BACKGROUND

Reaching space is hard. Lots of energy needs to be released to do so. Since the start of the space era, rocket engines have been producing this energy by chemical reaction known as combustion. Combustion requires two chemical reactants, an oxidizer and a fuel, called propellants. The heat carried by the combustion byproducts is turned into kinetic energy by letting the hot and pressurized gas to expand through the nozzle.

To put payload into orbit, one must use a large quantity of propellants. And every bit of mass must be useful to obtain an efficient rocket. This constraint led to design rockets with two or more stages so that the structural mass that became useless because the tanks have been emptied, can be jettisoned.

Currently, space transportation industry is dominated by two propulsion technologies: the liquid propellant engine and the solid propellant engine. Both of these technologies have their own advantages but they come with tradeoffs and drawbacks. They are far from ideal propulsion systems.

A third propulsion technology exists, called hybrid. Hybrid Rocket Engines have the potential of featuring the advantages of both liquid and solid propulsion technologies. They could become the best propulsion technology for space transportation in the near future!

Adapted from : Fundamentals of Hybrid Rocket Combustion and Propulsion - Chiaverini, M. I. and Kuo, K. K.

HYBRID ROCKET ENGINE

A HRE consists in using one liquid propellant (usually the oxidizer) and one solid propellant (usually the fuel). The liquid oxidizer is stored into a tank and feeds the combustion chamber which contains the solid fuel grain.

Using HRE, thrust can be modulated by controlling the oxidizer flow rate. Similarly, the engine can be shut down and reignited. Standard HRE have similar specific impulse than common kerolox engine. Solid fuel grain eases the use of metallic additives which can lead to propulsion systems featuring greater specific impulse than H2/O2. HRE is easy to develop and cheap to manufacture due to the low complexity and number of parts. It is also safe to operate. Even in case of failure, the propellants cannot be mixed to undergo violent explosion. This engine is robust and less sensitive to operational condition.

However, to date, you will not see HRE flying payload into orbit… That is due to two factors.

The first is economical. Most of the space propulsion developments take their roots during the Cold War when they received institutional funding. Still today, space propulsion development programs are expensive to run. Therefore, one prefers to use mature and proven technologies when it comes to starting a new rocket development.

The second is technological. Hybrid Rocket Engines face from a technological barriers: at large size (required by high thrust applications such as orbital rockets) classical HREs suffer from low solid fuel regression rate and inefficient combustion. In large HRE combustion chambers, the solid fuel does not receive enough energy from the flame and does not vaporizes fast enough for the thrust requirements. Besides, the gaseous reactants of the combustion do not mix properly resulting in incomplete combustion and inefficient engine.

At Hybrid Propulsion for Space, we are developing a breakthrough technology to unlock the technological hurdle. By providing to the space industry, ready-to-use high thrust HRE, we are removing all the barriers between the benefits and advantages of this new propulsion technology and space launchers sector.

OUR INNOVATION

The cofounders of Hybrid Propulsion for Space have developed a special injector that ensure the flame stays very close to the solid fuel surface and drive the turbulences at any given combustion chamber size. It has the effect of increasing the solid fuel regression rate (the speed at which the solid fuel is consumed) while ensuring that all the reactants are thoroughly mixed.

We designed, built and operated our first demonstrator called Lily. It is a combustion chamber test article, meaning it was not intended to produce any thrust, just to study the combustion process. We used gaseous oxygen up to 12 bars (1.2 MPa) combined with Poly (Methyl MethAcrylate) or PMMA (known as Plexiglas®) as both fuel and pressure containment. PMMA was used for testing purposes only, because it is transparent which allows to see directly into the combustion chamber. This has been widely done by the community.

We conducted more than 20 test firings from which we gained a lot of experience of running a test bench. Some firings were conducted without our innovation providing a baseline of performance which compares well with scientific data found in literature with similar setup. Other firings were done with prototypes of injector and the results showed up to 70 % of improvement of the fuel regression rate (denoted as ṙ in the figure below). With these encouraging results, the company was founded in May 2019 and a patent (still pending) application has been filed.

R&D ROADMAP

Our next step is to use a second demonstrator called Joker, this time to obtain propulsive performance. It is currently designed to produce up to 150 kgf (1500 N) of thrust to stay below ICPE regulations. A first firing campaign will be done using a standard bell-shaped nozzle. A second campaign is intended to test an aerospike nozzle. We also envision to integrate this engine into a student sounding rocket (if we shall find one) for a first flying demonstration.

The full scale demonstration will be conducted with the Terminator test article. It will be designed to produce a nominal of 30 kN of thrust. A first firing campaign will be conducted to debug the test article by incrementally increasing the duration and the thrust level from 5 kN to its nominal value. This is done thanks to the thrust modulation ability of HRE (alike liquid propellant propulsion). The second firing campaign will be a fast cycle of testing and improving the last constructive parameters (optimization process). A third campaign will test the Thrust Vector Control (TVC) system.

The first application for the Terminator class thruster is a single stage suborbital rocket. This vehicle is primarily for flying demonstration purposes. It might also find a commercial application as a tool for conducting scientific and technology experiments in microgravity or space environment, or even for advertisement with dramatic scenery behind luxurious products.

The main goal is powering orbital class rockets with our HREs. Our market analysis of the space sector led us to envision a microlauncher with a lift capability of 250 kg of payload to Low Earth Orbit (LEO) or 190 kg to Sun Synchronous Orbit (SSO). The upper stage will require 30 kN of thrust (hence the thrust of Terminator test article). The booster will require roughly 7 times more thrust, around 200 kN. This can be quickly achieved by clustering 7 HREs at the first stage.

Ultimately, the microlauncher we envision will use a single HRE per stage because it is easier to control and should be more reliable than clustered HREs. This last version of the microlauncher will require a new HRE development which is expected to be ran while the first version of the microlauncher will be operational. The new HRE could be a prelude to a smalllauncher program using the same roadmap than the microlauncher.

Lastly, if the market should prove to be difficult to address with a microlauncher, a HRE with 30 kN of thrust can be quickly turned into the main engine for a nanolauncher (less than 50 kg of payload capability). It can also propel all sorts of other vehicles (satellites, drones, hypersonic gliders).

WHY USING HRE?

When the oxidizer tank is not filled, HRE are basically an assembly of metallic, composite and polymer parts. There is no associated hazards to manufacture the parts and to assemble the HRE in town. The manufactory can settle virtually anywhere (it should not require a SEVESO infrastructure). Transportation is also eased. The microlauncher we envision has the right size to enable transportation in a standard 20 foot container. So it can be shipped away from the manufacturing plant by road, rail or sea without special care required. It also should be transportable inside a cargo hold of large military transport aircraft such as an A400M, for military purposes.

Our market analysis highlighted the need for IP protection of the payload. For some payloads, crossing borders is a no go. For others the associated paperwork is deterring even if a foreigner launch operator is cheaper than a local one. Therefore, the choice of launch location is strategic for a launch operator.

Having a launch pad on the ground does not offer the flexibility required. Having several is costly and mildly more flexible. Airborne launches are costly to develop and the size of the carrier plane limit the size of the rocket that can be carried to the size of a microlauncher. Launching from sea carrier seems to be a good alternative since 70 % of the Earth is covered by water and any rocket size could be launched this way…

Thanks to their relative simplicity, HRE are robust. They can be operated in a wider range of conditions. They can be made salt water compatible and thus lifting-off directly from sea water surface. No need for a launching platform, just a buoy. For reusability purposes, since it can lift-off from sea water surface, it can also land onto it. No need for landing legs or a landing platform or even a pinpoint guiding and landing system. This way, it can be offered to each client a launch from international water reducing IP and other governmental exportation issues.