Friday, May 29, 2015

A Vertical Landing SSTO - a "Space Shuttle" NASA Missed.

Copyright 2015 Robert Clark

Saturn V's S-II second stage as an SSTO.
The S-II stage during stacking operations of Apollo 6 in the VAB
The Saturn V launcher of the Apollo program was remarkable in the lightweight features of its upper stages, the S-II and the S-IVB. This page gives a list of the fueled weights and empty weights of the Saturn V stages:

Ground Ignition Weights

  The mass efficiency of the upper stages led to some proposals to use them together as an independent launcher, the Saturn II:

 However, when this Saturn II was proposed there were not high performance hydrolox engines that could operate at sea level. Therefore it was designed to use the upper stage engine already used on the Saturn V upper stages, the J-2. Since this was not designed to operate at sea level though, this limited the performance of the rocket.

 But in the late 70's when the space shuttle was being designed and built the high performance Space Shuttle Main Engine , the RS-25, was developed. Interestingly, if SSME's were used on the S-II stage you would get a fully reusable rocket as an SSTO that would have comparable payload to orbit as the space shuttle, i.e., no solid rocket boosters required.

 The later versions of Apollo had improved weight optimization. We'll use the specifications for Apollo 14. The "Ground Ignition Weights" page gives the Apollo 14 S-II dry weight as 78,120 lbs., 35,510 kg, and gross weight as 1,075,887 lbs., 489,040 kg, for a propellant mass of 997,767 lbs., 453,530 kg. 

 The SSME has a mass of 3,500 kg while the J-2 had a mass of 1,788 kg. We'll replace the 5 J-2's used on the S-II with 3 SSME's. This increases the stage dry mass by 1,560 kg. We'll use Dr. John Schillings Launch Performance Calculator to estimate the payload to LEO. 

 Enter in 37,070 kg for the stage dry mass, with the new SSME's replacing the J-2's. Enter in 453,530 kg for the propellant mass. Enter in the vacuum thrust with the max thrust at the 109% level in kilonewtons as 3*2,280 kN = 6,840 kN. Enter in the vacuum Isp of 452.3 s. Select "No" for the "Restartable Upper Stage" option, otherwise the payload will be reduced. And enter in an inclination of 28.5 degrees to match the latitude of the Cape Canaveral launch site of 28.5 degrees.

 Then the calculator gives the mass to LEO as 27,077 kg:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  27077 kg
95% Confidence Interval:  18536 - 37061 kg

"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters

Heat Shield and Landing Legs.
  We'll envision this as a VTVL (vertical take-off vertical landing) SSTO. Then we need to add heat shield, and landing legs. For weight of the heat shield, from the Apollo era it was about 15% of the weight of the reentry vehicle. However, SpaceX's Pica-X is about half the weight so about 8%. One scenario would be for the heat shield to be at the top with the stage entering top first and rotating to put the bottom down after reentry as with this proposal by SpaceX for the Falcon 9 upper stage:

 For the landing legs, that is commonly estimated as 3% of the landed weight:

 However, with modern composite materials we can probably get it to be half that. So call it 1.5%. 

Hovering capability for a VTVL vehicle.
 It was always taken as a given that a reusable VTVL rocket should have hovering capability. See for instance this discussion between noted space historian Henry Spencer and Mitchell Burnside Clapp who worked on both the DC-X and X-33 programs:

 Hovering ability and the low thrust capability this requires allows fine control of both the attitude and velocity in 3-dimensions for unexpected winds on landing. If you don't have that then that limits your ability to make small, fast corrections to attitude and velocity. The result will be repeated over-corrections and over-corrections to those over-corrections until time and space run out:

Hovering capability for the reusable Falcon 9.

 As with the Falcon 9 case, even one SSME would have too much thrust to allow this vehicle to hover. As discussed in the "Hovering capability for the reusable Falcon 9" post, you could apply various attachments to one of the engines to reduce the thrust on landing. You would though have to arrange the engines to be a in a straight-line to have a center engine rather than the clustered arrangement used on the Space Shuttle.

RL-10's as Landing Engines.
 Another possibiliuty to allow hovering would be to use multiple small engines for the landing. The RL-10A5 engine used on the DC-X would work. This is a version of the RL-10 with a shortened nozzle to operate at sea level:


 We'll use eight of them at the bottom of the stage arranged around the outer rim. Since the S-II stage had a 10 meter diameter these would still fit underneath the stage whether the three SSME's were arranged in the clustered format as with the Space Shuttle or in a straight-line.

 These will add 8*143 kg = 1,144 kg to the dry mass, but using them also at launch will add 8*64.70 kN = 517.6 kN to the vacuum thrust.

 Calculated Stage Weight for the Reusable Rocket.
  Adding on the 1,144 kg to the dry mass gives 38,214 kg. At 1.5% of the landed weight for the landing legs this would be 0,015*38,214 kg = 573 kg additional weight to the stage. The estimated propellant that needs to be kept on reserve for landing as discussed in the "horizontal vs. vertical landing" link, is about 10% of the landed weight. This would be about 3,879 kg reserve propellant. This plus the landing legs brings the reentry mass to 42,665 kg. The heat shield weighs 8% of this to bring it to 46,079 kg. 

Calculated Payload for the Reusable Rocket.
 We'll enter in now into the Schilling calculator 46,079 kg for the dry mass, subtract off the 3,879 kg kept on reserve from the propellant mass, and add on 517.6 kN onto the vacuum thrust. The result is:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  18706 kg
95% Confidence Interval:  10073 - 28808 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.

 This is less than the shuttle, but we can increase the payload as was done with the shuttle by using aluminum-lithium alloy for the propellant tank. This can shave off 25% from the propellant tank weight.

Payload Increase on Switching to Aluminum-Lithium Tanks.
 The propellant tanks on the S-II weighed 3% of the total stage weight:
SP-4206 Stages to Saturn
7. The Lower Stages: S-IC and S-II
The S-II turned out to be a comparatively advanced stage in terms of the existing state of the art. Although the S-II carried about 426 400 kilograms of liquid oxygen and liquid hydrogen, the tank structure, though supporting the structural mass, accounted for just a shade over three percent of the stage's total fueled weight. A common bulkhead much larger than that in any previous rocket averted the need for an [212] interstage between the oxidizer and fuel tanks; this reduced the total length of the stage by over 3 meters and saved about 4 metric tons of extra weight. In technical terms, the fabrication of the bulkheads called for unusually demanding accuracy in meridian welds that joined the bulkhead gores together. The welding operation joining the curved, 6-meter-long seams together had to be made to specifications allowing less than 0.33 millimeter of a mismatch. Then there was the problem of insulating the big liquid hydrogen tank, filled with thousands of liters of the super-cold propellant. Otherwise, the basic design elements of the S-II seemed conventional enough in that it consisted of eight major structural components and six major systems, all of which reflected the usual kind of basic elements associated with both the S-IC and the S-IVB.28

  Then the 3% of the S-II gross mass of 489,040 kg would be 14,671 kg for the propellant tank mass. A 25% savings by switching to Al-Li would be 3,667 kg. This would bring the payload to 22,373 kg, about the same as the Space Shuttle.

  Bob Clark

Friday, April 17, 2015

Hovering capability for the reusable Falcon 9.

Copyright 2015 Robert Clark

 We have now two landing attempts by the reusable F9 first stage onto the SpaceX barge. Both were unsuccessful. From the appearance of both failed landings it would appear that the capability to hover could have made both landings successful:

 About this latest test landing, Elon Musk in a Tweet has acknowledged that not being able to hover will result in a high g landing:


   This is confirmed by the rather high rate of descent apparent in the video, even though, as has been reported, the video is slowed down.

 Another disadvantage of not having hovering capability is apparent in the video. In correcting for mistakes in the angle of tilt, the engine having limited throttle capability will tend to over correct. That is, without hover and its low thrust capability, you can't make fine adjustments to the rocket orientation. Then at low altitude with little time to make corrections to the over-corrections, this can lead to failed landing due to the need for repeated adjustments and readjustments.

 In the blog post "Merlin 1A engine for a hovering Falcon 9 v1.1 first stage" I suggested one possible solution to the hovering question would be by using the lower thrust Merlin 1A engine. However, it would have to be made throttleable for this to work. In further updates to that blog post, I suggested either using the preburner exhaust or using variable size nozzles.

 Indeed all the proposals discussed in the "Altitude compensation attachments for standard rocket engines, and applications" post could also be used to make variable nozzle attachments to the engine nozzles to reduce the thrust when needed to allow hovering. For instance the carbon nanotube "rubber" attachment could be made to restrict the exit area to reduce the thrust and the "internal spike" proposal could be made to flare out to direct some proportion of the thrust laterally outwards rather than downwards to reduce the downwards thrust.

 These would reduce the efficiency of the engines, i.e., the Isp would be decreased for the hovering proportion of the flight. However, the altitude compensation attachments actually increase the payload perhaps as much as 25% for multi-stage rockets so overall the result will still be an improvement of the payload capacity.

 These altitude compensation attachment proposals do need more R & D work however, and SpaceX might want a quicker fix that can be attached quickly to the engines, or likely just the central engine for the landing phase.

 A possibility is suggested by this collapsible vegetable steamer:

 You would make an attachment like this that could flare out or be closed up, except it would have no holes on the sides. The open position would be usual formation used during the flight. The closed up position would be used only during hover to restrict the thrust.

 Another simple attachment might be the exhaust steering vanes used on rockets prior to the advent of engine gimbaling for steering:

      They could be used to direct a portion of the thrust laterally to reduce the downward thrust.

    Bob Clark

 UPDATE, April 21, 2015:

 Someone suggested to me another method to restrict the thrust to allow hovering, the variable nozzles put on some fighter jet engines:

Friday, April 3, 2015

Atmospheric humidity for water generation.

Copyright 2015 Robert Clark

*Suggestions for Water Production for California.*

California is still in the midst of a huge drought:

California’s About to Run Out of Water. We Have to Act Now.
Annie Sneed, Science, 03.23.15 7:00 am

 The state has now ordered a mandatory 25% drop in water usage.

 There appear to be however relatively simple methods of providing extra water. One is by deriving water from humidity in the air, the other by distilling the water from the ocean. Both would appear to have relatively low cost solutions.

 First the humidity solution. The amount of water in the form of water vapor is substantial, especially at high humidity. A key fact is air can store more water vapor at higher temperatures. But the point is the areas in California with the highest drought level are the areas with routine high temperatures, such as Los Angeles. So lets calculate the amount of water in air in Los Angeles. This page gives a graph of average relative humidity levels in LA:

Average Weather For Los Angeles, California, USA.
The average daily high (blue) and low (brown) relative humidity with percentile bands (inner bands from 25th to 75th percentile, outer bands from 10th to 90th percentile).

 You see that the relative humidity commonly reaches the 80% range and above, especially during the warmer months. The relative humidity is the percentage of the maximum possible water vapor the air can hold based on that temperature, called the saturated vapor density. This page gives a calculator for the saturated vapor density based on temperature.

Relative Humidity Calculation.

 For a temperature of 80 degrees F, the saturated vapor density is given as 25.4 gm/m3. And for a relative humidity of 80%, this corresponds to an actual amount of water in the air of 20.3 gm/m3, 0.0203 kg/m3.

 However, in actuality when the temperature is highest such as in the afternoon the humidity is lowest, and when the temperature is lowest such as in the early morning, the humidity is highest. So let's use some more representative temperatures and humidities. This page gives the  Los Angeles weather over a two-week period:

Weather past 2 weeks in Los Angeles, California, U.S.A.

 For the period as of today April 3, 2015, for a time during the day when both the temperature and the humidity are high, say, in the 60's range for both, this results in total amount of water in the air of about 10 gm/m3.

*Residential Solutions.*

I.) Some ceiling fans can move quite large amounts of air at a time resulting in quite large amounts of water vapor inflow at a time, if placed for example in residential windows. For instance some fans on this page at their higher speed settings can move 5,000 cubic feet per minute (CFM) of air while using only 30 watts:

ENERGY STAR Most Efficient 2015 — Ceiling Fans 52 inches and under.

  Lower speeds give better energy efficiency, but at the cost of lowered air flow amounts. We’ll use the high speed numbers since we want large air flow to get high water vapor inflow.

 At this high amount of hot air inflow though we would probably want the fan in a basement or attic since we would also need a window for the hot air, once the water is removed, to exit.

 How much water would need to be produced? An average person uses about 50 gallons per day of water:

City Utilities: Water Tips.

 So say a residence for a family of 4 needed 200 gallons of water per day. A gallon is 3.785 liters, so this is 757 liters per day. Water weighs 1 kilo per liter so this is 757 kilos per day. Then how long would it take for a fan blowing in 5,000 cubic feet per minute to bring in this much water(vapor)? At 5,000 cfm this is 5,000/3.283 = 141.7 meters per minute.

 At the the 80 degree F and 80% humidity level, each of these cubic meters of air would contain 0.0203 kg of water vapor. So this would amount to 2.88 kg per minute of water. This would take 262.8 minutes, about 4.4 hours. However, since the humidity actually goes down as the temperature goes up we would need to add up the total amount of humidity during a full 24-hour day. Looking up the representative amounts on the "Weather past 2 weeks in Los Angeles, California, U.S.A." page, we see the totals will still be above 200 gallons per month.
 This amount of power, only 30 watts used for 4.4 hours, is also quite small in energy costs compared to what the county of Los Angeles charges for water.

 An additional question to be resolved however is how can we convert this water vapor to liquid water? Air conditioners are able to do this, by accident, by chilling the air. Air dehumidifiers also commonly work this way. This causes water to condense out like happens for example in cool morning temperatures with morning dew. On our relative humidity calculator page, at 80 degrees F and 80% relative humidity, the dew point is only 73.4 F.  However, both air conditioners and air humidifiers use quite high power levels. We want to minimize additional power used.

 Some possibilities:

1.)If the water produced this way is an adjunct to the water received from the city, then we can use the cool water coming from the city water supply, typically around 50 degree F, to cool this air and get the water to condense out.

2.)It would be nice though, since the calculation showed this air-produced water alone is sufficient to supply the entire household water needs, to find a way that didn’t use the city water supply.

a.)Another type of air dehumidifier uses desiccants to absorb water vapor. The desiccant material is then heated to release the water as liquid and the same desiccant is used over again. However, this material typically is a silica compound and you would not want remnants of this to be left in the water. This also uses additional power for the heating step. If a desiccant could be found that is a type of mineral you would normally see for example in spring water then this might work. You would though need to find a way to get the water to be released as a liquid. Heating as with air dehumidifiers would work. However it may be at the high temperature of southern California would be sufficient so this would happen naturally.

b.) A similar possibility derives from the fact that rain droplets can frequently condense in the air out of water vapor at temperatures higher than they would normally do by having nucleation sites:

Cloud condensation nuclei.

 Then we could add nucleators into the air stream to get the air to condense. These nucleators though again would have to be a non-toxic if ingested. Ideal would again be some type of mineral commonly found in mineral water.

c.)To get the water to condense we could also expand the air flow. Rapidly expanding the air would cause the temperature to drop thereby chilling the water. A problem here though is the air flow is so large it might require an unreasonable size of expansion needed to get the needed temperature drop.

d.)Another possibility would be by increasing the pressure of the air. Just as increasing the pressure increases the temperature at which water makes the transition from liquid to gas, the boiling point, so also the temperature at which it makes the transitions from gas to liquid, the dew point, also increases. This page gives a calculator for how the pressure changes the dew point:

Dew Point Conversion Calculator.

 Enter in 73.4 degrees F in the known dew point field for our 80 degree F and 80% relative humidity scenario. Enter in 0 for the “psig” field, which measures how far this is above standard pressure in psi. Then a psig of only 4 gives a dew point of 80.3 F. That is an increase in pressure of less than 30% results in the water condensing out.

 We might be able to generate this amount of extra pressure by circulating the air around in a circle by centrifugal force. A problem though is the size of the air stream coming from a large ceiling fan size diameter might make the circle size needed impractically large. We could constrict the air coming from the fan into a smaller pipe diameter, but by the Bernoulli principle this would reduce the pressure. It still may be possible though that some combination of restricted pipe diameter and circulation diameter size could provide the needed pressure change at a practical size.

 Another possibility for using centrifugal force to effect the compression is by breaking up the airflow into multiple tubes at the microscale. The compression in this scenario is by the fact the acceleration around a circle is a = v2/r, v being the speed, and r the radius of the circulation. However, for our scenario the speed would only be around 3 to 4 mph, resulting in low acceleration and therefore compression for a large size pipe. However, by breaking the airflow up into multiple microtubes we can make the radius r very small giving high acceleration even for low speed.
 Support for the idea this can provide the needed compression comes from the several papers published on MEMs (micromechanical systems) rotors and turbines at the mini and microscales. Running at the hundreds of thousands of RPM's range these can provide pressure ratio's of from 2 to 3, sufficient for our purposes.

 At a slow airflow speed of 3 to 4 mph though these microtubes might need to be only be at the micron scale diameter to get such high RPM's. We might also constrict the pipe diameter first, which by Bernoulli's principle will increase the speed, before we break the air flow up into multiple microtubes (patent pending).

 For either the large pipe or microtube solutions using centrifugal force we could instead of bending the tubes around in a circle, have rifling inside the tube to cause the air to circulate helical fashion within the tube.

 II.) For the single fan bringing in 5,000 cfm it was surprising this could be done at only 30 watt power level. It would be interesting to find out then how large we could make the air inflow in for a single residential house. The ceiling-fan sized fan we used was about 1.5 square meters. For an average size house of a 100 m2 floor plan, this would provide about 66 times as much air flow and therefore water, if it were brought in through the roof. So this one home could provide the water needs for 66 other homes.

 This assumes that power requirements are just for running the fan and you would not need power for instance to do the condensing step. Since the power requirements are about at 1/10th that of the cost of water for Los Angeles, the home owner could sell the water at much reduced rates.

 However, it may be possible to reduce or eliminate the need for power to run the fan by using the wind flowing over the roof. A problem though is by the Venturi effect, the wind blowing horizontally over the roof would have the tendency to draw the air out of the house rather than drawing air in. If a solution could be found to draw the air in rather than out while using the air flow of the wind the costs would then be essentially nothing.

*Water Utility Solution.*

 For the water company using this method to produce extra water it becomes particularly simple. The large ceiling-fan type fans discussed in the residential solution only move air in the range of 3 to 4 miles per hour. But wind speeds commonly are above this speed especially in coastal areas. Then you would not even need to use fans for the water utility solution. You would just collect the air driven by the wind in large tubes for processing. Also, for California communities near the coast, using the ocean to supply the necessary cooling to condense the water vapor becomes especially simple (patent pending).

*Sustainable Water Resources*

  Up to 1 billion people don't have reliable access to clean drinking water:

 This would provide a simple, low cost means of providing clean water, assuming the condensing step could be done without extra power. If a fan is needed, the low amount of power could be supplied by a wind mill or solar power. However, it may be that the air flow from the wind itself would be sufficient and not even a fan would be needed.
*Individual Water Production.*

 For a scenario where one is stranded out in the wild or in a life raft without water, the amount of power needed for the fan is so small that it could probably be generated by hand for a single individual just for enough water to sustain life. And considering the wind speed needed is only in the range of 3 to 4 mph a fan might not be needed at all.

 Also, at least for an 80 F and 80% relativity humidity scenario, the 30% extra pressure needed to allow water to condense could be easily supplied by hand. By the Ideal Gas Law PV = RT, to get a 30% increase in pressure we would need to make less than a 23% decrease in volume, assuming we did the compression slowly so as not to increase the temperature. For instance, a piston in a foot long cylinder would only need to make a 3 inch compression to get the needed pressure.
 The life boat case would also be very simple because the condensing could instead be done by using the cool temperatures of the ocean water.

    Bob Clark

Sunday, February 15, 2015

Low cost Europa lander missions.

Copyright 2015 Robert Clark
 NASA has proposed funding for a future orbital mission to Europa. This is expected to have a billion dollar cost. A NASA-funded lander mission to Europa was expected to be too expensive. University of Arizona researcher Christopher Impey suggests doing privately funded missions to Europa:

Let’s Send a Private Mission to Europa, Expert Says.

 Indeed, by following the commercial space approach an actual lander mission, not just orbital, can be mounted for costs in the range of one of NASA's lowest cost Discovery-class missions.

 Some observations by the Hubble telescope were that Europa may have plumes like Enceladus though those observations have not been confirmed. If it does, then it may be the subice ocean on Europa could be reached simply by traveling though the fissures, like on Enceladus, not requiring melting through the ice.

 For getting funding for a privately-financed mission, imagine how much interest there would be among the public if we could actually access this subsurface ocean. As a point of comparison the Mars Pathfinder missions web site caused such overloads, with 40+ million hits per day, that some mirror sites crashed or had to have access limited:

Traffic on Mars
by Chuck Toporek
Asst. Managing Editor
Web Review
However, the most interesting and little known fact about the amount of traffic to the mirror sites comes from France, where the government actually pleaded with computer users to stop accessing the two Mars Pathfinder mirrors. You see, the phone systems in France carry all of the Internet traffic in the country, so when people started visiting the mirror sites at VisuaNet and Le Centre National D'Etudes Spatiales (CNES), they tied up the phone lines and basically disabled the country.

 Private companies that provided funding could be allowed to advertise on the mission web site.

Mission Delta-V requirements.
 The spacecraft first has to be sent on a trajectory towards Jupiter. This requires a delta-v of 6.3 km/s. Following a Hohmann trajectory it takes about 2.7 years to arrive at Jupiter. Aside from this Jupiter transfer delta-v, it will need to be placed into Jupiter/Europa orbit on arrival, and then sent towards a landing on Europa. According to this report this will require a delta-v in the range of 3.9 km/s:
European Cryo-Ocean Exploration Submersible (ECOES).
Preliminary Design Report. 
 We'll use a cryogenic, hydrogen-fueled stage for the injection into a trajectory towards Jupiter. And because of the long travel time, we'll use storable (hypergolic) propellants for the propulsive maneuvers at Jupiter.
Falcon 9 v.1.1 sized launcher.
 The expendable version of the Falcon 9 v1.1 has a payload to LEO of ca. 16.6 metric tons (mT). We'll use the Ariane H10-3 hydrogen-fueled upper stage for the Jupiter trajectory insertion. It has a 11.86 mT propellant mass, 1.24 mT dry mass, and 445 s Isp. Then it can carry 2.4 mT to the 6.3 km/s delta-v needed for the flight towards Jupiter:
445*9.81ln(1 + 11.86/(1.24 + 2.4)) = 6,324 m/s
 For the stage, at Jupiter we'll use the Integrated Apogee Boost Subsystem (IABS) stage.
 An artist's concept of a DSCS satellite being boosted by the IABS. Photo: U.S. Air Force
 This is a small kick-stage used to put geosynchronous satellites in their final orbits. It has a 1,578 kg gross mass and 275 kg dry mass, for a 1,303 kg propellant mass, and a 312 s Isp. Then this could provide a 220 kg spacecraft with the 3.92 km/s delta-v needed to land on Europa:
321*9.81ln(1 + 1303/(275 + 220)) = 3,947 m/s. 
 The Falcon 9 v.1.1 costs in the range of $56 million, the Ariane H10-3 cryogenic stage costs in the range of $12 million and the IABS stage costs in the range of $15 million, for a ca. $83 million launch cost.
 For a model of a low cost lander we might use the Mars Pathfinder mission. This weighed only 264 kg for the lander plus 10.5 kg for the rover. The development cost for the lander was less than $150 million. The development cost of  $150 million is quite low for a planetary mission. However,  privately financed it would be even less than that, perhaps as much as a factor of 10 cheaper.
  SpaceX was able to develop the Falcon 9 at a 90% (!) saving off  the development cost of a fully government-financed launcher of similar size. Planetary Resources Inc. plans to produce imaging satellites at a fraction of the usual cost.
Arkyd 100 space telescope.
 With mission costs this low we might want to produce a separate orbiter mission at Europa. As models for this we might use the small Mars orbiters Mars Odyssey and Mars Climate Orbiter.

Falcon Heavy sized launcher.

 The Falcon Heavy will allow a larger payload to be transported to the Europan surface. It is expected to be able to carry 53 metric tons to LEO. We'll use two Centaur stages together for the injection into a trajectory towards Jupiter. For the maneuvers at Jupiter, we'll use the storable propellant stage Delta K. This has a 6 mT propellant load, 0.95 mT dry mass, and 319 s Isp. Then it can provide a 1 mT spacecraft with the 3.9 km/s delta-v needed for the Europa landing:
319*9.81ln(1 + 6/(0.95 + 1)) = 4,397 m/s.
 The Centaur has several incarnations. It's propellant load is in the range of 20 mT, dry mass ca. 2 mT, and Isp ca. 451 s. Then two Centaurs together could provide the 6.3 km/s delta-v needed for the Jupiter flight carrying the 7.95 mT total mass of the Delta K stage + Europa lander:
451*9.81ln(1 + 40/(4 + 7.95)) = 6,500 m/s.
  The Falcon Heavy will cost in the range of $125 million. The Centaurs cost in the range of $30 million each and the Delta K stage costs ca. $4.35 million. Then the total launch cost will be in the range of $189.35 million.
 As a model of a 1 metric ton lander we might use the Mars Curiosity rover. This is a $1 billion mission however. The commercial space approach however should be able to produce a similar size spacecraft for a fraction of this cost.
   Bob Clark 

UPDATE, February, 20, 2015:

 The web traffic to the NASA web site for the Mars Exploration Rovers was even more extraordinary, measuring in the billions of hits:

NASA’s Web Site for 2005
By Digital Trends Staff — January 7, 2005
The U.S. National Aeronautics and Space Administration Web portal continues to drive high traffic numbers — more than 17 billion hits in 2004, report both NASA and Speedera Networks, a leading global provider of on-demand distributed application hosting and content delivery services. Speedera delivers content from the space agency’s portal to visitors seeking access to the site from around the world. Popular events on the NASA Web site, including the ongoing Mars Exploration Rover mission entering its remarkable second year, as well as upcoming major projects such as the launch and comet encounter of NASA’s Deep Impact satellite mission in 2005, are expected to drive continued high levels of traffic, according to NASA officials.

 It was estimated there were 142 million visits to the site during this period. So the question is how much advertising could be sold for a site this well visited?

Monday, February 2, 2015

Ariane 5 Core plus 4 Ariane 4 side-boosters as a manned launcher.

Copyright 2015 Robert Clark

  In the blog post "Ariane 4 for European manned spaceflight", I suggested resurrecting the Ariane 4 to produce a European manned launcher. In the blog post "A half-size Ariane for manned spaceflight", I suggested using a half-size Ariane 5, so as to be loftable by a single Vulcain 2, to serve as a manned launcher.

 Here, I'll suggest using a full Ariane 5 core but using 4 Ariane 4 liquid side boosters for a manned launcher. I'll use the smaller Ariane 5 "G" version to be loftable by the boosters and the single Vulcain 2 engine. By the on the Ariane 5G, the core has a 12 metric ton (mT) dry mass and a 158 mT propellant mass. By the Astronautix page on the Vulcain 2 used on the core, it had a 1,350 kN vacuum thrust and 434 s vacuum Isp.

 For the liquid-side boosters from the Ariane 4, they had a 4,493 kg dry mass and 39,279 kg propellant mass. The Viking 5C engine used had a 752 kN vacuum thrust and a 278 s vacuum Isp. Plugging these numbers into Dr. John Schilling's Launch Performance Calculator gives the results:

Mission Performance:
Launch Vehicle:   User-Defined Launch Vehicle
Launch Site:   Guiana Space Center (Kourou)
Destination Orbit:  185 x 185 km, 5 deg
Estimated Payload:   15537 kg
95% Confidence Interval:   11748 - 20004 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 15.5 mT payload is surprisingly high. Note this is without even using an upper stage.

   Bob Clark

UPDATE, February 10, 2015:

 The liquid-fueled boosters used on the ISRO's GSLV launcher are virtually identical to those used on the Ariane 4, so could be used for the purpose. Indeed the Vikas engines used on the GSLV were copied from the Viking engines used on the Ariane 4 as they were produced under license from the ESA:

GSLV Launch Vehicle Information.
Photo: Indian Space Research Organization

 Also, at a 15,500 kg payload capacity this Ariane 5 core plus Ariane 4 side-booster launcher could loft the Sierra Nevada Dream Chaser, at 11,300 kg.

 European companies are researching with Sierra Nevada the possibility of using the Dream Chaser as a manned spacecraft.

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.

     Bob Clark

UPDATE, February 10, 2015:

 In regards to the question of the suitability of the Ariane 4 for manned missions, i.e., whether it could be man-rated, note that it was considered for the purpose in the 1980's:

Multi-Role Recovery Capsule - BAe,1987.
Credit: NASA via Marcus Lindroos
    British manned spacecraft. Study 1987. Britain was the only European Space Agency member opposed to ESA's ambitious man-in-space plan, and the British conservative government refused to approve the November 1987 plan.
    However, the British aerospace industry did propose some interesting alternatives, such as the $2-billion 'Multi-Role Recovery Capsule'.
    British Aerospace Ltd. (BAe) regarded the French Hermes mini-Shuttle as too expensive and complicated. Instead, they felt a simple crew capsule would make more sense as an 8-man 'lifeboat' for Space Station Freedom (NASA issued a competitive request for proposals in late 1987). MRC was to be launched on the existing Ariane-40 rocket and the capsule could be flown manned or unmanned, for sensitive microgravity experiments in orbit.

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.

  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 .

 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