid part (Kerstens et al., 2021) [238]. Powder bed fusion is one of the most popular systems in the current metal AM industry. In the year 2020, powder bed fusion accounted for more than 50% of the market (Vafadar et al., 2021) [239]. Famous rocket engines like the ABL Space System’s E2 engine [240], Rocket Lab’s Rutherford engine, and SpaceX’s SuperDraco thruster (Lee et al., 2020) [241], to name a few, were all manufactured by the power bed fusion system.

4.4.2. Directed Energy Deposition

This process uses a laser, electron beam, or plasma arc to melt metal wire or powder as it is deposited onto a substrate, building up the part layer by layer. Relativity Space uses the wire-fed system to build the TERRAN 1 rocket’s fuselage, simplifying the manufacturing process. The powder-fed system was used by NASA to manufacture a diverged section of the rocket nozzle with complex internal cooling structures integrated into their study (Gradl et al., 2020) [242], which is otherwise very difficult to manufacture with traditional approaches.

4.4.3. Binder Jetting (Li et al., 2020) [243]

This process utilizes a binder to bind together small particles of metal powder, creating a solid part layer by layer. After the printing process, sintering is usually needed. During the process, non-homogenous shrinkage to the part can be hard to predict, which must be taken into account at the design stage. Parts requiring high tolerances may take advantage of traditional metal manufacturing methods to machine out functional features.

4.4.4. Fused Filament Fabrication (FFF)

This process forms parts via fused deposition modeling (FDM), in which they undergo a debinding and sintering process. The filament used a high-percent metal powder content combined with the binder. This makes the filament properties similar to a regular polymer-based filament, which can be used in regular FDM 3D printers. Just like binder jetting, non-homogenous shrinkage to the solid part can be hard to predict during sintering (Wagner et al., 2022) [244].

4.4.5. Sheet Lamination (Zhang et al., 2018) [245]

This process uses sheets of metal that are bonded together, layer by layer, using a laser, ultrasonic, and/or pressure to create a solid part. This method is currently rarely used in practice.

To take full advantage of these benefits of metal AM and create parts with good quality, it is important to understand every possible design approach to create an AM part. One needs to clearly define the design requirements before the design process. It includes determining the necessary strength, durability, and performance criteria for the solid part. Weight reduction methods, like topology optimization, iso-grid structure, bimetallic/multi-metallic design, and integrated design such as regeneratively cooled combustion chamber, nozzle, and plumbing, can be applied.

Topology optimization is especially useful for creating lightweight and efficient structures. This method determines the most efficient use of the material in a structure, while still meeting the strength and performance requirements. It can be applied to a wide range of engineering problems, including structural design, heat transfer, fluid flow, and electromagnetic design. Airbus uses AM to manufacture a lightweight wing bracket (Kellner, 2017) [246]. A great difference between traditional machining and AM can be observed. The AM part has several small trusses that are difficult or otherwise impossible to produce by a traditional five-axis milling machine. The AM part, on the other hand, can build up complex shapes layer by layer, therefore having little limitation to produce such a complex part.

A similar bracket design can be seen integrated into the Rutherford engine’s combustion chamber, which attaches a TVC actuator to provide steering capability for the rocket (Rocket Lab) [247]. Integrating parts together can not only reduce the complexity of the final product by reducing number of the parts and assembly time but also can almost eliminate the weight of mechanical joints, such as bolts and flanges. In the application of rocket engines, the manufacturing of complex piping and regenerating cooling channels can also be greatly simplified by AM (Lee et al., 2020) [241]. The Rapid Analysis and Manufacturing Propulsion Technology (RAMPT hereafter) Project at NASA employed blown powder=directed energy deposition to produce integrated channel-wall structures in the expansion section of a liquid rocket nozzle, which is a potential method to create a lighter structure (Clark and Tyler) [248].

Selecting the right material is also crucial for creating a good part. Factors such as strength, durability, and thermal properties must be considered when selecting the appropriate material. It is important to select a material that is compatible with the printing technique being used. Under NASA’s RAMPT Project (Lee et al., 2020) [241], an AM bimetallic combustion chamber was developed. A GRCop-84 copper-alloy liner printed by the laser powder bed fusion method was combined with an Inconel 625 structural jacket by cold gas spray, blown powder-directed energy deposition, and laser hot-wire cladding. The bimetallic design takes advantage of copper alloy’s high thermal conductivity and the mechanical strength of Inconel 625 under high temperatures, resulting in a lightweight combustion chamber. In addition to a bimetallic chamber, a composite overwrap can also serve the rule of the structural jacket.

An iso-grid is a structural design that uses a repeating pattern to create a lightweight and stiff structure. When combined with 3D metal printing, an iso-grid structure can be printed with high precision and minimal material waste on a complex surface. For example, a spherical pressure vessel with an iso-grid made by 3D metal printing was presented in Ossola’s research (Ossola et al., 2021) [249]. The results show that an iso-grid layout could provide up to 14% mass reduction, compared to plain shells. Furthermore, the study showed that AM can provide significant cost (−26%) and mass (−79%) savings, compared to traditional subtractive manufacturing.

In brief summary, the field of 3D metal printing has a considerable advantage for improving topologically complex problems and various lightweight designs, which are often used in the design of liquid rocket engines. In the field of HRE propulsion systems, it is more commonly used in oxidizer injector design (Musker et al., 2018) [250] and liquid-cooled nozzle design (Reaction Dynamics) [29] (Quigley and Lyne, 2014) [251] (Ercole et al., 2017) [252].

4.5. Recent Status of Hybrid Sounding Rockets and Launch Vehicles

As mentioned earlier, HRE offers superior throttling and re-ignition abilities, in addition to being safer and more cost-effective than LRE or even SRM (Estey and Whittinghill, 2017) [253]. As a result, they have become increasingly popular in recent years for use in sounding rockets, small satellite launchers, space-travel vehicles, and landers, among others.

Hybrid sounding rockets and suborbital launchers have been successfully launched by various organizations. Norway’s Nammo launched the world’s first hybrid rocket to cross the Karman line (Faenza et al., 2019) [47]. The Nucleus rocket, using a combination of H₂O₂ and HTPB as the propellant, reached a height of 107.4 km while carrying a 62-kg payload. The Advanced Rocket Research Center (ARRC hereafter) at Taiwan’s National Yang Ming Chiao Tung University launched the second stage of their HTTP-3A rocket [31], which features four hybrid rocket engines with active guidance and control for the first time in the community. Australia’s Gilmour has demonstrated their Rasta rocket (Cecil and Majdalani, 2016) [254] using a 3D-printed fuel grain and nitrous oxide, as the first step towards their goal of a low-cost satellite launcher. Recently, they have been preparing to launch an orbital rocket in 2024 by using hybrid rocket engines with hydrogen peroxide and a solid fuel grain. The HyEnD project at Germany’s University of Stuttgart has also presented several hybrid rockets using N₂O as the oxidizer (Kobald et al., 2017) [255] (Kobald et al., 2019) [54] (Oechsle et al., 2022) [256], and the N₂ORTH sounding rocket reached an altitude of 64.4 km in 2023 (Hybrid Engine Development, 2023) [257]. Also, in Germany, HyImpulse [258] launched the SR75 sounding rocket for technology qualification for their Small Launcher. Poland’s Warsaw Institute of Aviation launched the ILR-33 Amber 2K in 2024 (Okninski et al., 2021) [50] (Parsonson, 2024) [51], a hybrid rocket with two solid rocket boosters. The Hybrid Sounding Rocket Program (HSRP) at the University of KwaZulu-Natal from South Africa has also completed successful flight tests of their hybrid rockets using paraffin-based fuel with a 20% aluminum additive (Genevieve et al., 2015) [259]. The Aerospace and Compressible Flow Research group at the University of Calgary, from Canada, has also launched its own hybrid rocket (Messinger et al., 2019) [260].

Since the American Rocket Company (AMROC) in the USA prototyped the world’s first HRE propulsion satellite launch vehicle in 1990 (Kniffen et al., 1990) [261], many commercial launchers have been planning to use HRE as their primary propulsion system. For example, Gilmour in Australia [8] has qualified their Sirius HRE for installation on the Eris rocket, which is set to launch in the fourth quarter of 2024. Innospace in South Korea [6] has also completed a hot-fire test of their HyPER-15 HRE, which can provide 150 kN of thrust, and successfully launched the first test launch vehicle HANBIT-TLV from Brazil in 2023. HyImpulse in Germany is currently developing a launch vehicle with a 75-kN thrust HRE (Schmierer et al., 2019) [262]. Additionally, ATspace (formerly TiSpace) in Australia [27] plans to build a 3-stage launch vehicle capable of carrying a 350 kg payload to orbit.

In addition, HREs have also been used in space travel mainly because of their safety. Virgin Galactic’s SpaceShipOne spacecraft was powered by an HRE using N₂O and HTPB (Kelly et al., 2017) [52] (Thicksten et al., 2008) [53]. It was the first private suborbital flight to carry a pilot across the Karman line, launched from its White Knight mothership. In 2019, Virgin Galactic [23] launched its spacecraft, Virgin SpaceShip Unity, to an altitude of 90 km with a crew of three, including one passenger. In 2021, they further launched to 86.1 km, and they had been certified by the FAA to provide commercial spaceflight travel.

The throttling capability of HREs makes them a suitable choice for landers and ascent vehicles in planet exploration missions. In 2020, the ARRC in Taiwan successfully completed a first-ever hovering flight test of its HTTP-3AT rocket, which uses four throttle-able HREs (Wei et al., 2022) [48] with simple industrial ball valves, demonstrating a 10 s hovering at an altitude of two meters during its 25 s flight. Nammo in Norway has also verified the throttling technology for their hybrid rocket lander in the SPARTAN program (Sotto et al., 2015) [263], which has four engines with a fixed canted angle, even though it was not quite successful at that time. In the Mars Sample Return Campaign, NASA (Story et al., 2020) [264] has planned to use a single-stage-to-orbit hybrid rocket as the ascent vehicle and has completed several full-scale hot-fire tests of its main engines. Additionally, Delta V Space Technologies [32] is also developing a hybrid rocket lunar lander to send a rover to the moon. All these developments showed the potential space applications of the HRE, mainly because of its excellent throttling capability.

5. Conclusions

In summary, an extensive review of the development of hybrid rocket propulsion-related technologies in the past few decades is conducted in this paper. As hybrid rocket propulsion continues to advance and breakthrough, more and more new space teams are looking to adopt hybrid rocket propulsion technology. Even though the I_sp performance of the engine is lower than that of current, more advanced, liquid rocket engines, and low cost-effectiveness, simplicity, reliability, low complexity, and safety are still the primary conditions for mass access to space in this field. Especially except for the reused rocket, this may be a key to decreasing the launch costs for small launch vehicles. Not to mention that hybrid rockets also have the potential to be recovered and relaunched. Through the investigation and invention of various technical fields of hybrid rockets, many traditional problems of HRE have been preliminarily solved. Although there is still no best solution, this article could provide some major guidelines and tips for HRE researchers in further research. More importantly, even though HRE technology has grown rapidly in the past two decades, some research demonstrated HRE applications by demonstrating its advantages and capabilities, such as throttling, hovering, and sounding rocket flight demonstrations. Nevertheless, it is yet to be fully verified in its use in space transportation compared to SRM and LRE. Further investigations, research, and even the demonstrations of upcoming launch-vehicle flight tests using HREs are crucial for future applications in space transportation.