Saturday, November 15, 2014

Altitude compensation to allow the use of American engines on the Antares rocket.

Copyright 2014 Robert Clark

 In the blog post "Altitude compensation attachments for standard rocket engines, and applications", I noted that the idea that altitude compensation was only useful for SSTO's prevented their implementation and therefore their usefulness for multi-stage rockets was not realized. 

 An example of this is Orbital Sciences Antares rocket. The failed flight of the Antares in October, 2014 put renewed emphasis on the choice of 1960's era Russian engines AJ-26/NK-33. It is understandable why they were used since on the key performance metric of Isp, at 330+ s they were significantly better than American engines, at ca. 300 s. 

 However, by using altitude compensation the Isp of the low performance American rocket engines can even exceed that of the Russian engines. Orbital Sciences has decided not to use anymore of the Russian-derived engines on the Antares, and therefore need a replacement engine. I suggest investigating altitude compensation attachments that can made to already existing American engines so would be relatively low cost to implement.

 One possible engine that could be used would be the Rocketdyne RS-27A. It is used on the venerable Delta II rocket. Rocketdyne claims a 100% reliability record for the engine. You would need three of them though at ca. 200,000 lbs. thrust to make up for the two AJ-26/NK-33 engines at ca. 300,000 lb. thrust.

 How high could we get with the Isp on the RS-27A using altitude compensation? At an area ratio of only 12 to 1, the RS-27A only gets a vacuum Isp of 302 s. To see how much better we can do with a larger nozzle, we might make a comparison to the Russian RD-58, which gets a vacuum Isp of 349 s by using a high area ratio of 189 to 1 with not a particularly high chamber pressure of 78 bar. A better comparison might be to the Russian RD-0124 with a vacuum Isp of 359 s, but at a high chamber pressure of 162 bar. Unfortunately the area ratio of this engine is not specified, but it is certain to be high since it is an upper stage engine.

 Actually for vacuum Isp, just having a high nozzle area ratio is more important than the chamber pressure, a high chamber pressure being needed to insure a high sea level Isp. As a point of comparison, the hydrogen fueled RL-10B2 has only a chamber pressure of 39 bar but by using a nozzle extension to bring the area ratio to 280 to 1, it gets the highest Isp of any chemical engine at 465.5 s.

 Support for the idea a high area ratio on a kerosene engine can get a vacuum Isp of ca. 360 s even with a low chamber pressure is provided by the Rocket Propulsion Analysis program. Using the free Lite version you can estimate some fairly accurate vacuum Isp's for rocket engines, the sea level estimates though for the free version being not so accurate. Here are results using the specifications given on the Astronautix page on the RS-27A:

  The "Optimum Expansion" Isp number I've found to be a relatively accurate estimate for the actual vacuum Isp of existing engines. By the way, the negative values for the "Sea level" Isp are coming from the fact there would be severe losses for a low chamber pressure engine using such a large expansion ratio nozzle.

 Now compare this to the results if the chamber pressure were say 160 bar:

 You see the large increase in chamber pressure only adds minimally to the vacuum Isp, though it would have a great effect on the sea level Isp.

  So we'll take the vacuum Isp of the RS-27A with an adaptive nozzle attachment as 360 s. Now to calculate how much payload we can get on the Antares with these new engines I'll use the original's dry mass and propellant mass specifications here: Antares Launch Vehicle Information. The dry mass  of the first stage is given as 18,700 kg and the gross mass as 260,700 kg. 

 The two AJ-26 engines weighed 1,200 kg each for a total of 2,400 kg. The RS-27A weighs 1,000 kg, So three will be 3,000 kg. So the dry mass raises to 19,300 kg and the gross mass to 261,300 kg. I am assuming the adaptive nozzles can be made lightweight so as not to significantly increase the engine weight. The three RS-27A's though will have a lower liftoff thrust than the two AJ-26's. To make up for that I'll use a higher efficiency upper stage such as the hydrogen-fueled Ariane 4 H10-3 rather than the solid Castor stage now used.

 Now consider that we are assuming our adaptive nozzle will allow near optimal expansion from sea level to vacuum. Then note the RS-27A is a later edition of the RS-27 where the area ratio was increased from 8 to 1 to 12 to 1 to improve the vacuum Isp. But this reduces the sea level performance. The sea level Isp and thrust were reduced from 264 s and 93,357 kilogram-force (kgf) for the RS-27 to 255 s and 90,770 kgf for the RS-27A. But considering our adaptive nozzle I'll assume we are able to also get the 264 s Isp and 93,357 kgf thrust at sea level or perhaps do even better with a shorter nozzle equivalent at sea level. 

  At a 93,357 kgf liftoff thrust the total thrust at liftoff would be 280,071 kgf. The H10-3 stage has a gross mass of 13,100 kg. Then the total mass without payload will be 261,300 kg + 13,100 kg = 274,400 kg. This would result in a rather low thrust/weight ratio at liftoff which will reduce payload capacity through gravity drag.

 A couple of ways to improve this liftoff T/W ratio. First note on the page on the Antares linked above the specifications include the thrust at 108% of the "rated thrust". This is rather common that an engine can actually operate at a few percentage points above its rated thrust. This is the case for example with the Space Shuttle Main engines. If the RS-27A with adaptive nozzles can operate at 108% of its rated thrust that would bring the sea level thrust to 302,476 kgf.

  Another way to improve the liftoff T/W would be to reduce the propellant load by say 20,000 kg. As we'll see below the payload would still be rather high.

 We'll use Dr. John Schilling's launch performance calculator to estimate the payload possible. Select the Wallops launch site in the calculator and input the "inclination, deg" as 38, to match the Wallops site latitude.

 The calculator uses the vacuum values for the Isp and thrust inputs. This will be raised to 360 s for the Isp with our adaptive nozzles. But note also this increase in vacuum Isp also results in an increase in the vacuum thrust by a factor of the ratio of the Isp's, that is, by a factor of 360/302. Then the three RS-27A with adaptive nozzles will have vacuum thrust (360/302)*3*1054.20 kN = 3,700 kN.

 Input also the specifications for the Ariane 4 H10-3 for the second stage in the calculator. The HM7-B engine used on that stage has a vacuum Isp of 447 s. Then the results are:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Wallops Flight Facility
Destination Orbit:  185 x 185 km, 38 deg
Estimated Payload:  9458 kg
95% Confidence Interval:  7735 - 11589 kg

 The estimate of 9,458 kg is nearly twice the payload of the current Antares. Notably though this is using the high efficiency hydrogen-fueled upper stage.

 To address the low liftoff T/W I mentioned one way was to reduce the propellant load by, say, 20,000 kg. Doing this results in a payload of:

Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Wallops Flight Facility
Destination Orbit:  185 x 185 km, 38 deg
Estimated Payload:  8764 kg
95% Confidence Interval:  7166 - 10736 kg

 Still a pretty high result.  

 A consideration in regards to the accuracy of this estimate however is the effect of the altitude-compensating high vacuum Isp compared to the assumptions that go into the calculator. The Schilling calculator takes the vacuum Isp and thrust as inputs and automatically takes into account the reductions at sea level. However, since it assumes it is using a fixed nozzle it would assume the sea level Isp and thrust are much closer to the vacuum values than they would be in this scenario. On the other hand the altitude compensating nozzle would not have the losses of a fixed nozzle. Then more accurate payload calculators that take into account the variations of Isp and thrust with altitude would need to be used to get a more accurate estimate of the payload to orbit.

  Bob Clark

Saturday, November 8, 2014

Safety problems in the flight procedures for SpaceShipTwo.

Copyright 2014 Robert Clark

 Recent reports are that the co-pilot on the failed SpaceShipTwo flight unlocked the feathering mechanism early:

Two pilots who were close friends, now tied together by one fatal flight.
By Christian Davenport and Jöel Glenn Brenner November 3

 This article says the co-pilot "realized his error" after unlocking the feather and tried to shut down the engine. But it could be he noticed the feather deployed when it shouldn't have even when unlocked, and he then tried to shut down the engine.

 The flight procedures were that the feather should be unlocked at Mach 1.4, not at the Mach 1 it was unlocked on this flight. However, it is not known how much this was explained to be a mission critical element to the pilots. It may have been this was simply treated as something to do to follow the set timeline. Were there training sessions where this was explained that if you do this beforehand it will lead to vehicle disintegration? It is hard to imagine the pilots would make that mistake if it were emphasized the mission critical importance of when the feathering was unlocked.

 This article also quotes a Scaled Composites pilot as stating that normally the co-pilot would announce when Mach 1.4 was reached and the pilot would acknowledge it and command the feather to be unlocked. However, tape of an earlier SpaceShipTwo flight shows this didn't happen on that flight either.

 From the audio you can hear that one pilot state he is unlocking the feather when the motor is still burning. The feather doesn't deploy, correctly, until it is commanded to do so later after the rocket has ceased burning:

SpaceShipTwo's Intense Rocket Ride - Tail View and Cockpit Recording | Video.
Published on Sep 6, 2013

A camera was strapped to the rear of the Virgin Galactic vehicle to capture footage of the rocket engines and feather system at work. The vehicles 2nd powered flight occurred on September 5th, 2013.

 However, it is not announced that Mach 1.4 has been reached when it is unlocked. It is simply stated the feather has been unlocked by one of the pilots and the other acknowledges it.

 A key problem from listening to the video is that the pilots are not calling out the speed and altitude at any time during the burn. The only time they call out the altitude is a few seconds after the engine cutoff when they are close to max altitude. Note that when landing jet airliners when speed and altitude are both critical to a safe landing the pilots are calling these out to ensure they are within the correct range. The pilots should also be calling out both speed and altitude during the engine burn of SS2 to insure this mission critical step of the unlocking is done only at the right time.

 Another problem with the flight procedures also becomes apparent from this video. During that flight in September, 2013, the feathering was unlocked at about 16 seconds into the engine burn, and the feathering deployed correctly only later after engine cutoff. 

 But in the failed flight the catastrophic unlocking occurred only 9 seconds into the engine burn. That leaves a scant less than 7 second window to perform this action of unlocking that will lead to mission success or complete destruction of the vehicle. It's very disconcerting to know this would be the procedure as well for the passenger carrying flights. 

 Since the unlocking at 9 seconds was too early the window is actually shorter than that perhaps only 3 or 4 seconds. Note you can't unlock too late either since you want to ensure the feathering mechanism will be available before engine burnout, when you reach max altitude, when the feather would be needed for landing. Since that safe window for unlocking is so short in just a few seconds, there should be multiple redundant checks to ensure it occurs at the right time. 

 Actually, I'm not really comfortable with it being that short. An advantage of using liquid propulsion is that they have higher performance than hybrids and you can take a longer, more leisurely flight to altitude. This would have the additional advantage that the passengers would not be subjected to as high g-forces as becomes apparent from the pilots voices in the September, 2013 flight.

 In an earlier blog post I noted using liquid propulsion would have allowed Virgin Galactic to reach suborbital flight earlier and more cheaply:

Transitioning SpaceShipTwo to liquid fueled engines: a technology driver to reusable orbital launchers.

 Then in additional to that, there are flight safety advantages to using liquid propulsion.

    Bob Clark

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.

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:

N73- 12840 
NASA TM X-64690 
August 1972 

 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.

 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.

 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

1.)Discovery of New Molecule Could Lead to More Efficient Rocket Fuel.
 Dec. 22, 2010

2.)Carbon Nanotube Rubber Stays Rubbery in Extreme Temperatures.
Liming Dai
Angew. Chem. Int. Ed. 2011, 50, 4744 – 4746

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

4.)Flexible, high-strength polymer aerogels deliver "super-insulation" properties.
By Brian Dodson
September 27, 2012

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, 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.

 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.

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.

  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.

 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.

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.

 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.