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Saturday, October 25, 2014

Altitude compensation attachments for standard rocket engines, and applications.

Copyright 2014 Robert Clark

Advantages of Altitude Compensation.
 Methods of altitude compensation such as the aerospike or aeroplug have been investigated for decades now. The idea behind altitude compensation is that rocket engines get their best performance at high altitude, in near vacuum conditions. Because of the physics, this will be when they use large nozzles. However, such large nozzles can not be used on the ground because they can cause dangerous flow instabilities that can actually rip apart an engine. 

 Then rocket engineers use a nozzle of a compromise size for engines that need to operate at sea level, one that is short enough to operate at sea level but can get moderately good performance at high altitude, in near vacuum. So this compromise reduces performance both at sea level and in near vacuum. The design of the aerospike is to recover that performance by emulating a short nozzle at sea level and a long nozzle at high altitude. 

 A disadvantage though is that it requires a toroidal combustion chamber or numerous small engines arranged around a central spike that can act as a toroidal chamber.

Example of an aerospike nozzle with a subsonic, recirculating flow [from Hill and Peterson, 1992]

  This requires a whole new design for an engine. Better would be if we could just make an attachment onto already existing engines that would give them altitude compensation abilities. One possibility is already being used now but only on upper stage engines. It uses an attachment of a long nozzle to the engine that is retracted while the upper stage is not firing, but extends after stage separation just before the upper stage engine is ignited.

RL10-B2 engine

 However, the purpose of this retractable nozzle extension is not to do altitude compensation but to have a small enough engine that can fit within the upper stage. It is not made to extend while the engine is firing but only before ignition. But according to noted space historian Henry Spencer, Pratt and Whitney tested it while the engine was firing and it worked. Then this could be used for altitude compensation where it extends while in flight while the engine is firing.

 Another possible way this would work be an inflatable nozzle such as investigated by a Goodyear aerospace division back in the 1970's.

INVESTIGATION OF EXTENDABLE NOZZLE CONCEPTS.
Charles N. Scott, Robert W. Nordlie, William W. Sowa 
Goodyear Aerospace Corporation, Akron, Ohio 
Final Report GER 15240 
November 1972

 Another discussion of it appears in this report that discusses both the aerospike and the inflatable nozzle:

NASA TECHNICAL MEMORANDUM. 
N73- 12840 
NASA TM X-64690 
August 1972 
CHEMICAL PROPULSION RESEARCH AT MSFC.


 This uses woven metal strands to form the high temperature inflatable shroud. Another approach would be to use the high temperature ceramic material used with NASA's inflatable heat shield.

The Inflatable Re-entry Vehicle Experiment (IRVE-3) is an inflatable heat shield effective at hypersonic velocities.

  
  Another possibility would be the high temperature ceramic discovered by mathematician/engineer GW Johnson. According to Johnson it is extremely lightweight:




 BTW, instead of it being inflatable it may work for the extendable nozzle to be folded up and gradually extended by mechanical actuators as the rocket gains altitude.

Altitude Compensation for Multi-Stage Rockets.
  A remarkable aspect of the Isp of a rocket engine is that a small increase in Isp can have a large effect on the payload. For instance a rule of thumb among rocket engineers is that every 10% increase in the Isp results in a 100% increase in the payload,[1]. The feeling has been though that altitude compensation was only useful for SSTO's and since SSTO's weren't being developed altitude compensation was not further developed. This is unfortunate because in point of fact altitude compensation can improve performance even for multi-stage rockets. For instance as discussed in The Coming SSTO's: Falcon 9 v1.1 first stage as SSTO, Page 2, using altitude compensation on the Falcon 9 v1.1 first stage can improve the payload by about 25%. For a 16,600 kg payload for the expendable version of the F9 v1.1 this would put it in the 20,000 kg range.

 Note that the 100% increase in payload using altitude compensation for a single stage vehicle compared to the 25% increase for a multistage has importance in relation to the usefulness of SSTO's.  In the "The Coming SSTO's: Falcon 9 v1.1 first stage as SSTO, Page 2" post I calculated that taking further into account the fact that a reusable first stage has to reserve propellant to return to the launch site, thus losing payload, a reusuable SSTO can get better price per kilo than a reusable multi-stage vehicle.

 However the 25% increase in payload for a multi-stage vehicle has importance to their use as well. For the the F9 expendable this would put it in the payload class of the Ariane 5 at ca. 20,000 kg. The F9 is a much cheaper rocket than the Ariane 5 at $56 million compared to $200 million, but the Ariane 5 has an advantage in being able to lift heavier payloads. If this relatively low cost addition to the F9 would give it the same payload capacity as the Ariane 5 but at only one-fourth the price, then the Ariane 5 could become completely obsolete.

 However, such altitude compensation should also be applied to the Ariane 5 as well. Assuming this would also increase its payload by the approx. 25%, this would give it a payload in the 25,000 kg range. The importance of this is that the ESA intends to spend ca. $1.35 billion to develop the Ariane 5 ME to give the Ariane 5 an additional 20% payload capacity. Then a just an altitude compensating nozzle attachment to the core stage and side boosters engines could accomplish that or better at a much lower cost.

 Only one company seems to have realized the usefulness of altitude compensation even for multi-stage rockets and that is the smallsat launcher start-up, Firefly Space Systems. They propose to use multiple small engines arranged around a central spike or plug. 


Credit: Firefly Space Systems.

 This method could also be used on the F9 first stage though you would have to remove the center engine. This would appear to reduce the thrust on the first stage. But at least in regards to the vacuum thrust the loss could be minimal. The reason is by using altitude compensation you improve the vacuum Isp and thrust of the engines. For instance if the Isp can be increased to the 340 s Isp of the Merlin Vacuum compared to the usual 311 s Isp of the Merlin 1D, then removing the center engine reduces the thrust by a factor of 8/9 but the altitude compensation improves the thrust by a factor of 340/311 resulting in a change of thrust of (8/9)*(340/311) = .97, only a small reduction.

 However, it is known that the aerospike does not fully recover the performance of a full vacuum-optimized bell nozzle. Then methods that use an adaptive nozzle might be preferred. However, each of the 9 engine nozzles expanding to an vacuum optimized nozzle probably wouldn't fit within the F9's diameter. Another way it could be done would be to use a single large nozzle for all the engines. According Phil Bono, progenitor of so many SSTO concepts, this could actually improve your Isp:

Encyclopedia Astronautica

Chamber/single nozzle.
http://www.astronautix.com/engines/chaozzle.htm

 This would appear to increase the stage weight in having such a large nozzle but actually you either remove the original nozzles entirely or cut off a significant fraction of their length under this proposal. Either of the methods of having adaptive nozzles of having a nozzle extension that moves into place at altitude or an inflatable nozzle could be used in this scenario.

 A recent report also proposes using a single nozzle for several engines:


Epitrochoid Power-law Nozzle Concept for Reducing Launch Architecture Propulsion Costs.
http://www.dtic.mil/cgi-bin/GetTRDoc?Location=U2&doc=GetTRDoc.pdf&AD=ADA533330

 The title refers to the lobed shape of the nozzle. According to the authors this shape improves performance.

Some New Proposals for adaptive nozzles.
 The adaptive nozzle that uses a nozzle extension only has two settings. On the other hand the adaptive nozzle that uses an inflatable nozzle only has a conical shape, not the bell shape that optimizes performance. Ideal would be a nozzle attachment that would be maintain a bell shape from ground to vacuum and would also change size in accordance with the surrounding atmospheric pressure. 

1.) One possibility would make a slight adjustment to NASA's conical inflatable heat shield. Notice it consists of a series of increasing diameter inflatable tubes.




 Then we could obtain the bell shape by moving the tubes slightly downward and inward so that a bell-shape is obtained rather than a conical one. 

2.)Another possibility is suggested below. 

Novel high temperature carbon nanotube ‘rubber’ for adaptive rocket nozzles(patent pending). 
 Recently a high temperature ‘rubber’ was discovered using carbon nanotubes, [2], [3]. It maintains its elasticity at up to 1,000 degrees C. The idea would be to produce a nozzle of variable shape by use of this high temperature elastic material.

  BTW, another application of the variable nozzles would be to produce highly throttleable engines. This would be especially useful for stages intended to be reusable to allow them to hover. The variable nozzle could be used to reduce the exit size of the nozzle to reduce the thrust.

  The proposal would have various means of achieving this variability. One method would have it be hollow filled with either gas or liquid. Then change in size would be accomplished by varying the amount of contained fluid. 

  However, nozzles should have a certain curvature, i.e., bell shape, to optimize the exhaust flow velocity and direction. Simply increasing the fluid content would just lengthen the nozzle, without maintaining the optimal shape. Then the proposal is make the nozzle of varying wall thickness so that the thinner wall sections would lengthen more and the thicker parts less, thus resulting in a curved shape.

  This method of using fluid to change the shape would also have the advantage that the fluid could be made to flow within the hollow nozzle to cool it if needed. However, the adaptive nozzle could be attached to the bottom of a regular, short, first stage nozzle. As exhaust gases reach the end of a nozzle, they cool. Then with the high temperature resistance of this material no special extra cooling may be needed.

  An alternative method of using this material for an adaptive nozzle would be to simply stretch it over a metal framework in the shape of an extended nozzle. Then actuators would be used to pull the elastic material over the framework, thereby maintaining the optimal shape.

  However, as described in the research reports on this nanotube rubber, it is susceptible to burning in an oxygen atmosphere as the carbon nanotubes themselves are. One approach to address this issue would be to apply coatings to it to prevent oxidation or burning as done with the carbon-carbon composite leading edges on the space shuttle wings and with graphite nozzles on solid rocket motors. A question here though is whether the coating would be susceptible to cracking when the nozzle is stretched or compressed, though this might be addressed by actually infusing the coating throughout the material during the formation process.

  Another approach would be to adapt this method of producing high temperature rubber using nanotubes or nanofibers to using different materials other than carbon. For instance it has long been known such nanoscale tubes or “whiskers” can be made of metals such as iron and tungsten. Nanotubes have also been made of boron nitride and silicon, which might be used for the purpose. The carbon nanotube rubber obtains its elasticity from the multiple connections and reconnections formed among the individual nanotubes as the material is stretched and compressed. Then the same principle may work using nanoscale fibers of materials other than carbon.

  There is an analog to this in a recent development involving aerogels. A NASA research team wanted to produce aerogels like those used on the shuttle insulating tiles but with higher strength and more flexibility. An approach that worked was to use polymers to form aerogels replacing the silica commonly used in aerogels, [4].

3.) Another approach would be to maintain the gradually increasing diameter of a bell nozzle internally:
Altitude compensation nozzle by internal adjustable spike(patent pending).
 Firstly, another problem with the aerospike is that it has to do the pressure variation all the way from the 100 to 200 bar combustion chamber pressure to the pressure of the vacuum. It would be simpler if it only had to do this from, say, atmospheric pressure to the vacuum.

 Then we will attach our altitude compensation extension to the bottom of a regular nozzle, not to the bottom of the combustion chamber. The method will use a widened bell shaped extension, wider than a usual bell nozzle of comparable size. But inside there will be a variable position or expandable spike. This spike will be moved or expanded as the altitude changes to obtain the correct area ratio for that altitude.




  The appearance would be like an aerospike pointing inwards towards the engine instead of outwards. The spike would be shaped so that as it is either moved up or down or expanded in or out, it would maintain the desired area ratio by the area between the outer bell-shaped nozzle and the inner spike.

 As indicated there are two methods being considered for varying the area ratio. One by moving the inner spike in and out, and secondly by expanding/contracting it.  For this second method there are a couple of ways to do it. You could have it be filled with a fluid and expanded or contracted by varying the amount of fluid.  Or you could have it in the form of an expandable structure such as an umbrella.

 All of these methods would require a temperature resistant material for the spike. There are various high temperature canvas-like materials that can be used,for instance, the materials currently being investigated by NASA for inflatable heat shields. Another would be the tufroc material used on the X-37B. Still another might be the toughened ceramics being studied aerospace engineer G.W. Johnson. Lastly, what might also work would be the high temperature carbon nanotube rubber-like material recently discovered discussed above.

 At first glance the proposal of having an internal spike may appear to be the same as the expansion-deflection nozzle. The Skylon team for instance intends to use an expansion-deflection nozzle of their engines. A study though of the exp.-def. nozzle showed it not to have very good altitude compensation capacity. 


Expansion-deflection nozzle flow behavior at low altitude [from Sutton, 1992]
 However, the key difference is that here the spike would be shaped to give the correct area ratio as it is moved during the flight corresponding to the ambient pressure unlike the pintle used in the exp.-def. nozzle.

 Another advantage is that the shocks could be shaped or even canceled out by techniques such as the “Busemann biplane” method. This could result in increased efficiency of the nozzle.



Bob Clark

REFERENCES
1.)Discovery of New Molecule Could Lead to More Efficient Rocket Fuel.
 Dec. 22, 2010
 http://www.sciencedaily.com/releases/2010/12/101222071831.htm

2.)Carbon Nanotube Rubber Stays Rubbery in Extreme Temperatures.
Liming Dai
Angew. Chem. Int. Ed. 2011, 50, 4744 – 4746
http://case.edu/cse/eche/daigroup/Journal%20Articles/2011/Dai-2011-Carbon%20Nanotube%20Rubb.pdf

3.)Carbon Nanotubes with Temperature-Invariant Viscoelasticity from –196° to 1000°C.
Ming Xu1, Don N. Futaba1, Takeo Yamada1, Motoo Yumura1, Kenji Hata
Science 3 December 2010:  Vol. 330  no. 6009  pp. 1364-1368
http://www.sciencemag.org/content/330/6009/1364.abstract

4.)Flexible, high-strength polymer aerogels deliver "super-insulation" properties.
By Brian Dodson
September 27, 2012
http://www.gizmag.com/polymer-aerogel-stronger-flexible-nasa/23955/





UPDATE, October 29, 2014:

 Another proposed idea for an adaptive nozzle that could be attached to an existing engine involves shutters on the nozzle that could be opened on closed depending on the ambient atmospheric pressure:

Rocket motor thrust nozzle with means to direct atmospheric air into the interior of the nozzle
US 3469787 A.







Saturday, September 27, 2014

Falcon Heavy for Orbital Space Tourism.

Copyright 2014 Robert Clark


 A space tourism study concluded that in order for orbital tourism to take off, so to speak, the price would have to come down to $500,000 per person:

How the Space Tourism Business Could See Orbital Boom.
Mike Wall, SPACE.com Senior Writer   |   April 25, 2011 12:32pm ET
 To date, only seven people -- beginning with multimillionaire businessman Dennis Tito in April 2001-- have paid to launch into Earth orbit, and they've reportedly plunked down between $20 million and $35 million for the experience.
Those are not the numbers of a thriving industry. But things could change dramatically if prices drop significantly -- down to about $500,000 per seat or so. That reduced rate could lure in hundreds of thousands of customers for orbital tourist trips, potentially generating revenues in excess of $100 billion per year, according to the study.

 Note this is for flights to orbit, i.e., to LEO, not suborbital flights as with Virgin Galactic. Interestingly the Falcon Heavy might be able to make that price point by carrying airliner numbers of passengers. The Falcon Heavy is slated to cost $125 million for a flight to LEO. Then at a $500,000 price point it would have to carry 250 passengers.

 You would need a passenger cabin to carry the passengers on the flight to orbit. As an estimate for a comfortable cabin size for a short travel time as with airliner flight we might compare this to the Boeing 757. In the 757-200 single-class configuration it can carry 239 passengers.

 It has a cabin width of 3.54 m and cabin length of 36.09 m, giving it a cabin volume of 355 m^3.

 
To carry the 250 passengers I suggest using the TransHab modified from carrying 4 to 6 crew for long space missions to having several rows and levels of seats for passengers on a 2 day or shorter trip to an orbital space station. The Bigelow space hotel BA 330 could also be used.

Transhab Module 
NASA's Johnson Space Center proposed a much larger and lighter inflatable 8-meter diameter "Transhab" module that also could be converted into crew quarters for future manned missions to the Moon and Mars. 
Credit: NASA via Marcus Lindroos.


 According to the NASA page on the TransHab it has a inflated volume of 340 m^3, about the same as the Boeing 757-200 cabin. Then slightly changing the seat sizes or aisle size we should be able to fit 250 passengers in the TransHab volume.

 For this many passengers we also need to calculate the amount of consumables needed, oxygen, food, and water. The book Expedition Mars estimates it as 5 to 10 kg  per person per day. Then this would be at most 2,500 kg for a 2 day flight. According to the NASA page on the TransHab, it weighs 13.2 metric tons. Allow also 100 kg per passenger and assume each passenger is allowed a space shuttle style orange flight suit at 12.7 kg and 10 kg for the seat. Then we're only up to 46.3 metric tons, within the 53 metric payload capacity of the Falcon Heavy.

 However, it would be rather uncomfortable in that cabin  for that many people for two days even if allowed to float around the cabin once reaching orbit. We may want instead to do the Soyuz flight method that is able to cut the flight time from launch to ISS docking to only 6 hours.

 But where to go once reaching orbit? We would need space hotels already in orbit. I'll estimate each passenger having his own 10 m^3 cubicle with shower. This would be about the size of a room 6ft x 8ft x 8ft. The TransHab at 350 m^3 could hold 35 of these cubicles. Eight of the TransHabs could hold enough cubicles to house the 250 passengers. In deflated form, the eight TransHabs could be launched in 2 flights of the Falcon Heavy.

 For the cost for these TransHabs, the Astronautix page on the TransHab gives a price in 1998 dollars of $100 million:

Encyclopedia Astronautica.
Transhab Module.
 American manned space station module. Cancelled 1998. Cost overruns soon forced NASA to consider other options for the International Space Station's habitation module. The space agency originally intended to use the same 8.2-meter long habitation module as the final 1991 Space Station Freedom design. In late 1998, NASA's Johnson Space Center proposed a much larger and lighter inflatable 8-meter diameter 'Transhab' module that also could be converted into crew quarters for future manned missions to the Moon and Mars. It was also possible that the module could be built and paid for by private industry and leased to NASA, although the exact configuration wasn't clear. Transhab and the 8.2-meter module appeared to be equally expensive ($100 million in 1998) and NASA had not made a final decision.
Article by Marcus Lindroos.
http://www.astronautix.com/craft/traodule.htm

 Keep in mind though this would be the price charged to the NASA, i.e., the government, for this module to be attached to the ISS. However, SpaceX and Orbital Sciences proved both with launchers and space capsules that development costs could be cut by as much as a factor of 10 by following the commercial space approach. Consider also the production cost for each vehicle is always a fraction of the development cost. Then the cost for each of the TransHabs as privately financed could be under $10 million each.

 However,  even at a price of $100 million for the TransHab this might be feasible. For each of the 35 cubicles that would be $3 million that would have to be made up. Say the space hotel owner charged $100,000 per week for each cubicle. Then, at full occupancy, a TransHab could be fully paid for in 8 months.

 That would be quite an expense though if the cost for each TransHab really were $100 million each, $800 million total for eight.



   Bob Clark


Update, Sept. 30, 2014: 
 
The inflatable habitats by ILC Dover designed for lunar habitats might also be used:

Camping on the Moon Will Be One Far Out Experience.
02.23.07

http://www.nasa.gov/exploration/home/inflatable-lunar-hab.html


2nd Update, Sept. 30, 2014:

Bigelow Aerospace has suggested a price for its proposed Space Station Alpha at about three times the price point I was suggesting on a per volume basis. The Alpha station consists of two BA-330 modules. Bigelow is offering a 110 cubic meter space for lease for 2 months for $25 million. That amounts to about $280,000 for a 10 cubic meters space for one week.


August 05, 2014
Bigelow Aerospace is hiring and targeting Inflatable Space Station Alpha to start launching in about 2017 or 2018.
http://nextbigfuture.com/2014/08/bigelow-aerospace-is-hiring-and.html

  These prices though will undoubtedly come down as time goes on.

3rd Update, Oct. 2, 2014:

 The estimates of hundreds of thousands of passengers and over a hundred billion dollars in revenues per year at a $500,000 price point is extraordinary. To put this in perspective note this would be more than an order of magnitude higher then current launch revenues. See the graph of the satellite launch market here:

Satellite launch industry revenue worldwide from 2001 to 2013 (in billion U.S. dollars).
Revenue in billion U.S. dollars33.73.22.832.73.23.94.74.43.85.85.42001200220032004200520062007200820092010201120122013
01234567
http://www.statista.com/statistics/185969/worldwide-revenues-of-the-satellite-launch-industry-since-2001/

 Even if the estimate of a proposed passenger market is off by a factor of ten it would still double the current launch revenues. Launch providers have been criticized by knowledgeable observers of the industry as lacking in the will towards innovation, such as reusability. Their critique was that the launch companies reason "Why should we try to cut launch costs by, for instance, reusability when this would just shrink our yearly revenues?"

 However, this survey shows, if valid, that in fact they would increase the launch market better than ten times by finding a new market, paying passengers. Every launch company in the world should conduct independent surveys to verify the results of this study. If true, than every launch company in the world should convert to reusability to cut their launch prices at least by a factor of 5 to bring it in the range of the approx. $2,000 per kg range SpaceX is proposing to offer for the Falcon heavy. For otherwise, they would be ceding a $100 billion a year market to SpaceX alone.

 

Tuesday, September 23, 2014

A SpaceX Heavy Lift Methane Rocket, Page 2.

Copyright 2014 Robert Clark

 In the blog post A SpaceX Heavy Lift Methane Rocket I discussed some possibilities for a heavy lift rocket using SpaceX's planned methane engine, the Raptor. Those calculations were based on the initial released values for its thrust of 660,000 lbs, 300,000 kilograms-force. However, recently SpaceX has said the Raptor may have a thrust of 1,800,000 lbs in vacuum, 1,600,000 lbs at sea level. So I'll give some revised estimates for its payload. 

 As before I suggest using the same tooling for the core stage as that used for the SLS core to save on development costs. Corresponding to the new higher thrust of the Raptor I'll use the full tank size of the SLS core, which would hold 1,000 metric tons (MT) of hydrogen-lox propellant. Since methane-lox is 2.4 times as dense as hydrogen-lox, the SpaceX methane-lox core will hold 2,400 mT of propellant. The relative densities of methane-lox compared to hydrogen-lox are given in Table 1 of this report:

Alternate Propellants for SSTO Launchers.
Dr. Bruce Dunn
Adapted from a Presentation at:
Space Access 96
Phoenix Arizona
April 25 - 27, 1996

 In regards to the specifications of the methane rocket engine it should be noted the high vacuum Isp of 380 s  cited really would only be expected of a vacuum optimized nozzle. Such engines can not be operated at sea level. However, in an upcoming blog post I'll discuss some altitude compensation methods that will allow the engine to have this vacuum optimized Isp while still being able to operate at sea level. So I'll assume both the first stage and upper stage engines have the vacuum Isp of 380 s.

 As before I'll take the number of engines for the core as five and the propellant-to-dry mass ratio as 20 to 1. For the upper stage I'll take the propellant size as approx. 1/5th that of the core stage, at 500,000 kg with the same mass ratio, and use a single Raptor for the upper stage.

SSTO Case.
 We'll use again Dr. John Schilling's Launch Vehicle Performance Calculator. Input the thrust as the value of the vacuum thrust in kilonewtons of 5 Raptors as 5*8200 kN = 41,000 kN, and the Isp as the vacuum Isp of 380 s. Input the propellant mass as 2,400,000 kg and the dry mass as 1/20th of this at 120,000 kg. For the "Default Propellant Residuals?" option select "Yes", and for the "Restartable Upper Stage?" option select "No". Selecting "Yes" for this last option would reduce the calculated payload.

 Use the default altitude of 185 km. Select Cape Canaveral as the launch site with an orbital inclination of 28.5 degrees to match the latitude of the launch site. Then the Calculator gives the result:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  110395 kg
95% Confidence Interval:  75079 - 152433 kg

 So this larger version could give 110,000 kg payload as an SSTO.

Two Stage Case
For the two stage case, in the column for the 2nd stage, input the propellant mass as 500,000 kg and the dry mass as 25,000 kg. Input the thrust as the vacuum thrust of 8,200 kN, and the Isp as the vacuum Isp of 380 s. Then the calculator gives the result:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  196089 kg
95% Confidence Interval:  162727 - 236334 kg

  Close to 200 metric tons payload to LEO. Actually since the mass ratio of stages improve as you scale them up, quite likely the mass ratio will be better than just ca. 20 to 1, perhaps in the range of 25 to 30 to 1. This would then improve our payload to above 200 metric tons.

Cross-Feed Fueling for Multiple Cores.
 We'll just look at the 3 core case here. As described in the A SpaceX Heavy Lift Methane Rocket post, we emulate cross-feed fueling in the Schilling calculator by inputting for the two side boosters 2/3rds the actual  propellant mass, and also increase the first stage propellant mass by an additional 2/3rds. The rest of the specifications, dry mass, thrust, Isp remain the same. 

 So enter 2 as the number of side boosters. Then in the column for the boosters, input 1,600,000 kg for the propellant mass and 120,000 kg for the dry mass of the boosters. Enter the actual vacuum thrust of 41,000 kN and Isp as 380 s. Then the calculator gives the result:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Cape Canaveral / KSC
Destination Orbit:  185 x 185 km, 28 deg
Estimated Payload:  536113 kg
95% Confidence Interval:  449806 - 639526 kg
 Over 500 metric tons of payload! This rocket however would be truly massive at 3 times the mass of the Saturn V. This would likely require a new,expensive launch pad to handle a rocket this size.


  Bob Clark

UPDATE, October 25, 2014:

 Some methods of accomplishing the altitude compensation mentioned are discussed here:

Altitude compensation attachments for standard rocket engines, and applications.
http://exoscientist.blogspot.com/2014/10/altitude-compensation-attachments-for.html

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

Update, October 13, 2014:

 A correspondent to my Facebook page named David Whitfield suggested the possibility exhaust from the preburner alone for the Merlin 1D might be low enough to give the Falcon 9 first stage hovering capability. You might be able to use 1 to all 9 Merlin 1D preburners to provide the needed thrust.
 BTW, is this the same as the turbine exhaust that appears on the left on this image:

 UPDATE, October 25, 2014:

 Another suggestion for achieving low throttleability is to use variable size nozzles. This is discussed here:

Altitude compensation attachments for standard rocket engines, and applications.
http://exoscientist.blogspot.com/2014/10/altitude-compensation-attachments-for.html 


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