The Renaissance of Hybrid Rocket Engines: Advancing Small Satellite Propulsion

Authors: Victor Yevlakhov & Maya Ambrozhevich

Over the past two decades, there has been a resurgence of interest in hybrid rocket engines (HREs), in which the fuel and oxidizer exist in different states of matter. In the first blog of our HRE series, we explored their history, strengths and weaknesses, and the niche they occupy within the rocket engine family. HREs fit into the ‘neutral zone’ or ‘border area’ between liquid and solid propellant engines. This concept applies to all aspects of HREs, including design, manufacturing technology, cost, operating parameters, control and regulation systems, specific impulse, thrust, and engine operating time.

Fig. 1. A sharp increase in the number of small satellites launched into orbit over the past decade [1]

The renewed interest in HREs has been driven by the rapid increase in the number of small satellites launched into orbit (up to 500 kg in Low Earth Orbit) (Fig. 1). The engines used for such satellites typically fall between solid-fuel boosters and high-thrust liquid engines. This has prompted many to consider hybrid engines as a viable option for small satellite propulsion – an idea that has led to exciting projects, such as the one with hybrid propulsion engines described in detail in [1, 2] (Fig. 2).

Cycle Features

As hybrid rocket engines gain traction, companies worldwide are developing innovative solutions to maximize efficiency and performance. Let’s focus on one potential configuration, as shown in Figure 2.

Any propulsion system can operate in several modes, but two are constant: the start-up mode and the steady-state mode. Focusing on the steady-state mode, it involves two feed systems operating in parallel: the displacement system and the turbopump system (TPU). In the displacement system, helium (He) pressurizes the liquid ethanol (EtOH) tank. This ethanol is oxidized by liquid oxygen after being pumped through the Gas Generator (GG). The combustion products from the ethanol fuel drive a turbine (T) that powers the pump (P). After passing through the turbine, the combustion products enter a heat exchanger (HX), where they heat and evaporate the liquid oxygen (LOX) before it enters the oxidizer tank. The combustion products then expand in the converging-diverging (CD) nozzle, creating additional thrust, which can also be used for control. The main engines in this setup are hybrid rocket engines (HREs), where the fuel is a solid propellant based on polybutadiene (HTPB), and the oxidizer is LOX after the pump (P) [2].

Fig. 2. A novel small launcher project with dual hybrid marching engines [2]

Fig. 3. Hybrid Rocket Engine Modeling in AxSTREAM System Simulation software [2]

Key Performance Parameters of HRE Cycles

Regarding the thermodynamic cycle arrangement, a few key points must be addressed. The gas generator typically operates with an excess of oxidizer (or fuel) to maintain the required temperature for the gases entering the turbine (Fig. 3). While it is theoretically possible to burn the gas generated after the heat exchanger (HX) in an additional combustion chamber before expanding it in the CD nozzle, this was not implemented due to the small flow rate of the generator gas. Adding another combustion chamber would increase system complexity, weight, and reduce reliability. This option can be explored using CAE programs, such as AxSTREAM System Simulation. HREs operate with an Oxidizer-to-Fuel Ratio (OFR) close to stoichiometric (Fig. 7a), at which the maximum specific impulse of thrust is achieved.

From the thermodynamic design and analysis point of view, there are key parameters that should be precisely considered and evaluated.

• Turbopump Unit (TPU) Characteristics: The temperature of the combustion products entering the turbine inlet after the gas generator (GG) defines the characteristics of the turbopump unit. The temperature of the combustion products at the gas generator outlet, based on the OFR, is shown in Fig. 4. This is influenced by the strength characteristics of the turbine blade material, with temperatures ranging from 1000 to 1200 K. The higher the temperature, the better for the cycle.

Fig. 4. The dependence of methanol Combustion Products’ Temperature

• Engine Thrust: Engine thrust is a critical characteristic of the propulsion system. Thrust is influenced by the flow rates, including the mass flow rate of oxygen in the oxidizer tank pressurization system and the oxygen mass flow rate at the gas generator inlet. This is also important for the engine’s control system.

Fig. 5. Dependence of Gross Thrust on oxygen mass flow rate of in the Oxidizer Tank pressurization system and oxygen mass flow rate at the Gas Generator inlet

• Complexity of the Propulsion System: Even a simplified representation of the propulsion system is quite complex, as changes in one parameter can affect others across the cycle. For example, the pressure at the pump outlet influences pump power, the thermal power of the heat exchanger, engine specific impulse, and even the temperature of the combustion products exiting the auxiliary nozzle located behind the heat exchanger.

Fig. 6. Dependence of some parameters on the output pressure of the Oxygen Pump.

• Combustion Temperature, Specific Impulse, and Total Thrust Trends: It’s important to monitor how combustion temperature, specific impulse, and total thrust evolve throughout the system.

Fig. 7. a) The maximum specific impulse and the maximum combustion temperature in the combustor correspond to different OFRs; b) The maximum specific impulse and maximum thrust of the propulsion system correspond to different mass flow rates of fuel

Conclusions

Designing and analyzing HRE cycles requires a flexible approach to modeling and evaluating under different boundary conditions and configurations. Using AxSTREAM System Simulation software has confirmed the fundamental operability of the propulsion system and provided key operating characteristics for the system’s main elements. The 0D thermodynamic and 1D thermal-fluid network approach allows for different simulation levels, starting from preliminary thermodynamic design and extending to co-simulation analysis that accounts for accurate modeling of both systems and subsystems.

Interested in learning more about how AxSTREAM System Simulation can help your rocket engine design process? Get in touch by requesting a software trial here:

SOURCES:

This study is based on a publicly available research project performed by HyImpulse. Please find the original study linked below

1. https://digitalcommons.usu.edu/cgi/viewcontent.cgi?article=5035&context=smallsat
2. https://digitalcommons.usu.edu/cgi/viewcontent.cgi?article=4819&context=smallsat