Tuesday, December 30, 2014

A half-size Ariane for manned spaceflight, Page 2.

Copyright 2014 Robert Clark

  Whether Europe will build a manned launcher is not an engineering question. It's an economics and politics question. The European Union has the greatest economic might in the world as measured by GDP, including that of the U.S. It's greater than that of the space faring nations of Russia, China, and India combined. Moreover, Germany, France, Italy, and the UK, individually have greater economic power than India. Yet Europe has no plans to produce a man-rated vehicle. 


The currently accepted plan for the Ariane 6 is to cut down the size of the Ariane 5 core and spend funds on developing an upgraded version of the solid stages used on the Vega launcher to be used as side boosters. The ESA has agreed to spend $10 billion developing the Ariane 6 and upgrading the Vega to the Vega-C.

The half-size Ariane would have much smaller development cost since you would not need these new side boosters. Nor would you need the larger second stage using the Vinci engine. It would use the current cryogenic upper stage of the Ariane 5.

It would have a smaller payload than the Ariane 6 at about 4,800 kg for the two stage vehicle . But it would have about the same payload capacity of the upgraded Vega, the Vega-C. The Vega-C will have about 2 metric ton greater payload than the Vega, which will put it in the range of 4,500 kg.

The Vega already costs in the range of $50 million per launch. The cost of the Vega is in the range of $20,000 per kilo to orbit. The high cost probably deriving from its high development cost, in the range of $1 billion. The Vega-C needs an approx. 50% upgrade in size of the main solid stage, likely resulting in high additional development costs. Then judging by the approx. 50% upgrade in payload capacity, this would give it an estimated cost of $75 million.

In contrast, due to the low additional development needed for the half-size Ariane its cost would likely be comparable to other liquid fueled rockets in the range of $10,000 per kilo, or $48 million per launch.

Beyond that another very important advantage is that it could be made reusable if SpaceX succeeds in reusability. If SpaceX does succeed in cutting costs by reusability then the Vega and Vega-C immediately become obsolete. The half-size Ariane on the other hand would be able to keep pace with the price cuts by also being made reusable.

I mentioned the considerations on whether this could be undertaken were financial and political. The main financial reasons it should be undertaken are that it would have lower development cost than the Vega-C and would serve as a hedge against SpaceX succeeding in reusability.

Ironically, this might be the same reason why it might not be undertaken for political reasons, because it would undercut the Vega rocket, which is largely being built in Italy.


  Bob Clark

Sunday, December 28, 2014

A liquid-fueled Indian manned launcher. UPDATED.

Copyright 2014 Robert Clark


 India is progressing towards manned spaceflight:

India debuts GSLV Mk.III with prototype crew capsule.
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:

Mission Performance:
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:

                                 Boosters

# Boosters4
TypeLH40
Length19.7m
Diameter2.1m
Inert Mass~5,600kg
Launch Mass47,600kg
Tank MaterialAluminium Alloy
FuelUH25 - 75% UMDH, 25% Diazane
OxidizerNitrogen Tetroxide
Propulsion1 Vikas 2
Thrust763kN
Impulse293 sec
Engine Dry Weight900kg
Engine Length2.87m
Engine Diameter0.99m
Burn Time148sec
Chamber Pressure58.5bar
Mixture Ratio1.7 (Ox/Fuel)
Attitude ControlSingle-Plane Engine Gimbaling
Stage SeparationWith Core Stage

 The GSLV Mk. III core stage has specifications listed as:

Core Stage

TypeL-110
Length21.26m
Diameter4.0m
FuelUnsymmetrical Dimethylhydrazine
OxidizerNitrogen Tetroxide
Inert Mass10,600kg
Propellant Mass115,000kg
Launch Mass125,600kg
Propellant TanksAluminum Alloy
FuelUH25 - 75% UDMH, 25% Diazane
OxidizerNitrogen Tetroxide
Propulsion2 Vikas 2
Thrust (SL)677kN
Thrust (Vac)766kN
Specific Impulse293 sec
Engine Dry Weight900kg
Engine Length2.87m
Engine Diameter0.99m
Chamber Pressure58.5bar
Mixture Ratio1.7 (Ox/Fuel)
Turbopump Speed10,000rpm
Flow Rate275kg/s
Area Ratio13.88
Attitude ControlEngine Gimbaling
IgnitionT+110s
Burn Time200s
Stage SeparationActive/Passive Collets

 The cryogenic upper stage has specifications listed as:


Cryogenic Upper Stage

TypeC-25 Cryogenic Upper Stage
Length13.32m
Diameter4.0m
FuelLiquid Hydrogen
OxidizerLiquid Oxygen
Inert Mass~4,000kg
Propellant Mass25,000kg
Launch Mass~29,000kg
Propellant TanksAluminum Alloy
PropulsionCE-20
Engine TypeGas Generator
Thrust - Vacuum200kN
Operational Range180-220kN
Specific Impulse Vac443s
Engine Mass588kg
Chamber Pressure60bar
Mixture Ratio5.05
Area Ratio100
Thrust to Weight34.7
Burn Time580s
GuidanceInertial Platform, Closed Loop
Attitude Control2 Vernier Engines
Restart CapabilityRCS for Coast Phase

  Plugging these dry mass, propellant mass, and Isp values into the Schilling calculator gives these results:

Mission Performance:
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

TypeGS3 - C15
Inert Mass~2,500kg
Launch Mass~15,300kg
Length8.7m
Diameter2.8m
Tank MaterialAluminium Alloy
FuelLiquid Hydrogen
OxidizerLiquid Oxygen
Propellant Mass12,800kg
GuidanceInertial Platform, Closed-Loop
Propulsion1 ICE (CE-7.5)
CycleStaged Combustion
Thrust (Vacuum)73.5 to 93.1kN
Impulse454sec
Engine Dry Weight435kg
Engine Length2.14m
Engine Diameter1.56m
Burn TimeUp to 1,000sec
Chamber Pressure58bar
Attitude Control2 vernier Jets, each 2kN
RCS for Coast Phases
Stage SeparationMerman Band, Hot Staging

   Inputting these specifications instead into the Schilling calculator gives:

Mission Performance:
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
 

Could asteroidal impacts be the cause of the coronal heating problem?

 Copyright 2024 Robert Clark   A puzzle in solar science that has existed for 150 years is the corona heating problem: Why is the sun’s coro...