Budget Moon flights: lightweight crew capsule.

Copyright 2013 Robert Clark

 In the post Budget Moon flights, I argued that by using a capsule half-sized to the Dragon capsule at ca. 2,000 kg dry mass, that we could launch a manned lunar landing mission carried to LEO by a single Delta IV Heavy or Ariane 5 ME, with a separate man-rated launcher to carry the capsule and crew. And if using the Falcon Heavy at the stated $1,000 per pound price point it could be done at launch costs of less than $100 million.

 I'll discuss such a half-sized capsule here. This report discusses a capsule with such low dry mass:

Phoenix: A Low-Cost Commercial Approach to the Crew Exploration Vehicle

Abstract: Since the announcement of President Bush’s Vision for Space Exploration (VSE) in early 2004, the architecture of Project Constellation has been selected. The system will be centered around the Crew Exploration Vehicle (CEV), which has been dubbed by NASA administrator Michael Griffin as “Apollo on steroids”. The CEV is to be launched on a new launch vehicle, derived from existing shuttle technology. The development of this new
spacecraft and launch vehicle is a very costly proposition. An alternate approach is proposed in this study. The Phoenix is a smaller spacecraft designed specifically to be launched on the Falcon 5 vehicle under development by SpaceX. Because the SpaceX vehicle will cost only a fraction of today’s launch costs, the Phoenix is estimated to cost less than half of the price of the CEV. This reusable three person capsule utilizes an innovative re-entry concept, which allows for a cylindrical spacecraft with greater interior volume. This extremely cost-effective spacecraft is an attractive option for fulfilling VSE requirements.

 Below is page 3 from this report:

  ...
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  ...

 Our crew capsule would not need its own propulsion system, so get the dry mass by subtracting off the masses in the "Propulsion (SPS)" section to get 2,160 kg. We will retain the RCS system though. Note the 2,160 kg dry mass actually also includes the mass for a crew of three and the mass for their space suits, as well as the food and water for a mission to and from the Moon.
 The heat shield is of the innovative "parashield" design:

This with its truss structure/drive system would amount to about 15% of the capsules dry weight. This is about the same as Apollo era heat shields. However, the lightweight PICA-X material used on the Dragon capsule can withstand lunar return velocities and would only weigh half as much. This would subtract about 160 kg from the dry mass.

 This report on a "Phoenix" capsule only envisions this crew module to carry the crew from LEO to lunar orbit and back with a separate module to be used on a lunar lander, a la the Apollo architecture. But following the Early Lunar Access architecture we could use this one single crew module for the entire flight.

 It is notable that Orbital Sciences Cygnus capsule is of similar size to this "Phoenix" crew capsule.

Artist's rendering of Cygnus spacecraft approaching the International Space Station.

CREDIT: Orbital Sciences Corporation

 The Cygnus dry mass is 1,500 kg. This includes the propulsive service module at the base. The service module is based on Orbital's Star satellite bus. According to Astronautix this has a dry mass of about 800 kg. So the capsule itself is 700 kg. Adding on the life support elements as given in the "Phoenix" capsule report, would result in a dry mass of about 2,000 kg.
  This is important because the Cygnus is built by Thales in Italy so it means Europe could make all the components for the Moon mission, including the man-rated launch rocket as described in post The Coming SSTO's: multi-Vulcain Ariane.
 The "Phoenix" capsule report estimates billions of dollars in development cost. But this is using traditional NASA costing estimates. However, SpaceX has shown that development costs can be cut by a factor of 10 by private financing both for launchers and for crew capsules. From its half-size compared to the Dragon we might estimate its development as privately financed at half of the $300 million spent developing the Dragon, so to only $150 million.


   Bob Clark

Update, May 1, 2014:

 For the passage above highlighted in red, I am now inclined to think that 1,500 kg dry mass does not include the mass of the service module, which is, the propulsion unit at the bottom. It's because of this page:

Cygnus Spacecraft Information

Pressurized Cargo Module

Cygnus standard and Cygnus enhanced use Pressurized Cargo Modules built by Thales Alenia Space, Italy. The module is based on the Multi Purpose Logistics Module that was flown on multiple Space Shuttle Missions to deliver pressurized cargo to ISS.

It is 3.07 meters in diameter and has a length of 3.66 meters in its standard configuration and 4.86 meters with one segment added in the enhanced configuration. The standard PCM has a dry mass of 1,500 Kilograms and the enhanced version weighs 1,800kg. 2,000 Kilograms of cargo can be packed into the standard PCM while the enhanced version allows 700kg more to be loaded. The cargo module has a pressurized volume of 18.9 cubic meters in the standard configuration and 27 cubic meters in the enhanced configuration. Power consumption of the PCM is less than 850 watts.

http://www.spaceflight101.com/cygnus-spacecraft-information.html

 This page also has a separate section on the service module which suggests that mass amount above was only for the pressurized module itself.

Budget Moon flights.

Copyright 2013 Robert Clark 

In the blog post SpaceX Dragon spacecraft for low cost trips to the Moon, I argued that manned flights to the Moon could be mounted for costs in the few hundred million dollars range. This is compared to the $100 billion cost estimated for the now defunct Constellation program.
 The key to the low cost was using the SpaceX Falcon Heavy booster, and using already existing components, such as the Centaur upper stages. But I wondered could we bring the cost down even further using even smaller upper stages? Perhaps to even a few 10's of millions of dollars??
 I'll use again a combination of hydrogen-fueled stages, two copies of the Ariane 4 third stage, version H-10-III this time. There are variations in the cited specifications for this stage in various sources:

Ariane 4.
http://www.b14643.de/Spacerockets_1/West_Europe/Ariane/Design/Ariane_2.htm 

Die Oberstufen H-8, H-10 und ESC-A.
http://www.bernd-leitenberger.de/h-10.shtml 

Ariane-44L H10-3.
http://space.skyrocket.de/doc_lau_det/ariane-44l_h10-3.htm 

Ariane H10-3.
http://www.astronautix.com/stages/arieh103.htm

ARIANE 4 SPECIFICATIONS.
http://www.braeunig.us/space/specs/ariane.htm 

 It may be some sources are including the weight of the Vehicle Equipment Bay (VEB), others not. The VEB carried the avionics and telemetry equipment for the Ariane 4. We may suppose these functions carried out by the crew capsule, and at lighter weight than that used on the Ariane 4, first launched in the mid-90's.
 I'll use the numbers on Braeunig's "ARIANE 4 SPECIFICATIONS" page. It gives the dry mass of the stage as 1,240 kg and the propellant mass as 11,860 kg. The HM-7B engine used on that stage has an Isp of ca. 445 s.

 We'll use again this table of Earth/Moon delta-V's:

Delta-V budget.
Earth–Moon space.

https://en.m.wikipedia.org/wiki/Delta-v_budget#Earth%E2%80%93Moon_space%E2%80%94high_thrust

 With aerobraking on the return to the Earth, the total round-trip delta-V is 8,650 m/s.  We'll  use the architecture that the landing stage is used to return all the way back to Earth, not just to link up with a stage waiting in lunar orbit. And just a single crew capsule will be used that carries the crew all the way from Earth to the Moon and back again, no separate command and lunar modules, as with Apollo. This is analogous to the Early Lunar Access proposal of the early 90's.
 Then we could bring a payload of 2.4 metric tons(mT) to the Moon and back, sufficient for a half-Dragon sized capsule:

445*9.81ln(1 + 11,860/(1,240 + 13,100 + 2,400)) + 445*9.81ln(1 + 11,860/(1,240 + 2,400)) = 8,660 m/s.

 Various weight saving techniques can be used to save further weight on the stages.  The propellant tanks are made of aluminum rather than the heavier steel of the Centaurs. But they can be made lighter in the rang of 15% to 25% by using the aluminum-lithium alloy used on the later versions of the shuttle external tank (ET). Further weight saving techniques would be to use the common bulkhead and "balloon tank", i.e., pressure-stabilized, design of the Centaurs. 
 The tanks can be made even lighter by using composites, perhaps by 30% over aluminum-lithium:

The Composite Cryotank Technologies and Demonstration Project.

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20120002943_2012002501.pdf

 This means composites can save as much as 50% (!) off the weight of standard aluminum. Aside from tanks, composites can be used on the structural members to save weight. Composites could be used on the thrust structure, the intertank, the payload and interstage adapters, and the tank structural support members.
 A new composite structural technique is that of the isotruss:

Isotruss

 According to the manufacturer, it weighs only 1/12th as much as steel at the same strength:

isotruss strength comparison

 Boeing has already developed a 2.4 meter wide composite cryogenic tank: 

Boeing Develops Game Changing Composite Propellant Tank

 This would be sufficient for the H10-III stage. By next year they expect to have a 5.4 meter wide tank. This would allow a 20 mT standard sized Centaur to have the composite tanks. This size tank would also work for a scaled-up 40 mT Centaur-style stage. ULA has argued a stage this size following various weight saving techniques could get a 20 to 1 mass ratio. This would allow a stage that could make a manned round-trip lunar flight from LEO in a single stage.  

 ULA has discussed improving mass ratio as the best way to improve performance at the NewSpace 2012 conference. 

 From the graphs shown it is apparent propellant mass fraction, or mass ratio, is far and away the best way of increasing performance.  
Then to maximize the payload we can deliver to planetary targets, including on manned missions, it is important to implement such weight saving methods. For instance a single in-space stage for a lunar mission could be made reusable, thus cutting the cost of the in-space stage. 
 Conceivably just as important is that high mass ratio would also allow a single stage to orbit vehicle, SSTO. This has relevance to another ULA proposal of establishing propellant depots. It is known that a single stage capable of reaching LEO could also with orbital refueling make the round-trip from LEO to the lunar surface and back again as a single stage. So this one stage could make the entire trip from Earth to the Moon, with that one stop for refueling.
 The DC-X test VTVL test vehicle will be having its 20th anniversary this year. The event is to be celebrated by the participants in the program in August this year. It will also be marked by restoration of the DC-X for display in the New Mexico Space Museum.
 The DC-X itself was not intended to reach orbit. But interestingly the H-10-III stage, of similar size to the DC-X, may have this capability when lightweighted with composites. The single HM-7B engine it has though does not provide sufficient thrust for liftoff. You may need to add two to three additional engines. The increased thrust  however would add additional stress to the structure. It may require additional strengthening mass.   

COST. 
 For the lunar flight scenario, the total weight for the two H-10-III stages and the 2.4 mT capsule would be 28.4 mT. Using the estimated $2,000 per kilo cost for the Falcon Heavy launcher this would be a launch cost of $56.8 million. There is the cost also of the H-10-III stages. According to the "Ariane H10-3" Astronautix page, it's listed as $12 million. So the total cost of the launch would be $80.8 million.
 There is also however the cost of the capsule. We may suppose though the capsule is reusable so its cost per use might be only in the few million dollars range. Actually at a weight penalty of a few hundred pounds of propellant kept in reserve, which would subtract a proportionally smaller amount from the payload, the first H-100-III stage might be returnable to Earth to also be reusable. The second stage, used as the lander, could already be reusable since under the Early Lunar Access architecture it would also serve as the propulsive stage to return the capsule all the way to Earth. This would reduce the per use cost for the upper stages as well.
You would not be limited to using the Falcon Heavy, though this would be the lowest cost. The Delta IV Heavy with upgraded RS-68a engine has a 28 mT payload capacity to LEO, slightly less than the 28.4 mT needed at around a $300 million launch cost. However, the Delta IV Heavy is not expected to be man-rated anyway so you might as well launch the capsule on another rocket and link up with the propulsive stages in orbit.

ESA VERSION.
 This may be an architecture that could be implemented by the ESA. The Ariane 5 ME  is to have a 20% increase in payload to GTO to 12 mT. If the increase in payload to LEO is also 20% then that would bring the payload capacity to 24 mT. This would be enough to carry the propellant of the two H-10-III stages. Neither the current Ariane 5 nor the ME version will be man-rated however. We will need a separate man-rated rocket to carry the crew to orbit. This would have to lift the two H-10-III dry masses plus the capsule mass for 2*1.24 mT + 2.4 mT = 4.88 mT. I'll show in an accompanying blog post this is well within the capability of a vehicle made from a just a single stage of the Ariane 5 core with a second Vulcain 2 engine added. 
 The cost of the Ariane 5 is about $200 million. If the cost is also increased by 20% on the ME this would bring it to $240 million. And the cost for the manned launcher? I'm informed by a member of ESA that the estimated cost for a modified Ariane 5 core with a second Vulcain 2 engine would be ca. $60 milion. Plus the $24 million cost of the two upper stages brings the cost to $324 million.



  Bob Clark