Copyright 2014 Robert Clark
India is progressing towards manned spaceflight:
December 17, 2014 by William Graham
The current plan is for a 2021 launch for the manned system. However, the most recently developed rocket the GSLV Mk. III uses two large solid side boosters. Space engineers in general do not like solid rockets for manned launchers since they can not be shut down. However, there is an all liquid alternative for India for a manned launcher that actually would be cheaper than the GSLV Mk. III.
It would use 4 of the liquid-fueled strap on boosters used on the earlier design the GSLV Mk. II attached to the GSLV Mk. III core stage. We'll use the specifications on the GSLV Mk. II and GSLV Mk. III on Ed Kyle's SpaceLaunchReport.com page. The strap-ons for the GSLV Mk. II have a gross mass of 48.2 metric tons (mT) and propellant mass of 42.6 mT so a dry mass of 5.6 mT. The vacuum thrust of the single Vikas 2 engine is 70,360 kgf, 690 kN, with a vacuum Isp of 281 s.
The gross mass of the GSLV Mk. III is 125 mT with a propellant load of 110 mT, so a dry mass of 15 mT. The vacuum thrust of the two Vikas 2 used is 140,720 kgf, 1380 kN. The cryogenic upper stage has a gross mass of 30 mT, with a propellant load of 25 mT, so a dry mass of 5 mT. It's engine CE-20 engine has a thrust of 20,000 kgf, 196 kN, with an Isp of 450 s.
Input this data into Dr. John Schilling's Launch Performance Calculator. Select the Satish Dhawan launch site and a launch inclination of 13.9 degrees to match the latitude of the launch site. Then the calculator gives a payload to LEO of:
| Launch Vehicle: | User-Defined Launch Vehicle |
|---|---|
| Launch Site: | Satish Dhawan Space Center |
| Destination Orbit: | 185 x 185 km, 14 deg |
| Estimated Payload: | 6493 kg |
| 95% Confidence Interval: | 4829 - 8498 kg |
"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters
This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.
About 6,500 kg. It is notable though reading the description of the failed launches of the GSLV on the SpaceLaunchReport.com page that some involved engine failures. There would need to be multiple successful unmanned launches of this configuration before it is certified for manned launches.
Bob Clark
UPDATE, January 15, 2015:
I found that the www.spaceflight101.com site provides more accurate info on the GSLV Mk. II and GSLV Mk. III launchers than the Spacelaunchreport.com page. Using these values gives an even better performance for this proposed all-liquid launcher.
The GSLV Mk. II side boosters have specifications listed as:
UPDATE, January 15, 2015:
I found that the www.spaceflight101.com site provides more accurate info on the GSLV Mk. II and GSLV Mk. III launchers than the Spacelaunchreport.com page. Using these values gives an even better performance for this proposed all-liquid launcher.
The GSLV Mk. II side boosters have specifications listed as:
Boosters
| # Boosters | 4 | |
| Type | LH40 | |
| Length | 19.7m | |
| Diameter | 2.1m | |
| Inert Mass | ~5,600kg | |
| Launch Mass | 47,600kg | |
| Tank Material | Aluminium Alloy | |
| Fuel | UH25 - 75% UMDH, 25% Diazane | |
| Oxidizer | Nitrogen Tetroxide | |
| Propulsion | 1 Vikas 2 | |
| Thrust | 763kN | |
| Impulse | 293 sec | |
| Engine Dry Weight | 900kg | |
| Engine Length | 2.87m | |
| Engine Diameter | 0.99m | |
| Burn Time | 148sec | |
| Chamber Pressure | 58.5bar | |
| Mixture Ratio | 1.7 (Ox/Fuel) | |
| Attitude Control | Single-Plane Engine Gimbaling | |
| Stage Separation | With Core Stage |
The GSLV Mk. III core stage has specifications listed as:
Core Stage
| Type | L-110 | |
| Length | 21.26m | |
| Diameter | 4.0m | |
| Fuel | Unsymmetrical Dimethylhydrazine | |
| Oxidizer | Nitrogen Tetroxide | |
| Inert Mass | 10,600kg | |
| Propellant Mass | 115,000kg | |
| Launch Mass | 125,600kg | |
| Propellant Tanks | Aluminum Alloy | |
| Fuel | UH25 - 75% UDMH, 25% Diazane | |
| Oxidizer | Nitrogen Tetroxide | |
| Propulsion | 2 Vikas 2 | |
| Thrust (SL) | 677kN | |
| Thrust (Vac) | 766kN | |
| Specific Impulse | 293 sec | |
| Engine Dry Weight | 900kg | |
| Engine Length | 2.87m | |
| Engine Diameter | 0.99m | |
| Chamber Pressure | 58.5bar | |
| Mixture Ratio | 1.7 (Ox/Fuel) | |
| Turbopump Speed | 10,000rpm | |
| Flow Rate | 275kg/s | |
| Area Ratio | 13.88 | |
| Attitude Control | Engine Gimbaling | |
| Ignition | T+110s | |
| Burn Time | 200s | |
| Stage Separation | Active/Passive Collets |
The cryogenic upper stage has specifications listed as:
Cryogenic Upper Stage
| Type | C-25 Cryogenic Upper Stage | |
| Length | 13.32m | |
| Diameter | 4.0m | |
| Fuel | Liquid Hydrogen | |
| Oxidizer | Liquid Oxygen | |
| Inert Mass | ~4,000kg | |
| Propellant Mass | 25,000kg | |
| Launch Mass | ~29,000kg | |
| Propellant Tanks | Aluminum Alloy | |
| Propulsion | CE-20 | |
| Engine Type | Gas Generator | |
| Thrust - Vacuum | 200kN | |
| Operational Range | 180-220kN | |
| Specific Impulse Vac | 443s | |
| Engine Mass | 588kg | |
| Chamber Pressure | 60bar | |
| Mixture Ratio | 5.05 | |
| Area Ratio | 100 | |
| Thrust to Weight | 34.7 | |
| Burn Time | 580s | |
| Guidance | Inertial Platform, Closed Loop | |
| Attitude Control | 2 Vernier Engines | |
| Restart Capability | RCS for Coast Phase |
Plugging these dry mass, propellant mass, and Isp values into the Schilling calculator gives these results:
"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters
This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.
A payload of 9,000 kg is quite high. It's close to the 10,000 kg payload of the GSLV Mk. III without needing the two huge, expensive solid side boosters the Mk. III uses. Note also this all-liquid configuration with 4 liquid side boosters is similar to that of the venerable Soyuz launcher used for decades for manned launches.
To get a faster operational manned rocket we might use the smaller cryogenic upper stage already used on the GSLV Mk. II. On the December 2014 test flight of the GSLV Mk. III, its cryogenic stage was still in development and was not part of the test. So to use all operational stages instead we could use the Mk. II's cryogenic stage. It's specifications are listed as:
Inputting these specifications instead into the Schilling calculator gives:
"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters
This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.
The 6,800 kg payload is still high, sufficient for a manned launcher.
Bob Clark
| Launch Vehicle: | User-Defined Launch Vehicle |
|---|---|
| Launch Site: | Satish Dhawan Space Center |
| Destination Orbit: | 185 x 185 km, 14 deg |
| Estimated Payload: | 9005 kg |
| 95% Confidence Interval: | 7106 - 11299 kg |
"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters
This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.
A payload of 9,000 kg is quite high. It's close to the 10,000 kg payload of the GSLV Mk. III without needing the two huge, expensive solid side boosters the Mk. III uses. Note also this all-liquid configuration with 4 liquid side boosters is similar to that of the venerable Soyuz launcher used for decades for manned launches.
To get a faster operational manned rocket we might use the smaller cryogenic upper stage already used on the GSLV Mk. II. On the December 2014 test flight of the GSLV Mk. III, its cryogenic stage was still in development and was not part of the test. So to use all operational stages instead we could use the Mk. II's cryogenic stage. It's specifications are listed as:
Third Stage
| Type | GS3 - C15 | |
| Inert Mass | ~2,500kg | |
| Launch Mass | ~15,300kg | |
| Length | 8.7m | |
| Diameter | 2.8m | |
| Tank Material | Aluminium Alloy | |
| Fuel | Liquid Hydrogen | |
| Oxidizer | Liquid Oxygen | |
| Propellant Mass | 12,800kg | |
| Guidance | Inertial Platform, Closed-Loop | |
| Propulsion | 1 ICE (CE-7.5) | |
| Cycle | Staged Combustion | |
| Thrust (Vacuum) | 73.5 to 93.1kN | |
| Impulse | 454sec | |
| Engine Dry Weight | 435kg | |
| Engine Length | 2.14m | |
| Engine Diameter | 1.56m | |
| Burn Time | Up to 1,000sec | |
| Chamber Pressure | 58bar | |
| Attitude Control | 2 vernier Jets, each 2kN | |
| RCS for Coast Phases | ||
| Stage Separation | Merman Band, Hot Staging |
Inputting these specifications instead into the Schilling calculator gives:
| Launch Vehicle: | User-Defined Launch Vehicle |
|---|---|
| Launch Site: | Satish Dhawan Space Center |
| Destination Orbit: | 185 x 185 km, 14 deg |
| Estimated Payload: | 6810 kg |
| 95% Confidence Interval: | 5373 - 8577 kg |
"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters
This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.
The 6,800 kg payload is still high, sufficient for a manned launcher.
Bob Clark
2 comments:
There are basically two (credible) failure modes with solids as boosters (when designed and tested properly). (1) Seals can leak as in the Challenger disaster. (2) voids or cracks in the propellant cause overpressure-burst disasters. The third issue is poor motor design from a fundamentals standpoint, root of the AFRPL thrust stand explosion that caused a billion-dollar write-off for what was Hercules Aerospace in the 1990's.
Item (1) is a seal-and-joint design problem. Item (2) is a processing quality control problem. Simple as that, and just as complex, too. Solve those and there is no reason you cannot use solids in a manned system. They can be solved, but the solutions are unknown to government labs. Not all contractors can do these "right" either. Some cannot even do that third egregious item correctly. The SLS SRB instability is one of those.
GW
Large solids also suffer from vibration problems. These were mitigated on the space shuttle by the presence of the ET between the SRBs and the orbiter. However, the inline design of Ares I proved to be troublesome due to vibration issues.
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