Propulsion Group

The two projects currently being studied include the development of a 40N high test peroxide (HTP)/kerosene rocket engine and the long-term storage of hydrogen peroxide. Both projects made considerable progress last year and two papers were presented at the green propellants conference at ESTEC last June. The bi-propellant rocket engine has progressed from an un-cooled heat sink chamber design to a practical gas-cooled system utilising flight propellants for cooling without any need for external cooling.

The performance of the engine has been calculated to be 90 per cent of theoretical, demonstrating good combustor efficiency, translating to an expected vacuum specific impulse (I_sp) of 260 seconds. This, when combined with the high specific gravity of H₂O₂, produces a density I_sp (DI_sp) that actually exceeds that of commercially-available small bi-propellant (N₂H₄/N₂O₄) systems, making it very attractive for small satellites.

The long-term storage work reached a conclusion on suitable tank materials for the containment of peroxide. Modern fluoropolymers, namely PFA and FEP, are good at containing peroxide but their intrinsic compatibility is not as good as pure aluminium. However, such polymers do permit oxygen to permeate and this could form the basis of a new tank design that is currently under investigation.

Hybrid rockets have characteristics that are well suited to small spacecraft missions. Environmental friendliness, unparalleled safety for a given level of performance and the promise of highly reduced propulsion system cost make this technology a contender for the upcoming generation of highly maneuverable and cost effective small spacecraft.

This research program focuses on the packaging of hybrid motors to accommodate small spacecraft designs. In addition to investigating techniques for employing short and squat axial flow hybrid motors, this program is investigating a novel "pancake" geometry motor that can be incorporated into a multifunctional separation system and thereby leverage some of the separation system mass into the rocket motors combustion chamber. The "pancake" design has the additional benefit of an external mount providing the best scenario for radiating heat to space.

One of the primary obstacles that has kept hybrid technology from being employed on spacecraft is their historically long and slender geometry. Long and slender geometry coupled with the requirement to mount the motor inline with the spacecraft centre of gravity presents a prohibitive combination to the small spacecraft designer. In addition, the heat generated by a hybrid rocket can present thermal difficulties especially when the motor and catalyst pack are embedded deep within a small spacecraft.

We have developed a novel configuration for a hybrid motor that fits within the volume constraints of a small satellite. Its "pancake" structure ensures good mixing of oxidizer and fuel within a confined space, making use of vortical flow within the chamber. The use of this unique hybrid technology provides a very high thrust propulsion system for our satellites.

The vortex flow pancake (VFP) hybrid engine completed its testing phase with the design of an engine capable of providing 800 m/s of delta-V on a 100 kg microsatellite. The engine demonstrated exceptional efficiency, with performance within 1 per cent of theoretical, unheard of for a hybrid rocket engine. The design also provided exceptionally smooth combustion. A paper on the VFP was presented at the Green Propellants Conference last June at ESTEC, attracting a lot of attention from various parties. Work on an unusual-geometry hybrid rocket was inspired by the tight volumetric constraints imposed by small spacecraft. As such it has engendered serious interest from mission designers who are grappling with severe volume constraints yet require relatively high performance propulsive capability.

Nitrous oxide (N2O) gas has been considered as a propellant for small satellite propulsion. An ability of self-sustaining exothermic decomposition, non-toxicity, storability, and self-pressurisation feature of this liquefied gas benefit towards its use for cold-gas, resistojet, monopropellant, and bipropellant propulsion. This gives the opportunity of combining any of these propulsion options into one simple, inexpensive system sharing the same propellant and capable of providing the spacecraft with all necessary propulsion functions to fulfil its mission.

We have developed a novel configuration for a hybrid motor that fits within the volume constraints of a small satellite. Its "pancake" structure ensures good mixing of oxidizer and fuel within a confined space, making use of vortical flow within the chamber. The use of this unique hybrid technology provides a very high thrust propulsion system for our satellites.

To develop monopropellant and bipropellant propulsion techniques, we investigated the possibility of using nitrous oxide. This propellant was investigated as it has the property of being able to generate a self-sustaining decomposition. This had been previously observed by Tim Lawrence during vacuum tests of one of his resistojets at Edwards U.S. Air Force Base in 1998.

Recent experience of storing nitrous oxide on-board the UoSAT-12 mini-satellite for more than one year indicates that storage of the gas in-orbit is not a problem.

We have studied the catalytic exothermic decomposition of nitrous oxide to evaluate the economic and power saving advantages of this propellant. This has encouraged us to consider the feasibility of developing a nitrous oxide monopropellant thruster for small satellite applications.

Self-sustaining catalytic decomposition has recently been demonstrated during sea-level test firing of prototype assembly of nitrous oxide monopropellant thruster at Westcott, Royal Ordinance test facility of the University of Surrey.

The vortex flow pancake (VFP) hybrid engine completed its testing phase with the design of an engine capable of providing 800 m/s of delta-V on a 100 kg microsatellite. The engine demonstrated exceptional efficiency, with performance within 1 per cent of theoretical, unheard of for a hybrid rocket engine. The design also provided exceptionally smooth combustion. A paper on the VFP was presented at the Green Propellants Conference last June at ESTEC, attracting a lot of attention from various parties. Work on an unusual-geometry hybrid rocket was inspired by the tight volumetric constraints imposed by small spacecraft. As such it has engendered serious interest from mission designers who are grappling with severe volume constraints yet require relatively high performance propulsive capability.

This research was supported by Surrey Space Centre and partially funded with research grants from the European Office of Aerospace Research and Development under Contract #994100.

Most propulsion systems rely on the use of combustion or electrical power input to impart energy to propellants, for thrust and specific impulse benefits. The sun, however, is a natural heat source, providing over a kilowatt per square meter in earth orbit. Through the use of a concentrating mirror with concentration ratios on the order of 1,000 to 10,000:1, we can focus intense sunlight onto a refractory metal or ceramic cavity filled with propellant and obtain very high performance propulsive capability. Isp figures of 400-450 seconds are believed to be achievable. Since a small satellite can only carry a small amount of fuel, this research provides an opportunity to put a small satellite into a deep space orbit, or escape the gravitational pull of the Earth entirely.

Preliminary analysis has confirmed that a micro-scale solar thermal engine, using a storable, stable propellant such as ammonia (NH₃), hydrazine (N₂H₄), or water, will be capable of moving microsatellites between low earth orbit and geosynchronous orbits, put into lunar orbit, or even interplanetary space. Velocity changes (delta-Vs) on the order of 1,000-3,000 m/s are realisable, at thrust levels of up to several Newtons. Low-cost mirror technologies, analogous to terrestrial telescope optics, and judicious selection of high-temperature cavity materials, will be critical to the cost-effectiveness of the design.

Research into this approach is being conducted in collaboration with The Boeing Company and the U.S. Air Force Research Laboratory, who will provide several key components and access to system test facilities. A proposal for a flight experiment, the Microscale Solar Propulsion Experiment (MSPEx), is currently in development and could fly aboard a Surrey microsatellite as early as May 2005.

Since 2001, the Surrey Space Centre (SSC) has undertaken a comprehensive investigation of microsatellite-based solar thermal propulsion, including likely mission applications, satellite and launch vehicle constraints, trajectory analyses, requirements definition, and design. The resulting microscale STP system is currently undergoing component level ground testing in preparation for flight opportunities arising after 2005.

Two solar cavity receivers (denoted Mk. I and Mk. II) have been built and electrically tested in vacuum at temperatures of up to 2,000 K. The 400-gram Mk. II cavity receiver, an insulated ceramic structure, has survived numerous cycles to 2,000 K with no sign of damage or deformation. This receiver is expected to provide roughly 500 N-s of impulse per engine firing, at a thrust level of between 1 and 5 N. Predicted specific impulse (Isp) with ammonia propellant is between 300 and 400 s. Test results have validated our extensive thermal modelling, with close agreement between simulation and actual test data. Receiver flow testing with inert gas, nitrogen, and ammonia propellants, in high-temperature vacuum, is underway. Thrust and Isp data, as well as material degradation and mass loss information, are being collected in a series of thermal cycling tests, intended to simulate a multiple-kick firing profile.

A purpose-built 56-cm diameter concentrating aluminium mirror has demonstrated high concentration ratio (> 10,000:1) and input powers of up to 150 W, at ground level test on-sun. This concentrator is now being used to characterize concentrator-receiver interactions, testing the full optical path. A number of small (14-cm) mirrors, diamond-turned from a plastic substrate, have also been produced and coated at very low cost. These demonstrate moderate concentration ratios (1,225 to 1,600:1), which could be improved with the addition of small hyperboloidal secondary lenses near the focal point, or more simply through the substitution of aluminium for plastic as the substrate material. Aluminium is susceptible to smaller form errors during machining, and thus allows for higher concentration ratios. Three 14-cm aluminium mirrors are being procured for further testing.

Testing with high numerical aperture optical fibre is underway. Recent success with low-loss high-flux solar radiation transmission over fibre lines has been reported by researchers in Israel, for surgical and nanomaterial production applications. SSC is now attempting to duplicate these results and apply them to solar thermal propulsion. Decoupling the solar receiver from the concentrating mirror's focal point permits a single receiver to accept light from multiple mirrors, which can be made smaller at much less expense. The mirror-fiber scheme is under test, with estimates of end-to-end efficiency approaching 75 per cent.

A space test hardware suite, the Microscale Solar Propulsion Experiment (MSPEx), is being proposed by the Surrey Space Centre for incorporation into a microsatellite slated for launch in the 2005/6 timeframe. Two 20-centimetre mirrors, coupled to a single small refractory ceramic receiver, supplies approximately 60 W of power for heating. Estimated thrust levels range from 20 to 100 milliNewtons. Predicted Isp could range as high as 390 s. If successful, this experiment would represent the first space-based test of solar thermal propulsion, and a stepping stone to larger operational systems.