Liquid propellant engines remains the mainstream power plant in most purpose built space launch vehicles.
At Copenhagen Suborbitals we locked our choice of main engine about a year ago. This happened after a series of static tests with medium and large liquid propellant rocket engines using LOX and 75 % ethyl alcohol as propellant.
These propellants are cheap, non toxic, easy avalible by the ton, and environmental safe. Since CS is committed to sea launch, there are additional requirements to the handling of the propellants. One of them is that it must be fairly safe and environmentally manageable if propellants are lost at sea. A LOX spill at sea is not a problem at all, and a spill of ethyl alcohol in the sea is much less a problem than a spill of any hydrocarbon fuel as, say, RP-1. Furthermore, alcohol cannot burn on the surface, because it is diluted by the sea water. The auxiliary propellant – 80 % hydrogen peroxide – is also relatively benign if lost at sea. So the propellant choice is more than just density and specific impulse, it takes into consideration even the unlikely situation that a large amount of propellants is lost at sea.
As was the case in the early days of spaceflight, the cooling properties of water / alcohol mixtures are far better than those of hydrocarbon fuels. This is another important factor in not using hydrocarbons.
Read more about our newest BPM-5 rocket engine here
In the case of our old workhorse, the 65 cm diameter HEAT X, we use a simple pressure feed propulsion system. In this system high pressure helium is used to expel LOX, and high pressure nitrogen is used to expel the alcohol fuel. The system uses a pressure blow down cycle, so the tanks are filled only 65 % – the remainder being an ullage containing the pressurization gas. The helium is stored in the tank at LOX temperature -183 C.
CS liquid propellant rocket engines is based on the first generation LPRE, as known from the V2 missile, and – more precisely from the US Army´s PGM-11 Redstone missile.
This means that we operate at relatively low pressures, and at relatively low thermal loads on the combustion chambers. This in turn means that simple low carbon steel can be used as the primary construction material for the combustion chambers. The combustion chambers are both regeneratively and film cooled, as was the case of the V2 engine and Redstones North American A6 rocket engine.
The “TM65 Tordenskjold” rocket engine comes in a 45 kN form with a pressure blow down feed system. This version is used on the HEAT X workhorse.
The main feature of the pressure blow down system is the build-in simplicity. However, to build such a system on a large scale means manufacturing propellants tanks on an similarly large scale. The 65 cm HEAT X workhorse launch vehicle represents a scale where the manufacturing of the tanks, and the amount of expensive helium gas used, is still manageable. However, the HEAT X rocket is not powerful enough to launch a manned capsule into space. To do so, CS need to develop a larger booster that can handle a payload with a diameter of 1,6 meters.
This is the HEAT 1600.
In the early design phase of this vehicle, it was considered to make a simple scale up of the HEAT X. It was found that the cost of the manufacturing of the propellant tanks, and to no small extent, the cost of the helium gas to be used, both in welding and as pressurisation media, was prohibitively high.
So, what would it really take to build a crude, low pressure turbine pump ? We started to do the math, and explored standard pump options for conversion to turbopump use. This research showed that, at the size of the HEAT 1600, a crude turbine pump would be a feasible solution. It was decided early on to use the historic V2 / Redstone system with a gas generator cycle powered by 80 % hydrogen peroxide.
The use of a turbine pump versus a pressure blow down system is controversial for amateurs. However, the manufacturing of the tanks for low pressure was not only far cheaper and simpler, it also removed some 1700 kg of weight from the tank system. 1700 kg that could go into the turbine plant, and still be competitive.
The cost of a compact, high performance, high speed turbopump is high. The cost of a bulky, low pressure, crude turbopump based on off the shelf pump components is likely less.
We designed a pilot scale plant for the making of 80 % hydrogen peroxide, and is currently developing the turbine pump plant for the TM65 baseline rocket engine. It will weigh far less then 1700 kg, and the first tests has already been done.
The launch vehicle that uses the turbo rocket engine is descibed in another section on this site. The TM65 is not powerful enough, even in the turbo edition, to launch a fully fueled HEAT 1600, so a TM260 rated at 260 kN thrust is planned for its ultimate version.