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Tuesday, August 26, 2014

Merlin 1A engine for a hovering Falcon 9 v1.1 first stage.

Copyright 2014 Robert Clark

  Because of the high thrust of the Merlin 1D engine and the lightweight of the reusable Falcon 9 v1.1 first stage, the stage can not hover on its return to Earth. The firing of the engine has to be precisely timed so that the rocket reaches about zero relative velocity to the Earth once it reaches the landing point. SpaceX has referred to this non-hovering mode of landing as "hover-slam". It is due to the fact the thrust to weight ratio of the stage is still above 1 with single Merlin 1 D firing when the stage is nearly empty on landing:

More on Grasshopper’s “Johnny Cash hover slam” test.
http://www.newspacejournal.com/2013/03/09/more-on-grasshoppers-johnny-cash-hover-slam-test/

 Still for safety reasons I would prefer a stage that did have the capability to hover. One concern without the ability to hover for example would be unexpected large changes in wind speed and direction known as wind shear. Airline pilots know these when they are low to the ground as "microbursts"

Microburst schematic from NASA. Note the downward motion of the air until it hits ground level, then spreads outward in all directions. The wind regime in a microburst is completely opposite to a tornado.

 These can be potentially dangerous for pilots during takeoffs and landings since they result quickly in a large change in the aircraft's apparent airspeed, important for maintaining lift. There have been several airline accidents with wind shear identified as the cause. 

While wind shear is particularly dangerous for aircraft when near to the ground because it gives the pilots limited time to react, it does also occur at altitude. For both Space Shuttle accidents wind shear is suspected to have been a contributing factor. For the Challenger accident wind shear occurred about the same time as the shuttle reached Max Q, maximum aerodynamic pressure. This may have increased the stresses on the vehicle leading to a breach in the solid rocket boosters. In the case of Columbia, unusually strong wind shear occurred also close to Max Q that might have weakened the wing before the impact of the insulating foam.

  Recently, SpaceX had to destroy its Falcon 9R test vehicle during its last test flight:





  SpaceX has not released the cause of the accident but the rocket appeared to pitch over during the flight. There could be variety of reasons for this and not wind, but unexpected wind changes could cause it.

 The ability to hover gives you more leeway about where you land and some leeway when. You could then avoid the wind shear like an aircraft doing a go-around.

 So how to give it the ability to hover? One way would be to use a smaller engine for the landing engine. In fact SpaceX already has it in its inventory: the original Merlin 1A.

 The page on the Falcon 1 by Ed Kyle gives the Merlin 1A engine a sea level thrust of 34,900 kgf (kilograms-force). And Kyle's page on the Falcon 9 v1.1 gives the total sea level thrust using 9 Merlin 1D engines as 600,000 kgf. So one would be 66,000 kgf. Then replacing the Merlin 1D with the 1A would result in a loss of 31,000 kgf thrust. This is only a 5% loss of sea level thrust.

 Kyle's page on the F9 v1.1 though gives it dry mass of 19 metric tons (mT). Typically rocket engines leave some residual propellant left in the tanks at about 0.5%. This would be about 2,000 kg.This would give the first stage a mass at landing at about 21 mT. Then the Merlin 1A would need to be throttleable down to 60%.

 However, the Merlin 1D was made to be throttleable down to 70%, but the Merlin 1A never was. Then for this method to work the Merlin 1A would also need to be made throttleable.


       Bob Clark

Friday, August 15, 2014

Dave Masten's DARPA Spaceplane, page 2: an Air Launched System.

Copyright 2014 Robert Clark

 In the blog post Dave Masten's DARPA Spaceplane, I discussed using SpaceX Falcon 1 or Falcon 9 stages to achieve DARPA's XS-1 reusable first-stage spaceplane. Another DARPA program ALASA seeks to send smaller payloads of 45 kg to orbit for $1 million using air-launch. 

 DARPA has already awarded a contract to Boeing to produce the ALASA system:

Boeing Targets 66 Percent Launch Cost Reduction with ALASA.
By Mike Gruss | Mar. 28, 2014
http://www.spacenews.com/article/military-space/40023boeing-targets-66-percent-launch-cost-reduction-with-alasa

 However, using the Falcon 1 upper stage may provide a fast, low cost means to produce such a system. Masten Space Systems could develop this as well since it would provide a much reduced cost proof-of-principle for their larger spaceplane that in itself would still be profitable.

 SpaceX has said the Falcon 9 first stage accounts for 3/4 of the cost and the upper stage, 1/4. If we assume a similar ratio for the $8 million Falcon 1, then we might estimate the cost of the upper stage as $2 million. However, unlike with the Falcon 9, the Falcon 1 upper stage uses a much smaller and simpler engine in the pressure-fed Kestrel and it is a much smaller stage in comparison to the first stage than is the case with the Falcon 9. Then I'll estimate its cost to be, say, $1 million. That would already be at the $1 million max cost DARPA wants per launch for the ALASA system.

Solid Rocket Motor Expendable Stage version.
 To get the launch cost below $1 million we would need reusability. If we got 10 launches from the Kestrel powered booster, that would be $100,000 per launch for just this lower stage. Then staying below the $1 million max cost would depend on the cost of the upper stage. There were some small solid rocket stages that were below $1 million in cost such as the Star 17 solid rocket motor:

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

 This is not currently in production but there are probably some remaining in storage or some of comparable size.

 According to Ed Kyle's page on the Falcon 1, the F1 upper stage had a 0.36 metric ton (mT) dry mass, 3.385 mT propellant mass and 327 sec Isp. However, the Kestrel engine used only had a 3,175 kgf (kilogram-force) thrust, i.e., less than the stage weight. So we'll cut down the propellant load to 2.5 mT.

 Now we'll use the fact that airlaunch actually can result in a significant reduction of the required delta-v that needs to be supplied by the rocket to reach orbit and therefore a significant increase in payload. This is described in some reports by Sarigul-Klijn et.al.:

Air Launching Earth-to-Orbit Vehicles: Delta V gains from Launch Conditions and Vehicle Aerodynamics.
Nesrin Sarigul-Klijn University of California, Davis, CA, UNITED STATES; Chris Noel University of California, Davis, CA, UNITED STATES; Marti Sarigul-Klijn University of California, Davis, CA, UNITED STATES
AIAA-2004-872
42nd AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, Jan. 5-8, 2004
http://pdf.aiaa.org/preview/CDReadyMASM04_665/PV2004_872.pdf  [first page only]

 The conclusions are summarized in this online lecture:

A.4.2.1 Launch Method Analysis (Air Launch).
For a launch from a carrier aircraft, the aircraft speed will directly reduce the Δv required to attain LEO. However, the majority of the Δv benefit from an air launch results
from the angle of attack of the vehicle during the release of the rocket. An
ideal angle is somewhere of the order of 25° to 30°.
A study by Klijn et al. concluded that at an altitude of 15250m, a rocket launch with the
carrier vehicle having a zero launch velocity at an angle of attack of 0° to
the horizontal experienced a Δv benefit of approximately 600 m/s while a launch
at a velocity of 340m/s at the same altitude and angle of attack resulted in a
Δv benefit of approximately 900m/s. The zero launch velocity situations can
be used to represent the launch from a balloon as it has no horizontal velocity.
Furthermore, by increasing the angle of attack of the carrier vehicle to
30° and launching at 340m/s, a Δv gain of approximately 1100m/s
was obtained. Increasing the launch velocity to 681m/s and 1021m/s produced a
Δv gain of 1600m/s and 2000m/s respectively.
From this comparison, it can be seen that in terms of the Δv gain, an airlaunch is 
superior to a ground launch. As the size of the vehicle decreases, this superiority 
will have a larger effect due to the increased effective drag on the vehicle.
https://engineering.purdue.edu/AAE/Academics/Courses/aae450/2008/spring/report_archive/reportuploads/appendix/propulsion/A.4.2.1%20Launch%20Method%20Analysis%20(Air%20Launch).doc

 A speed of 340 m/s is a little more than Mach 1, while subsonic transport aircraft typically cruise slightly below Mach 1. So the delta-V saving could still be in the range of 1,000 m/s with air launch even using a standard subsonic jet, a significant savings by the rocket equation.

 And this study found by using a supersonic carrier aircraft you could double the payload of the Falcon 1:

Conceptual Design of a Supersonic Air-launch System.
43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
8 - 11 July 2007, Cincinnati, OH
http://www.ae.illinois.edu/m-selig/pubs/ClarkeEtal-2007-AIAA-2007-5841-AirLaunch.pdf

 The idea had been that airlaunch can't result in much of an improvement in payload since jet transports typically cruise only around 300 m/s, so, it was thought, you would only subtract this off the delta-v needed to reach low Earth orbit (LEO), which is about 9,100 m/s. However, there is also the altitude the aircraft can achieve and another key factor is the high altitude launch means you can use the higher Isp and higher thrust vacuum versions of the engines. The Isp advantage can be quite significant. For instance the Merlin 1C only had a vacuum Isp of  304 sec, but the Merlin Vacuum, being optimized only to operate at near vacuum conditions, had an Isp of 340 sec.

 The Boeing version of the ALASA system will use the F-15E fighter jet for the airlaunch. This has a Mach 2.5 maximum speed at altitude and can carry 10,400 kg payload. So we'll use this also for our system. Following the Sarigul-Klijn et.al. paper, the Mach 2.5+ max speed of the F-15E  is above the 681 m/s air launch speed needed to reduce the delta-v to orbit by 1,600 m/s. This will bring the delta-v that needed to be delivered by the rocket down to about 7,500 m/s.

 Using the reduced propellant load for the Falcon 1 upper stage of 2.5 mT then with a 45 kg, 0.045 mT, payload, an (F1 upper stage + Star 17) rocket could get a delta-v of:

327*9.81ln(1 + 2.5/(0.36 + 0.124 + 0.045)) + 280*9.81ln(1 + 0.110/(0.014 + 0.045)) = 8,500 m/s.

 This is high enough that a cheaper subsonic carrier, which according to Sarigul-Klijn can still subtract off about 1,000 m/s from the required delta-v, could be used instead of the Mach 2.5 F-15E. 

 Let's also estimate how much higher payload we could get using the reduction of delta-v to 7,500 m/s allowed by using the F-15E. Taking the payload to be 80 kg, 0.08 mT, we get:

327*9.81ln(1 + 2.5/(0.36 + 0.124 + 0.08)) + 280*9.81ln(1 + 0.110/(0.014 + 0.08)) = 7,560 m/s.

 So we could actually exceed the DARPA requirements to get 80 kg to LEO.

Two Falcon 1 Upper Stage Version.
 Instead of using an expendable solid rocket as the upper stage, we could use instead a second Falcon 1 upper stage. This will allow the possibility of getting a fully reusable system. We'll have both stages firing in parallel to be able to get a T/W greater than 1. We'll also use cross-feed fueling to maximize payload. For the upper stage that reaches orbit, we'll give it the full 3.385 mT propellant load since this stage doesn't have to have a T/W greater than 1. Then using the reduced 7,500 m/s required delta-v to orbit, we could transport 240 kg to LEO:

327*9.81ln(1 + 2.5/(0.36 + 3.745 + 0.240)) + 327*9.81ln(1 + 3.385/(0.36 + 0.240)) =7,530 m/s.

 We could also improve the mass ratio of these stages and increase the payload by switching to lightweight aluminum-lithium alloy for the propellant tanks. This could save as much as 25% off the tank weight. 


  Bob Clark


Monday, August 11, 2014

Dave Masten's DARPA Spaceplane.

Copyright 2014 Robert Clark

 Dave Masten's Masten Space Systems was recently announced as a winner of an award from DARPA to produce a reusable first-stage booster for a small orbital system:

Masten Space Systems selected by Defense Advanced Research Projects Agency for XS-1 Program.

 It is notable their version will be a winged booster. Previously Masten had worked on VTVL, i.e., vertical, propulsive landing vehicles. Masten describes the decision to go with a winged VTHL, i.e., horizontal landing, vehicle in a video interview on SpaceVidcast:



 At about the 41 minute mark Masten describes the fact that the need to return to the launch site to maintain low cost reusability after a high Mach flight, suggests high lift/drag ratio design and therefore wings.

 However, it would also work to use a lifting body. I discussed resurrecting the X-33 for this purpose in the post DARPA's Spaceplane: an X-33 version. It turns out the problem of getting conformally-shaped composite tanks, which doomed the X-33, becomes a non-issue if the vehicle is only to be used as a first stage booster. The reason is a first stage does not have to be as mass-ratio optimized so you can just use metal tanks. Still, despite that, in an up coming post I'll describe how it IS possible to get the lightweight tanks originally envisioned for the X-33 so in fact it is to possible to produce a SSTO VentureStar.

 In the interview, Masten also discusses a key difficulty is getting low cost engines that would be reusable that fit within DARPA's low cost requirements. He mentioned possibly using the engines XCOR is developing. I want to suggest the possibility also of using the Merlin engines as used on the SpaceX Falcon 1 first stage.

 The last quoted price for the entire Falcon 1 according to Ed Kyle's SpaceLaunchReport.com page on the Falcon 1 was $7.9 million from 2008. Based on that one would expect the cost of the engine alone would be less than that. Actually rather than developing a whole new first stage from scratch on this high risk project, as a preliminary development Masten might want to base a first version of his booster on the Falcon 1 first stage. By the specifications on Ed Kyle's page the Falcon 1 using the Merlin 1C only had a 470 kg payload to LEO, well less than the 1,400+ kg DARPA wants. Still this would lead to a faster and cheaper development to a reusable winged booster rather than creating everything from scratch. There is also the fact SpaceX is committed to launcher reusability and might even donate surplus Falcon 1's now in storage to the project. And Masten himself said during the SpaceVidcast interview he is spending much of his time working on the aerodynamics of such a winged booster rather than such questions as the propulsion.

  If SpaceX ever constructed the Falcon 1e, then Masten possibly might be able to use the Falcon 1e, which was to have double the size of the Falcon 1's first stage. According to Ed Kyle's page this was to have about a 1,000 kg LEO payload, closer to the DARPA requirements.

 Another possibility might be to use the Falcon 9 v1.1 upper stage. According to Elon Musk in discussing reusability, the first stage of the F9 is 3/4ths the cost and the upper stage 1/4th. So the cost of the upper stage would be in the range of $14 million. With 20 to 30 reuses this would be well within the $5 million per flight DARPA requirement for the program.

 A problem though is that as is usually the case with an upper stage the thrust is less than the full stage weight. We would have to cut down the tank size. According to Ed Kyle's page on the F9 v1.1 the propellant mass for the upper stage is estimated to be 93 metric tons (mT) and the dry mass 6 mT. Cutting the stage to be half size, take the propellant mass as 46 mT and the dry mass 3 mT.

 We need also to replace the Merlin Vacuum which can not operate at sea level with the Merlin 1D. By Ed Kyle's page the total thrust of the 9 engines on the F9v1.1 first stage is 600,000 kgf (kilogram-force). So one Merlin 1D would have sea level thrust of  67,000 kgf. Then using the 311 sec vacuum Isp of the Merlin 1D,  this could carry 12 mT to 4,280 m/s delta-v:

311*9.81ln(1 + 46/(3 + 12)) = 4,280 m/s.

 With approx. 1,200 m/s losses due to gravity and air drag this should be close to the Mach 10 DARPA requirement for the reusuable first stage, carrying a 12 mT total load for the upper stage and payload.


  Bob Clark



Monday, June 30, 2014

The Commercial Space Approach to Beyond Low Earth Orbit Spaceflight.

Copyright 2014 Robert Clark

 The National Research Council has released a report on NASA's plans for human spaceflight beyond low Earth orbit (BEO). It confirmed what many observers of the space program had already realized that NASA's current approach is unfocused and uninspiring, and most importantly unlikely to succeed:

NASA could not deliver humans to Mars, says new strategy report.
Published time: June 05, 2014 02:39
Landing humans on Mars is unattainable for NASA if the space agency’s current strategy and level of funding are not modified in the near future, according to a new congressionally-mandated report.

Of the three pathways to Mars that NRC suggested, two were associated with a return to the moon. A lunar landing and habitat would hone technologies that could later be employed on a Mars mission, the report said.
The Obama administration has publicly expressed distaste for continued, expensive moon landings. In outlining US space policy in 2010, President Barack Obama said, “I just have to say pretty bluntly here: We’ve been there before.”
The third option outlined by the report includes the Asteroid Redirect Mission, a plan still in the study phase but currently endorsed by the Obama administration.
Such a mission would send robotic spacecraft to essentially grab and re-orbit an asteroid passing near Earth, allowing astronauts to take samples of the rock.
That mission, though, is not preferred by authors of the report. Safety issues and development of “dead end” technologies render the asteroid mission inferior if NASA wants to reach Mars, it said.
The asteroid option “cannot provide the flight frequency required to maintain competence and safety,” the report posits.
http://rt.com/usa/163736-mars-nasa-funding-strategy/


 A major problem is that NASA is taking the Apollo approach to such missions, which required huge expenditures:

It’s time for NASA to abandon the Apollo mission model.
by John K. Strickland
Monday, June 23, 2014
What, then, is the opposite of the Apollo Mission model? It is the concept of a set of continuing missions designed to operate with a set of reusable space vehicles (both for crew and cargo), allowing the creation (and construction) of enduring infrastructure in specific locations in space and at surface destinations beyond low Earth orbit, using local materials, and making each subsequent mission easier, safer, and cheaper. One very critical aspect of this concept is the use of fully reusable spacecraft in an integrated cislunar transport system. Despite its name, such a cislunar system is not focused exclusively on access to the Moon, but to multiple cislunar and other inner solar system destinations. While at least one booster is about one year away from being operationally reusable, almost all spacecraft and other boosters are not, and according to some are in fact prohibited from being reused by NASA.
http://thespacereview.com/article/2537/1


  The Apollo program absorbed 5% of the federal budget at the time. This is compared to the 0.5% NASA currently gets. If NASA were funded to the same extent as during the Apollo era, it would be getting in the range of $180 billion per year(!) That's not happening.

 However, there is a way to get similar results without the huge payouts, and that is to follow the commercial space approach. Both SpaceX and Orbital Sciences were able to develop both launchers and capsules at 1/10th the cost of the usual fully-funded government projects. That's four separate systems able to cut 90% off the development costs. That is not just a coincidence.

 Detractors of commercial space have complained the name is a misnomer for they still get federal funding. However, the name you apply to it is irrelevant. What is important is the results you get for an order of magnitude lower cost. For the commercial space approach to work though the companies involved have to be convinced they can make a profit on the projects, for it involves a public and private partnership where the companies share part of the development cost, rather than it being fully government funded.

 But what BEO projects can be expected to make a profit? Well, SpaceX believes they can make a profit on making passenger flights to Mars, and the Golden Spike Company and Bigelow Aerospace believe they can make a profit by making crewed flights to the Moon. Also, Planetary Resources Inc. and Deep Space Industries believe they can make a profit on asteroid mining.

 This will be dependent though on these companies being willing and able to provide the needed funding for their part of the costs. For companies like SpaceX and Bigelow with wealthy, visionary leaders such as Elon Musk and Robert Bigelow this will likely be possible. But for the other start-up companies clearly this will depend on the size of the investment they need to make. Then a quite key fact to keep in mind is actually such missions can be accomplished at orders of magnitude lower costs than the amounts NASA estimates.

 The example of a lunar lander illuminates this quite clearly. NASA has said that we can't afford to return to the Moon because a manned lunar lander would cost $10 billion to develop. But the examples of the Masten XEUS lunar lander and the NASA Morpheus lunar lander show a manned lander actually can be developed for only a few ten's of millions of dollars in development cost. And the costs that need to be borne by the companies will be even less with the cost-sharing with government space agencies.

  I used the term "government space agencies" instead of NASA intentionally. The costs that needed to be borne by the companies would be reduced even further when more than one space agency contributed to the costs. For instance both NASA and ESA are contributing to the development of the Sierra Nevada Dream Chaser spacecraft. You could conceivably have all of NASA, ESA, JAXA, and Roscosmos contributing to the development cost of the BEO projects. Then the costs financed by the companies might only be 1/5th that of the already low commercial-space-approach development cost.


  Bob Clark

Thursday, June 19, 2014

Let's Do Project M, Robonaut to the Moon.

Copyright 2014 Robert Clark

 Project M was conceived in the aftermath of the cancellation of the Constellation program to return us to the Moon, with no manned missions beyond low Earth orbit (BEO) in sight. It appeared that a manned return to the Moon was too expensive. Project M would send a humanoid robot to the lunar surface as inspiration for us to do such manned missions BEO:





 The estimated cost was $450 million including launch cost. Part of the high cost though would be due to using the Atlas V 411 as the launcher. The SpaceX Falcon 9 v.1.1 rocket though would have comparable performance and would only cost $56 million. The lander used would be the Morpheus lunar lander. Considering the low cost of the Morpheus and the Robonaut, the cost conceivably might be in the $60 million range, certainly well less than $100 million.

 The non-reusable version of the F9 v1.1, which does not need to keep a fuel reserve for return to the launch site, has an estimated payload to geosynchronous transfer orbit (GTO) of 5,745 kg. The delta-v to GTO is 2,500 m/s and the total delta-v from LEO to the lunar surface is 5,930 m/s. Then 3,430 m/s delta-v would have to be supplied by the Morpheus lander. 

 According the Wikipedia specifications on the Project Morpheus lander, it has a propellant mass of 2.9 metric tons (mT), a dry mass of 1.1 mT, and an engine Isp of 321 s. Then it could transport 350 kg the rest of the way to the lunar surface:

321*9.81ln(1 + 2.9/(1.1 + 0.35)) = 3,460 m/s .

 Since Project M was proposed in 2010 there has been significant advance in humanoid robot locomotion, particularly in regards to walking in the vicinity of obstacles and on uneven terrain. DARPA has conducted a competition for humanoid robots completing physical tasks humans could normally do. NASA's entry was a version of the Robonaut given legs called Valkyrie at a 125 kg mass:



 However, Valkyrie did not perform particularly well during the DARPA robotic trials. It might be better to use the Robonaut torso only while using some more advanced robots for the legs such as Boston Dynamics' Atlas: 

 In this prior post I discussed the possibility of one of the Google Lunar X-prize entrants observing Earthrise from the lunar surface:

Full phase Earth from the lunar surface?

  This would have an even greater inspirational effect if done by a humanoid robot.
  However, Project M accomplished with the Morpheus lander would have benefits beyond just that of inspiring an interest in BEO spaceflight. Because the Morpheus lander was developed at such low cost it would give support for the idea a Morpheus-derived manned lander could also be low cost.
 
 This has importance to those of us who think that a return to the Moon is important to accomplish in itself for man's expansion into space, but also because of the fact it can be used as a stepping-stone for further interplanetary missions to Mars or to near Earth asteroids. 
  And in any case the NASA plan of having a station at L2 from which to direct robotic operations on the lunar surface would still benefit from humanoid robots performing those operations and of those robots being delivered by low cost landers.

  Bob Clark


Update, June 28, 2014: 

  Note the delta-v to many Near Earth Asteroids (NEA's) is less than that to the surface of the Moon:

Near-Earth Asteroid Delta-V for Spacecraft Rendezvous.
http://echo.jpl.nasa.gov/~lance/delta_v/delta_v.rendezvous.html 

 Then we could also do low cost Robonaut missions to NEA's using the Morpheus lander. This will have importance to accomplishing manned missions to NEA's as well as asteroid mining operations.

  

Wednesday, June 18, 2014

Full phase Earth from the lunar surface?

Copyright 2014 Robert Clark


The full Moon sometimes looks huge when it's low on the horizon. I was
wondering if the same effect would occur on the Moon if we saw a full phase
Earth from the lunar surface.
There haven't been many pictures taken from the surface of the Moon that
show the Earth. Here's one:





The Earth can look pretty large low on the horizon from lunar orbit even
when not at full phase as shown by this Apollo 11 image taken from orbit:




  
Are there images taken from the lunar surface either by robots or manned
missions that show the Earth rise at or close to full phase?

Nice article by Dr. Pauls Spudis:

Earth Rising.
Earth as seen from the Moon is always in the same place – true or false? It
depends.
By Paul D. Spudis
May 15, 2014
Thus, there are places on the Moon from which we can stand and
contemplate the sheer beauty and magnificence of a slowly rising Earth.
Given the sea change in global perspective provided by the famous Earthrise
picture taken by the Apollo 8 crew almost fifty years ago, what societal
impacts will occur when a human being stands on the lunar surface and
watches the Earth slowly rise above the horizon?  I suspect that a similar
shift in planetary perspective will occur.  If history is any guide, such a
shift will have profound psychological and political implications  – both
positive and negative – in our reach for the stars.
http://www.airspacemag.com/daily-planet/earth-rising-180951474/#ixzz321FRfV82

 I wonder if a Google Lunar X-Prize entrant could land at one of the
locations where Earthrise would be visible.


   Bob Clark

Friday, June 13, 2014

The Morpheus lunar lander as a manned lander for the Moon.

 Copyright 2014 Robert Clark

 Nice article here on the Morpheus lunar lander:

Project Morpheus Concludes Successful Flight Test Campaign With Spectacular Night Launch.
By Mike Killian
http://www.americaspace.com/?p=61298&cpage=1


The project leader notes it could be scaled up to be a manned lander. Based on specifications of the lander I estimate it would need to be scaled up by a factor of three to form a descent stage while using the original sized version for the ascent stage. The delta-V from low lunar orbit to the lunar surface is 1,870 m/s each way:

Delta-V Budget.
Earth-Moon space.
http://en.wikipedia.org/wiki/Delta-v_budget#Earth.E2.80.93Moon_space

 According to the given specifications, the Isp of the Morpheus engine is 321 s and the propellant load is 2.9 mT and dry mass, 1.1 mT. So with a 2 mT lunar capsule mass, the ascent stage consisting of a single Morpheus would have delta-v of:

321*9.81ln(1 + 2.9/(1.1 + 2)) = 2,080 m/s.

 The descent stage consisting of the Morpheus scaled up three times would have a 8.7 mT propellant load and 3.3 mT dry  mass. Carrying the 4 mT of the ascent stage and the 2 mT capsule, the descent stage would have delta-v of:

 321*9.81ln(1 + 8.7/(3.3 + 4 +2)) = 2,080 m/s.

  Another nice article describes the origin of the idea of the Morpheus and its innovative, low cost approach:

A father-son chat leads to first-of-its-kind NASA spacecraft.
By Thom Patterson, CNN
updated 8:00 AM EDT, Mon May 19, 2014 |
http://www.cnn.com/2014/05/18/tech/big-idea-morpheus-lander/

 Based on a $14 million development cost for two prototypes, one scaled up by a factor of three might cost $21 million. So $28 million for both stages. Actually by the Wikipedia page on Project Morpheus, the parts to build the Morpheus version 1.5B were only $750,000. So construction of a single Morpheus was probably well less than $7 million, and the cost for one three-times scaled up one well less than $21 million.

  Instead of scaling up the Morpheus, we could also combine three of the original size to form the descent stage, with the same development cost of $21 million. This would have an advantage of a quicker time to producing a flight capable prototype. Another problem with the scaled up version of the descent stage is that based on the 12 foot height of the original version I estimate an 18 foot height of the descent stage. That would be a high climb down for the astronauts. Constructing the descent stage of three copies of the original-sized Morpheus though would allow you to connect them together on a single level so the climb down would still be 12 feet.

  In any case we see again, just as with the Masten Xeus lander, a manned lunar lander can be made for 10′s of millions of dollars rather than the $10 billion of the Altair lunar lander.

      Bob Clark