Copyright 2014 Robert Clark
In the blog post A SpaceX Heavy Lift Methane Rocket I discussed some possibilities for a heavy lift rocket using SpaceX's planned methane engine, the Raptor. Those calculations were based on the initial released values for its thrust of 660,000 lbs, 300,000 kilograms-force. However, recently SpaceX has said the Raptor may have a thrust of 1,800,000 lbs in vacuum, 1,600,000 lbs at sea level. So I'll give some revised estimates for its payload.
As before I suggest using the same tooling for the core stage as that used for the SLS core to save on development costs. Corresponding to the new higher thrust of the Raptor I'll use the full tank size of the SLS core, which would hold 1,000 metric tons (MT) of hydrogen-lox propellant. Since methane-lox is 2.4 times as dense as hydrogen-lox, the SpaceX methane-lox core will hold 2,400 mT of propellant. The relative densities of methane-lox compared to hydrogen-lox are given in Table 1 of this report:
Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996
In regards to the specifications of the methane rocket engine it should be noted the high vacuum Isp of 380 s cited really would only be expected of a vacuum optimized nozzle. Such engines can not be operated at sea level. However, in an upcoming blog post I'll discuss some altitude compensation methods that will allow the engine to have this vacuum optimized Isp while still being able to operate at sea level. So I'll assume both the first stage and upper stage engines have the vacuum Isp of 380 s.
As before I'll take the number of engines for the core as five and the propellant-to-dry mass ratio as 20 to 1. For the upper stage I'll take the propellant size as approx. 1/5th that of the core stage, at 500,000 kg with the same mass ratio, and use a single Raptor for the upper stage.
SSTO Case.
We'll use again Dr. John Schilling's Launch Vehicle Performance Calculator. Input the thrust as the value of the vacuum thrust in kilonewtons of 5 Raptors as 5*8200 kN = 41,000 kN, and the Isp as the vacuum Isp of 380 s. Input the propellant mass as 2,400,000 kg and the dry mass as 1/20th of this at 120,000 kg. For the "Default Propellant Residuals?" option select "Yes", and for the "Restartable Upper Stage?" option select "No". Selecting "Yes" for this last option would reduce the calculated payload.
| Launch Vehicle: | User-Defined Launch Vehicle |
|---|---|
| Launch Site: | Cape Canaveral / KSC |
| Destination Orbit: | 185 x 185 km, 28 deg |
| Estimated Payload: | 110395 kg |
| 95% Confidence Interval: | 75079 - 152433 kg |
So this larger version could give 110,000 kg payload as an SSTO.
Two Stage Case
For the two stage case, in the column for the 2nd stage, input the propellant mass as 500,000 kg and the dry mass as 25,000 kg. Input the thrust as the vacuum thrust of 8,200 kN, and the Isp as the vacuum Isp of 380 s. Then the calculator gives the result:
| Launch Vehicle: | User-Defined Launch Vehicle |
|---|---|
| Launch Site: | Cape Canaveral / KSC |
| Destination Orbit: | 185 x 185 km, 28 deg |
| Estimated Payload: | 196089 kg |
| 95% Confidence Interval: | 162727 - 236334 kg |
Close to 200 metric tons payload to LEO. Actually since the mass ratio of stages improve as you scale them up, quite likely the mass ratio will be better than just ca. 20 to 1, perhaps in the range of 25 to 30 to 1. This would then improve our payload to above 200 metric tons.
Cross-Feed Fueling for Multiple Cores.
We'll just look at the 3 core case here. As described in the A SpaceX Heavy Lift Methane Rocket post, we emulate cross-feed fueling in the Schilling calculator by inputting for the two side boosters 2/3rds the actual propellant mass, and also increase the first stage propellant mass by an additional 2/3rds. The rest of the specifications, dry mass, thrust, Isp remain the same.
So enter 2 as the number of side boosters. Then in the column for the boosters, input 1,600,000 kg for the propellant mass and 120,000 kg for the dry mass of the boosters. Enter the actual vacuum thrust of 41,000 kN and Isp as 380 s. Then the calculator gives the result:
| Launch Vehicle: | User-Defined Launch Vehicle |
|---|---|
| Launch Site: | Cape Canaveral / KSC |
| Destination Orbit: | 185 x 185 km, 28 deg |
| Estimated Payload: | 536113 kg |
| 95% Confidence Interval: | 449806 - 639526 kg |
Over 500 metric tons of payload! This rocket however would be truly massive at 3 times the mass of the Saturn V. This would likely require a new,expensive launch pad to handle a rocket this size.
Bob Clark
UPDATE, October 25, 2014:
Some methods of accomplishing the altitude compensation mentioned are discussed here:
Altitude compensation attachments for standard rocket engines, and applications.
http://exoscientist.blogspot.com/2014/10/altitude-compensation-attachments-for.html
UPDATE, October 25, 2014:
Some methods of accomplishing the altitude compensation mentioned are discussed here:
Altitude compensation attachments for standard rocket engines, and applications.
http://exoscientist.blogspot.com/2014/10/altitude-compensation-attachments-for.html
