guest blogger john hare
John Bossard over at Plasma Wind http://plasmawind.typepad.com/ introduced me to air turborockets after a propulsion conference in 2000. My hard references on the subject were from his recommendations. He has built operating engines on this cycle. Never being willing to let well enough alone, I suggest a few upgrades. He may take the ideas apart if he has time and interest, and it doesn’t conflict with his business.
A normal turborocket is an air breathing engine with a rocket gas generator driving a turbine which drives the air compressor. The fuel rich turbine exhaust mixes with the compressed air to burn in the afterburner. The result is an engine with over twice the thrust to weight of a jet engine with immunity to flameouts. It can reach higher altitudes and higher airspeeds than any turbojet, and is simpler to operate. The down side is that it sucks down far more fuel than a jet, and won’t operate in vacuum like a rocket.
The turborocket is a niche engine for air breathing acceleration from mach zero to mach five or so. It can actually do the things people try to claim a ramjet can do. It is also good for relatively short duration cruise when engine mass is very critical and the good thrust to weight ratio outweighs the high fuel consumption. It also uses far less fuel than a rocket when in its’ proper working environment.
A rocket may have an Isp (Dense Fuels) of 300 with a T/W of 100. A turbojet may have an Isp of 3,000 with a T/W of 10. A turborocket might have an Isp (dense fuels) of 750 with a T/W of 20. Clearly the turborocket needs either a better Isp, or better T/W to make it a clear win over the more traditional engine cycles. For this particular cycle, hydrogen makes a lot of sense.
Hydrogen gives a 30-40% Isp boost to rockets, at the expense of tripling propellant tankage per unit of propellant mass, when H2/O2 is compared to Kero/O2. It is an ongoing argument whether LH2 is good for launch vehicles. Upper stages benefit far more from the extra performance. Hydrogen on air breathing engines though, exhibit a 300% increase in Isp.
Hydrogen has about 3 times kerosenes energy per pound when burned with atmospheric air, so even being 10 times as bulky per pound, it only has tanks a bit over 3 times the size of the kerosene per BTU. Since acceleration engines have to lift the take off mass, cutting your fuel load by two thirds has to have an effect on engine mass, along with wing and landing gear mass of course.
Hydrogen has another valuable trait in its’ specific heat characteristics. It is the best possible coolant with about 16 times the cooling capacity of air. 16 pounds of air introduced to 1 pound of colder hydrogen will meet it about halfway on temperature. Hydrogen can cool enormous quantities of air either through a physical heat exchanger, or with Mass Injection Pre Cooling (MIPC). Many studies have used precooling in some form, though most of them just get too complicated for the simple functions we need for a launch system.
I suspect a straight switch to hydrogen from dense fuels will less than double Isp, considering the oxydizer mass to be carried. I suggest using the cooling properties of liquid hydrogen to precool a core flow for a turborocket. This potentially increases mass flow through the gas generator while much decreasing the on board oxydizer requirements.
After the initial compression of air, about a sixth of the total flow is sucked into the core for gas generator use. A pound of LH2 as injection precooling for each sixteen pounds of air per second will almost double the density compared to uncooled air that has been through the same compressor. True gas generators operating in ejector mode compress the air/hydrogen mixture and burn it in the middle chamber that drives the turbine. The combination of precooling and ejector compression should give a pressure in the middle chamber of about four times that of the main chamber. With a turbine pressure ratio of four, and a core mass flow of 20% of total compression, a fairly high main compressor ratio (for turborockets) should be achievable.
The turbine can be cooled by the remaining hydrogen required to reach stoichiometric burn in the main chamber. I suspect an Isp in the 2,000 range can be reached during acceleration periods, with cruise phase considerably higher. This while retaining most of the strengths of the turborocket system.