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Saturday, January 24, 2015

Ariane 4 for European manned spaceflight.

Copyright 2015 Robert Clark


The Hermes spaceplane because of its size was intended to be carried by the Ariane 5. However, that plan was cancelled because of cost. But if you use a smaller capsule then it could be carried by the Ariane 4.

Two versions would work for a fully liquid fueled launcher, the Ariane 42L and Ariane 44L, the first with two liquid-fueled side boosters and the second one with four. Versions of the Ariane 4 using solid side boosters were also made however for this manned spaceflight application I'm only considering all-liquid fueled launchers.

According to Astronautix, the Ariane 42L could carry 7,900 kg to LEO and the Ariane 44L, 10,200 kg.

Ariane 42L V56 
Ariane 42L V56 - COSPAR 1993-031

Ariane 44L 
Credit: Arianespace

 A crewed version of the Cygnus capsule probably could be produced to mass in the range of 2,000 kg dry mass:

Budget Moon flights: lightweight crew capsule.
http://exoscientist.blogspot.com/2013/04/budget-moon-flights-lightweight-crew.html



     Bob Clark




Thursday, January 15, 2015

NASA Technology Transfer for suborbital and air-launched orbital launchers.

Copyright 2015 Robert Clark

 I have become enamored of NASA's Morpheus lunar lander project. In the post "NASA Technology Transfer for manned BEO spaceflight", I discussed how it can be used to produce a manned lunar lander, or asteroidal lander, for a few 10's of millions of dollars, far less than the $10 billion estimated to be needed by NASA. And in "NASA Technology Transfer for Orbital Launchers", I discussed how its engines could be used for the small orbital launch system Firefly, resulting in a significant reduction in the launcher's development costs.

 I don't think NASA fully appreciates the usefulness of the Morpheus development. Here I'll show how the Morpheus itself can be used to produce suborbital launchers, and also the stages for orbital launchers. For instance the Morpheus can be used to provide the solution to DARPA's ALASA air launched, small orbital system.

 The Wikipedia page on the Morpheus gives its propellant load as 2.9 metric tons (mT) and dry mass as 1.1 mT. Its methane/LOX engine has an Isp of 321 s with a thrust of 24 kN, 2,450 kilogram-force (kgf).


 Note this means when fully fueled the single engine could not lift the vehicle in Earth's gravity. The single engine of course would be fine for its intended purpose as a lunar lander at 1/6th gravity. However, for a Earth launch system we'll use a half-size vehicle to be launchable with a single engine. Rounding off this gives it a propellant mass of 1.5 mT and dry mass of .5 mT. Compared to the full Morpheus this will have only two spherical propellant tanks instead of four, one each for the liquid methane and LOX.

 Since this will be reaching high velocity through Earth's atmosphere it will have to be streamlined. Then we'll place the two propellant tanks inline vertically. We'll also need an aeroshell. To save weight we could make the aeroshell composite. Another possibility would be to make the aeroshell inflatable. Since the aeroshell would not need to be load-bearing and with the possibility to make it inflatable we'll assume it adds only a small proportion to the weight. We could save additionally weight by making the tanks out of aluminum-lithium alloy, titanium, or composites. Alternatively, we could use a cylindrical tank to hold the propellants to eliminate the need for an aeroshell.

 Suborbital Case.

 This page gives the required delta-v for a suborbital flight as in the range of ca. 2,400 m/s:

Flight Mechanics of Manned Sub-Orbital Reusable Launch Vehicles with Recommendations for Launch and Recovery.
Mechanical and Aeronautical Engineering Department, University of California, Davis, CA 95616-5294
Marti Sarigul-Klijn Ph.D. and Nesrin Sarigul-Klijn*, Ph.D.
An approximate delta V to reach 100 km is 7,000 to 8,000 fps (2,100 to 2,400 m/s) for vertical takeoff, with slightly less delta V needed for air launch, and significantly more required for horizontal takeoff.
http://www.spacefuture.com/archive/flight_mechanics_of_manned_suborbital_reusable_launch_vehicles_with_recommendations_for_launch_and_recovery.shtml

  Now, at a 1.5 mT propellant load, .5 mT dry mass, .25 mT payload, and 321 s Isp, the vehicle can do a delta-v of 321*9.81ln(1 + 1.5/(.5 + .25)) = 3,460 m/s, sufficient for a suborbital flight.

 There are commercial opportunities for suborbital flight with NASA. Also using two to four copies or scaled up that many times this could also be used for a suborbital tourism vehicle.

DARPA Air-Launched Orbital Vehicle.

 DARPA is funding research into a small air-launched system called ALASA, As described in the blog post "Dave Masten's DARPA Spaceplane, page 2: an Air Launched System", high altitude supersonic air-launch at Mach 2 can cut 1,600 m/s from the delta-v needed for low Earth orbit. This would reduce the delta-v that needed to be supplied by the rocket from 9,100 m/s to 7,500 m/s.

 We'll use two copies of the half-size Morpheus firing in parallel and cross-feed fueling. Cross-feed fueling allows the upper stage to have its full level of fuel after staging, unlike the usual case with parallel staging. As in the earlier blog post, we'll again use the Star 17 solid stage as the final, orbital stage:

Encyclopedia Astronautica.
Star 17
Solid propellant rocket stage. Loaded/empty mass 124/14 kg. Thrust 19.60 kN. Vacuum specific impulse 280 seconds.
Cost $ : 0.580 million.
Status: Out of production.
Gross mass: 124 kg (273 lb).
Unfuelled mass: 14 kg (30 lb).
Height: 0.98 m (3.21 ft).
Diameter: 0.44 m (1.44 ft).
Span: 0.44 m (1.44 ft).
Thrust: 19.60 kN (4,406 lbf).
Specific impulse: 280 s.
Specific impulse sea level: 220 s.
Burn time: 18 s.
Number: 25 .
http://www.astronautix.com/stages/star17.htm

 Then we can get a payload of 55 kg to orbit by supersonic air-launch:

321*9.81ln(1 + 1.5/(.5 + 2.0 + .124 + .055)) + 321*9,81ln(1 + 1.5/(.5 + .124 + .055)) + 280*9.81ln(1 + .110/(.014 + .055)) = 7,690 m/s.


  Bob Clark






Tuesday, January 6, 2015

NASA Technology Transfer for manned BEO spaceflight.

Copyright 2015 Robert Clark

  The Morpheus lunar lander developed by NASA demonstrates how successful NASA can be when it takes a low cost approach to developing new systems. Not only can the Morpheus be used for unmanned robotic landers, but scaled up it can produce a manned lunar lander at orders (plural) of magnitude lower cost than the $10 billion NASA previously estimated for a lunar lander. Then Project Morpheus becomes a prime example of how NASA's Technology Transfer program can be utilized.



 The orders of magnitude reduction in cost is important for enabling the low cost commercial space approach to BEO (beyond low Earth orbit) spaceflight. For this to work, first, the companies that engage in the commercial space approach have to be convinced they can make a profit on such ventures. This is much easier to do when the development costs are in the ten's of millions of dollars range rather than the billions of dollars range

 Now note that by taking a smaller, modular approach to BEO spaceflight, as exemplified for example by the Early Lunar Access proposal, it is only a lander that would be needed to be developed for lunar flights.

Early Lunar Access. Credit: NASA

This actually is also the case for asteroidal flights since the delta-v to near Earth asteroids is actually less than that to the Moon. No hugely expensive super heavy lift launchers would be required. Only currently existing launchers, or those fully paid for by the developer such as the Falcon Heavy, would be used. Then that huge proportion of the development cost of such missions for the HLV would be eliminated.

 Then at least the development costs, aside from the launch costs, can become manageable for private financing. Then private BEO proposals such as by Golden Spike Company. Planetary Resources, Inc., Deep Space Industries, etc. become feasible.


  Bob Clark

Saturday, January 3, 2015

NASA Technology Transfer for Orbital Launchers.

Copyright 2015 Robert Clark


 NASA's Technology Transfer Program intends to partner with U.S. companies to commercialize technology developed at NASA. One example is the autonomous landing system for robotic lunar landers (ALHAT) developed by NASA which will be used by the Moon Express team for their Google Lunar X-prize entrant. 

 I suggest another example that has more important commercialization potential are the methane engines developed for NASA's Project Morpheus robotic lunar lander. The Firefly launch company intends to make small methane fueled orbital launchers. The largest single development cost for a launcher are usually the engines. By using the engines already developed for the Morpheus lander for the Firefly, a significant proportion could be cut from the development cost.


  Bob Clark

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