Saturday, November 29, 2014

A half-size Ariane for manned spaceflight.

Copyright 2014 Robert Clark

 The current agreed  upon design for the Ariane 6 is to use a slightly reduced in size Ariane 5 core with strap-on solid boosters about half-size to the solids used on the Ariane 5:

Ariane 6.

   I believe this is a preferred solution for the Ariane 6 than the version using all solid lower stages. For one thing, if SpaceX succeeds in producing a reusable first stage, then ESA can keep pace by making the core stage of the Ariane 6 reusable.

 My ideal solution however would have used two to three Vulcain engines on the core stage. This would have an additional advantage of being able to be used as a manned launcher with no solids attached:

Friday, March 29, 2013
The Coming SSTO's: multi-Vulcain Ariane.
Copyright 2013 Robert Clark
http://exoscientist.blogspot.com/2013/03/the-coming-sstos-multi-vulcain-ariane.html



Single-Stage To Orbit Case. Still we can get a manned launcher retaining a single Vulcain II on the core and shrinking the size of the stage, to half-size. As discussed in the "The Coming SSTO's: multi-Vulcain Ariane" post, the propellant mass of the Ariane 5G core is 158,000 kg, with a 12,000 kg dry mass. We may remove a forward skirt called the "JAVE" used to attach the solids to the Ariane 5. This massed 1,700 kg bringing the dry mass down to 10,300 kg. The propellant tank on the Ariane 5G weighed 4,400 kg. So half-size this will weigh 2,200 kg, bringing the dry mass down to 8,100  kg. 
 In Dr. John Schilling's Launch Performance Calculator, enter in now also 79,000 kg for the propellant mass, 1,350 kN for the vacuum thrust and 434 s for the vacuum Isp. Select Kourou as the launch site with a launch inclination of 5.2 degrees, to match the launch site latitude. The "Restartable Upper Stage" option should be checked "No" even for a single stage, otherwise the payload will be reduced. Then the calculator gives:


Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Guiana Space Center (Kourou)
Destination Orbit:  185 x 185 km, 5 deg
Estimated Payload:  2528 kg
95% Confidence Interval:  1064 - 4248 kg
"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters
This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.

  This payload of 2,500 kg is for a single stage to orbit vehicle. As discussed in the blog post "Budget Moon flights: lightweight crew capsule", this may be sufficient for a 3-person capsule to LEO. For instance the Cygnus capsule given life support may fit within this size range. I have discussed though an SSTO reaches its best performance when using altitude compensation: "Altitude compensation attachments for standard rocket engines, and applications".

 By using altitude compensation the vacuum Isp can be raised to 466 s and the vacuum thrust to 1,350 kN*(466/434) = 1,450 kN. Schilling's calculator now gives a result of: 

Mission Performance:                  
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Guiana Space Center (Kourou)
Destination Orbit:  185 x 185 km, 5 deg
Estimated Payload:  4544 kg
95% Confidence Interval:  2894 - 6480 kg
"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters
This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.



 As discussed in the "altitude compensation" blog post though characteristics of how the Schilling calculator makes its estimates may make it less accurate in a scenario using altitude compensation. A more accurate analysis that varies the Isp from ground to orbit may be needed in this case.

Two-Stage To Orbit Case.

 We can get a higher payload manned launcher by making it TSTO. We'll use the cryogenic upper stage the Ariane H10-3. The Astronautix page gives it a gross mass of 12,310 kg and dry mass of 1,570 kg, for a propellant mass of 10,740 kg. The Isp is listed as 446 s with a vacuum thrust of 62.70 kN. However, this extra mass for the upper stage would mean the single Vulcain II on the core could not loft it.


 Then we'll reduce the propellant load in the core stage. It might also work to run the Vulcain at some percentage above the rated thrust, or use a varied mixture ratio at launch compared to high altitude. But using a reduction of the propellant load method, we'll lessen the propellant in the first stage by the mass of the upper stage, so by 12,310 kg. This brings the propellant load of the first stage to 66,690 kg. There is about a 35 to 1 ratio of propellant to tank mass so this will reduce the tank mass of the first stage by 12,310 kg/35 =350 kg. Then the dry mass of the core becomes 7,750 kg.  Then the calculator gives:


Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Guiana Space Center (Kourou)
Destination Orbit:  185 x 185 km, 5 deg
Estimated Payload:  4891 kg
95% Confidence Interval:  3970 - 5982 kg
"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adapters
This is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.


 We'll also estimate the payload for the altitude compensation case. Again take the first case vacuum thrust as 1,450 kN and the vacuum Isp as 466 s. But also improve the thrust and Isp for the upper stage, The thrust becomes 62.70 *(466/446) =  65.5 kN, with vacuum Isp also 466 s. Then the Schilling calculator gives:


Mission Performance:
Launch Vehicle:  User-Defined Launch Vehicle
Launch Site:  Guiana Space Center (Kourou)
Destination Orbit:  185 x 185 km, 5 deg
Estimated Payload:  6075 kg
95% Confidence Interval:  5016 - 7334 kg

"Payload" refers to complete payload system weight, including any necessary payload attachment fittings or multiple payload adaptersThis is an estimate based on the best publicly-available engineering and performance data, and should not be used for detailed mission planning. Operational constraints may reduce performance or preclude this mission.

 Again however this estimate for the altitude compensation case would have to be confirmed with more accurate estimation methods.


  Bob Clark

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  
http://www.washingtonpost.com/news/business/wp/2014/11/03/two-pilots-who-were-close-friends-now-tied-together-by-one-fatal-flight/

 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.
http://exoscientist.blogspot.com/2014/01/transitioning-spaceshiptwo-to-liquid.html

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

    Bob Clark

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