Concept and objectives

Hybrid propulsion

The basic operation of a chemical rocket engine is the reaction of two (or more) propellants in an open volume. The heat released causes an expansion of gases through a nozzle, resulting in a force on the exhaust which propels the rocket forward. Depending on the propellant they use, they can be distinguished in three general classes: bi-liquids, solids and hybrids. A hybrid system, as its name suggests, uses propellants in two different phases: one solid and the other either gaseous or liquid.

The schematic view of classical hybrid propulsion operation is depicted in Figure 1. The liquid propellant is stored in the propellant tank and pressurized with an inert gas (e.g. Helium) stored in the pressurant tank. The desired amount of liquid propellant is vaporized on to the solid propellant through regulation of a flow valve. Once in contact, the propellants ignite and the products of combustion exhaust through the nozzle, providing the intended thrust.



Figure 1: schematic operation of hybrid rocket engines (Source: Hokkaido University)

In the case of HYPROGEO, the liquid oxidizer is hydrogen peroxide H2O2, at a very high concentration grade. It is gasified by flowing through a catalyst bed and then burns with the pyrolysis gases coming from the ablation of the solid propellant. The hot combustion gases are passing through a nozzle to produce thrust.

Hybrid propulsion has numerous advantages over classical types of propulsions: it is as efficient, but cheaper and greener. “Greener” means that the propellant is environmental friendly, easier and safer to handle than conventional ones. Despite these advantages, it suffers from specific issues (unburnt fuel, regression of fuel grain, combustion oscillation, thrust stability, etc.), that have limited the development of large thrust applications (i.e. first/second stage of large launch systems). All these issues are all addressed by HYPROGEO, which aims for long and stable firing time. Consequently, HYPROGEO will enable new applications, or a true renaissance for hybrid propulsion. 

Table 1: advantages of hybrid propulsion over solid and bi-liquid rockets. (Source: Space Propulsion Gropu Inc.)

State of the art

Hybrid propulsion is not a new technology. The first successfully launched hybrid engine flew on the Soviet GIRD-09, a sounding rocket that used bi-liquid oxygen to burn gelled petroleum, on the 17th of August 1933. However, most of the development were done much later throughout the 60s until the 80s at several large aerospace companies and research centres. In Europe, the main actors were the ONERA, Nord Aviation (today Astrium Defence and Space), and Volvo.

Today, the use of hybrid rocket engines is mostly limited to low and moderate thrust applications: sub-orbital flight such as sounding rockets and tourism. A famous example is the SpaceShipOne, the first private manned spacecraft, which was powered by SpaceDev's hybrid rocket motor (Figure 2).

Figure 2: similarly to SpaceShip One, SpaceShip Two uses hybrid propulsion. Here a flight test of its rubber-fuelled rocket engine. (Credits: Virgin Galactic)

Used in conjunction with electric propulsion (EP), hybrid propulsion offers the advantages of:

  • Equivalent or higher performance (ISP and thrust) than bi-liquid technology.
  • Simpler design (compared to bi-liquid and to electric propulsion), with one single tank.
  • Shorter transfer time with higher thrust (reduced cost of operations compared to EP). The time from launch to operations is often a key factor for commercial missions and hence can be a limiting factor for EP transfers (despite them being more fuel efficient). The shorter transfer results also in reduced radiation exposure from fewer transitions through the van Allen belts.
  • Shorter and simpler filling operations before launch (safer and reduced cost of operations compared to traditional chemical propulsion).
  • Globally, the main impact will be a much lower cost: 30% savings are expected on a future operational system. Last but not least, the proposed hybrid propulsion technology is green: the chemical reaction produces CO2 and H2O.

A very innovative exploitation derived from this R&D could be a detachable apogee module for the communication satellites (satcom) market (Figure 3). When combined to current or future European launchers and to the emerging EP technology on satellites, such a modular approach will increase the diversity of achievable mission profiles and improve the competitiveness of European access to space. The impact of this R&D project is thus two-fold:

  1. Pave the way for a new pillar of the independent and cost-effective access to space in Europe, with a new generation of upper stage or transfer module.
  2. Contribute to the structuration of the related industrial supply chain in Europe.

Figure 3: Conceptual view of one of the disruptive innovations from HYPROGEO; a transfer module using hybrid propulsion. The toroid is the interface with the launcher's upper stage. The tanks are the blue spheres on the side, and the nozzle is the pruple extrusion.

Challenges of the project

The strategic objective of HYPROGEO is declined in two main operational objectives:

  1. Study a new cost-effective subsystem for access to space using hybrid chemical propulsion technology while confirming its benefits as a complementary and cost-effective solution compared to traditional chemical engines and full electric propulsion.
  2. Provide significant scientific progress and increase the European excellence and know-how in critical technologies and components of an operational hybrid propulsion system:
  • the combustion chamber;
  • the high endurance nozzle;
  • the catalytic injector;
  • the production, storage and use of high concentration hydrogen peroxide propellant.

These four elements will influence the performance of the system, and scientific research is required in order to propose the development of an operational system. Let us now briefly cover the specific technical challenges of each element.

The combustion chamber

The objective is to propose a combustion chamber design for long duration burns and stable/constant low to moderate thrust in vacuum. Indeed, with current regression rate, the typical channelled type of combustion chamber design envisaged for rocket applications would lead to a large and long combustion chamber to achieve the desirable thrust level, thus incompatible with the highly demanding spacecraft accommodation constraints. Moreover, as opposed to solid and bi-liquid technologies, the amount of ablated fuel evolving during the firing causes instability the oxidizer to fuel mixture that impacts directly the propulsive performance through the ISP level and the combustion efficiency. This is acceptable for short duration firings of rocket applications but not for the targeted long duration transfer for which thrust and combustion efficiency have to remain constant.

The high endurance nozzle

The main challenge is to design, manufacture and test a lightweight nozzle capable of withstanding the specific erosion caused by the long term exposure of the reaction products of the hybrid engine. Ceramic C/C-SiC based composite materials are good candidates.

The catalytic injector

Even though the hybrid chemical engine can operate with a bi-liquid atomizer for the injection of the oxidizer in the combustion chamber, the combustion efficiency is greatly improved when a catalytic injector is used to pre-decompose the oxidizer into a hot gaseous phase that enhances the mixture with the pyrolysis gas from the solid fuel grain. Moreover, the use of catalytic injector avoids the necessity of a dedicated pyrotechnic igniter to start the engine. That allows engine “cold start”, and increases the engine reliability for the multiple re-ignitions during the transfer phase. Thus the challenge is to design a catalytic injector compatible to highly concentrated H2O2 and working at high temperature capable of rapid ignition.

The high concentration H2O2

The H2O2 concentration strongly influences the specific impulse (ISP). The target is thus to use the highest concentration possible, well above the 87.5% grade currently available on the market in order to achieve similar performance to bi-liquids technology. However, although low level TRL  techniques have been tested to produce H2O2 at concentration higher than 97%, none of them have yet been developed to large scale production. Also, within current European transportation regulations there is no approved supply chain for high test peroxide > 97 wt% established; and none or only few instruments and equipment systems are approved for the use with high test peroxide > 97 wt%. As a consequence, the biggest challenge for the use of high test peroxide in the space industry at concentrations above 87.5% is the accessibility to an industrial scale production and implementation of a setup regarding safe and reliable supply chain. 

Structure of the project

These R&D activities are structured by 4 main work packages. A system study ensures the global vision in coherence with an economic analysis, the identification of technical challenges and the consolidation of scientific results. A last work package performs the dissemination of results.

Figure 4: Overall work logic of HYPROGEO.

The Project

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