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Project Apollo - NASSP 6.4.3 released!
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31  Orbiter Mars Direct / Planning / Re: Computer on: July 20, 2006, 04:58:25 PM
I've always thought that a developed planetary society with approximately current technologies:

Each planet probably has a Clarke orbit (for Earth it is at 38768km, Mars is 17023km, Venus and Mercury are unlikely to have any both because they rotate very slowly and they are close to the sun (and have closer libration limits.)

Satellites that would RF to the surface would then transmit to another planet via laser, which would then RF the signal to the destination planet's surface.  These would gradually supplant the current set of synchronous earth satellites (very, very gradually mind you.)
32  Orbiter Mars Direct / Planning / Re: Computer on: July 20, 2006, 04:50:45 PM
Quote from: aftercolumbia
You'd obviously want a computer that can run Orbiter  Very Happy
wouldn't that be a bad omen if on mars encounter day you botch the simmed entry and landing?  Wink

No...the day before; encounter day would be too busy to so a sim.
33  Orbiter Mars Direct / Planning / Re: Mars ERV Truck (SP-100 Deployment Ideas) on: July 20, 2006, 04:45:27 PM
Terry, I don't know where you are getting the idea of orbital docking over Mars....this really has nothing to do with the truck at all.

 Bangs Head  Duh!  Yuck!  Zip Lip  What? I'm never going to figure out how this happens so frequently...I was comparing docking trucks with docking spacecraft...not suggesting it.
34  General Projects / Mars Challenger / Re: Generic Mars Ascent Information on: July 09, 2006, 12:59:46 AM
Additional information on other LV families:

Delta II (standard flared biconic fairing-2.9m diameter):
7326: 500kg; $40 mln. (?, uses Star 37FM)
7325: 550kg; $41 mln. (?, uses offloaded Star 48)
7426: 550kg; $44 mln. (?, uses extra strap-on)
7425: 650kg; $45 mln. (?)
7926: 900kg; $59 mln. (?)
7925: 1060kg; $60 mln. (Futron; LEO On The Cheap by John R. London III, MGS and MER injection mass quoted; confirmed from source docs.)
7926H: 1050kg; $64 mln. (?; not recommended; lousy stage points)
7925H: 1200kg; $65 mln. (Roving Mars by Steve Squyres, Going to Mars by Garfield et. al.)

Notes: All version use spin-stabilized unguided solid upper stage motors that spin up to 60rpm and back, putting violent loads on propellant management devices and requiring traps for despin if tanks are located with outlets on axis.  A dynamic balancing analysis and test is also required (probably adds about $2 mln. to the S/C development.)

Delta 7925H and 7926H are not available at Vandenburg, keep this in mind if you plan to use an ascent declination above about +-70deg.

Delta III is no longer available.

Delta IV
M: 2117kg; $90mln (? this is a completely wild guess at cost)
M+(4,2): 3278kg; $100mln (? - lighter/higher C3 payloads are more expensive on this booster)
M+(5,2): 3200kg?; $110mln (?)
M+(5,4): 3686kg; $120mln (?)
H: 7812kg; $240mln. (?, I got this number from somewhere)

Orbital Sciences LVs
Pegasus: 0kg; $15mln (Futron, Pegasus is incapable of ascending directly to high energy trajectories and has no standard options to do so.)

Taurus:
2130: 210kg; $20mln (?; 18g burnout accelleration)
2230: 185kg; $20mln (?; 20g burnout accelleration)
3113: 343.4kg; $25mln (?; 13g burnout accelleration)

These use both spin stabilized unguided upper stages; meaning payloads need to be dynamically balanced; on the 3113 (and probably on the others, which are deprecated from 1996 edition and may no longer be available), the payload provides command ignition of the Star 37FM fifth stage and SFTS beyond 3rd stage separation if required.

Minotaur: 0kg; not commercially available, no standard high energy options

Minotaur IV: ?kg; not commercially available, no standard high energy options; Sec 8 Optional Enhancements says it can take a Star 48B but has no performance information.
35  Meadville Space Center / Site News / Re: New Forum on: July 08, 2006, 11:53:23 PM
Thanks once again for making a good home forum for After Columbia Project
36  Orbiter Mars Direct / Development / Re: ERV Development Thread on: July 08, 2006, 11:51:08 PM
The latest pics are busted on my system Sad
37  Orbiter Mars Direct / Development / Re: Mars Direct Rover on: July 08, 2006, 11:48:00 PM
If it's inflatable, it'll wind up being round.  Pressure likes to inflate round shapes, I can't imagine it being easy to rigidize an inflatable box too well with internal pressure.  Also, if its inflatable, there is little need to worry about fitting the final shape into the lander, only its stowed shape.  An inflatable would wind up operationally resembling a tent-trailer.

I think it would be best for it to have its main power rely on propellant provided from the SP-100 fuel plant and use the solars mostly as a back-up, or as aux power if it uses internal combustion engines for locomotion (i.e. if electrical power is incapable of moving the vehicle, it would be best to use two engines so that you aren't stranded if one fails; Apollo limited themselves to walk-back distance in the event the LRV failed, so it not-quite-doubled walking radius.  We have a year and a half, not a day and a half, such a restriction would be a major pain.)
38  Orbiter Mars Direct / Development / Re: Spint Development Thread on: July 08, 2006, 11:33:08 PM
Hullo...is anyone still here besides me?
39  General Projects / The After Columbia Project / Brief Introduction to In-space Propellant Management on: July 04, 2006, 12:44:23 AM
The basic question is:

"My motor is counting on getting pure liquid propellant.  In space, the propellant likes to float around in clinging blobs and bubbles...how do I make sure I'm getting the blob instead of the bubble?"

The expert on this sort of thing is D. E. Jackle Jr., who wrote all of the PMD related Technical Papers at http://www.psi-pci.com, if you are desperate for answers on in-space propellant tanks, click there now.  If you just want a basic overall perspective:

The type of performances for PMDs (Propellant Managment Devices), depend on a number of factors:

- How much propellant do you need in a single off-axis maneuver
- How much accelleration is the off-axis maneuver going to generate when the tank is nearly empty
- From where are the off-axis maneuver accellerations going to be imposed?
- How long between off-axis maneuvers

Why do I keep saying "off-axis"...because an "on-axis" maneuver will pool the propellant right over the outlet, in which case your PMD needs are vastly simplified.

- For on-axis maneuvers, how much propellant will you need before the propellant pool settles?  From which come:
- How long will it take the pool to settle?
- How much propellant is your motor gobbling up in the meantime?

So...there are a bunch of demand options:

- Ignition supply (for on-axis maneuvers)
- Specific one-time demand system
- Specific refillable demand system
- Flexible demand system

The specific choices for devices depend primarily on the volume needed to fill the demand of an on-axis ignition or specific off-axis maneuver (most often station keeping(GEO), drag reboost(LEO), or minor course correction(Planetary).)  The usual "specific one-time" maneuver is despinning from the top of a PAM type upper stage or after the spin-stabilized firing of a solid-fuelled orbital insertion or apoapsis kick motor.  In increasing order of complexity (therefore cost and residuals) and decreasing order of cost and reliability:

- vanes: accellerations less than 7mm/sec2 (hydrazine), often used to refill...
- sponges: accellerations less than 70mm/sec2 (hydrazine), demand volumes less than 4L.  Can go with somewhat higher accellerations and lower demand volumes, or vice versa.  Vanes and sponges can get residuals down to 0.10-0.25%
- big control sponges: pool falls out under any significant accelleration, but comes back.  Used to control the propellant center of gravity and bubble position (Viking Orbiter: pointing accuracy for cameras and ability to vent pressurant if tank pressure is dangerously high)  Bigger sponge means more residuals (probably about 1%)
- Trough: any demand volume, accelleration limits depend on ability to refill after maneuver, but can be up to about 1000mm/sec2 (tetroxide)
- Non-refillable trap: one-time demand for off-axis maneuver.  Typically up to 500mm/sec2 (hydrazine).  At great expense, up to 5000mm/sec2 (hydrazine).  If off-axis demand exceeds 10% of tank volume, troughs and galleries become more practical.
- Refillable trap: repeated demand for off-axis maneuver, or ability to hold propellant for on-axis maneuver start while doing off-axis maneuvers from another supply.  Typically up to 500mm/sec2 (hydrazine).  At great expense, up to 5000mm/sec2 (hydrazine).
- Gallery: up to 100mm/sec2 (tetroxide); 1000mm/sec2 (tetroxide) at great expense.  Gallery type devices are often used inside traps and troughs.
- Gallery pickup: a bit like an inflexible version of the hose inside an aerosol can; there can be one or many and don't each have to have their own tube.
- Gallery channel: A vane-like arm with a channel covered by a screen or perforated sheet wrapping around inside the tank walls where propellants need to be picked up
- Gallery liner: a liner around the inside of a tank with perforated sheet or screen covered holes.  Performs like poo with lots of residuals (1-2% or more) and complex construction when used in large tanks.  Use only for small tanks where building channel and pickup type devices is a pain.
- Diaphragm: All of the liquid propellant is held behind a flexible rubber barrier with little to no gas in it.  Flexible demand to very high accellerations (i.e. 2g or more).  The downside is that the residuals quantity is second worst behind the gallery liner.  Diaphragm rubber is not compatible with nitrogen tetroxide, and so is generally used only in monopropellant systems (because tetroxide has only 27% of the surface tension of hydrazine, bipropellant tanks often have different PMD in the fuel and oxidizer tanks.)

Diaphragms are useless for vapour pressurized systems because the evolved pressurant would wind up inside it.  Other vapour pressurized systems must be careful not to evolve (boil) pressurant gas behind the last gas barrier in the outlet.  They also need to limit boiled vapour in front of the last gas barrier, in which case it could bubble through if there is too much (i.e. no or too small a liquid path to the outlet.)

So...if you're looking to put a propellant tank on your spacecraft, I hope this gives you a good idea of how to figure one out that can survive all of your spacecraft's needed acrobatics (the highest of which in Orbiter is probably the time-overaccellerated entry which can send you out of the solar system at a high multiple of the speed of lignt.)
40  General Projects / Mars Challenger / Generic Mars Ascent Information on: June 27, 2006, 12:39:34 AM
The information posted herein comes from http://www.aoe.vt.edu/~cdhall/courses/aoe4065/NASADesignSPs/tm208533.pdf ... a must read for all MDC competitors

Interplanetary Mission Design Handbook: Earth-to-Mars Mission Opportunities and Mars-to-Earth Return Opportunities 2009-2024 by L.E. George and L.D. Kos, NASA MSFC 1998.  It is basically a porkchop smorgasbord, enough numbers to make any space nut crack.

I've sorted out the basics for the convenience of anyone whose interested, including other MDC competitors.  I've ignored all of the "human mission" stuff and went with Type I trajectories both ways, with no official limits on transit times.  Lower energy Type II's were available for departure as standard ISRU missions carry liquid hydrogen, it may be desirable to get to Mars quickly where you can start using it rather than watch it boil away during the cruise.

Dec 2013-Jan 2014 Departure C3=10km2/sec2 gives a 24 day window, which should be plenty.  Arrival at Mars is about 11 July 2014.  First following departure from Mars is Oct 2015-Dec 2016 followed by arrival at Earth in July 2016.

Feb-Mar 2016 Departure C3= 10km2/sec2; 21 day window.  Arrival at Mars is about 19 August 2016.  First following departure from Mars is Feb-April 2018 with arrival at earth in Sept-Nov 2018.

This information pretty much ignores entry C3's, which you might want to consider if your concept uses aerocapture, orbital capture, or has a weakened or new type of thermal protection system.

Both missions produce a generic design requirement of 10km2/sec2 for Earth ascent and 7km2/sec2 for Mars ascent.  You can use these numbers in finding booster performance information and in designing your ERV booster.  Here's the applicable Atlas V performances (good numbers from www.ilslaunch.com), and costs (educated guesses):

401: 2700kg (eq. 1125m/s GTO); $90mln. (from MRO press release)
411: 3400kg (eq. 1125m/s GTO); $100mln.
421: 4000kg (eq. 1125m/s GTO); $110mln.
431: 4500kg (eq. 1150m/s GTO); $120mln.
501: 2400kg (eq. 1200m/s GTO); $110mln.
511: 3300kg (eq. 1175m/s GTO); $120mln.
521: 4000kg (eq. 575m/s GDO); $130mln.
531: 4600kg (eq. 675m/s GDO); $140mln.
541: 5100kg (eq. 675m/s GDO); $150mln.
551: 5300kg (eq. 675m/s GDO); $160mln.
HLV: 8800kg (based on interpolated 675m/s GDO and 2200m/s equiv. offload performance...in other words, an educated guess.); 240mln.

Geosync Transfer and Delivery Orbit equivalents are there mostly to ease my mind that the performances I read from the Interplanetary Graphs and the more extensive Geosync tables were equivalent.

The Atlas numbers work like this:
- First number is the fairing diameter (4 or 5m)
- Second number is the number of solid strap-ons (0 to 5)
- Third number is the number of RL10-A4 engines in the Centaur (1 for high energy missions; 2 for low energy ones.)

Extra: Performance information has been updated based on 10km2/sec2 instead of 11km2/sec2.  Some are 100kg or 200kg higher.  The

Proton and Angara are also at www.ilslaunch.com
Other boosters are available at www.sea-launch.com, www.spacex.com, www.boeing.com/delta, www.arianespace.com, and www.starsem.com; if you want to go really small, go to www.orbital.com...the Taurus does something like 250kg.

Greenstar, is on hold right now and will probably not be available for the MarsDrive Contest.  If it becomes "available" before the deadline I will post good 10km2/sec2 performance numbers because I don't think it would be fair to use it without offering it to other competitors.

For some reason, JPL seems to prefer the Atlas V to the Delta IV and had this trend in motion before 21 December 2004, when the first Delta IV-H suffered a cavitation glitch that fooled it into thinking it was out of propellant in its first stage 7 seconds before normal shutdown...leading to a ~400m/s loss of performance.  Lucky for them, it was not an operational satellite but an Air Force test vehicle.

Let the games begin... Cool
41  General Projects / Bluestar / Bluestar Overview...the Limbo tour on: June 25, 2006, 10:04:54 PM
There are two editions of the Bluestar crunched out, and one more is now under development:

- INSEA Bluestar (INSEA stands for International Space Exploration Alliance, an "alternate history" to the space race I started working on before founding ACP; more details in the Cyanstar Overview; between INSEA and ACP concepts, Bluestar is Terry's oldest concept at Meadville...but the unsuccessful Greystar development is actually a couple months older.)
- ISTS Bluestar (ISTS for International Space Transportation System, one of After Columbia's original two projects from February 2003, the other being the Sprint Program.)
- Ascent Bluestar

I crunched INSEA Bluestar in late 2004 to early 2005, shortly after Greystar crapped out on its numbers (I may have posted late 2005 for Greystar in the Cyanstar Overview...that would be incorrect and I will try to edit it.)  INSEA Bluestar entered service in 1990.  In addition to Cyanstar, which was the heritage for Bluestar's Booster stage, the Bluestar had Volkstar (a successful but uneconomical minishuttle that served from 1986 to 1991) and Yellowstar (from 1981 to 1986, this tested full size Bluestar Orbiter configurations by launching them from the 25 tonne lift Lilmax expendable, most similar to the Saturn IB of contemporary history.)  Bluestar, as ugly as the numbers turned out, was still feasible for the INSEA universe.  The ISTS numbers, which required pre-extant turbine engine technology, failed to attain orbit with a feasible Orbiter dry mass fraction.  ISTS Bluestar used the technical capabilities of the Cyanstar booster (which is probably an optimistic basis for the real world as well) and added oxykerosene engines to the Booster.  The required Orbiter mass fraction (some 16% with higher than SSME performance oxyhydrogen) from the resulting 1500m/s staging point is not feasible.

INSEA Bluestar entered service in 1990, burying the Cyanstar with a tenfold reduction in LEO cost/kg ascent costs.  At entry into commercial operations in 1992 (some commercial flights in 1991 were made subsidized and with limited free insurance from INSEA), a flight charter cost $2.5 million and lifted 8000kg.  By 1994, it was down to $2 million.  Today (in 2006), Bluestar Series II has come out and lifts up to 10000kg in tanker configurations for $1.5 million.  These seemingly $150 to $250/kg numbers are not quite as rosy as their first impression gives.  Bluestar's payloads and dunnage have to be extensively qualified for flight on the $800mln combined Bluestar vehicle with it's four crew members (two in the Booster, two in the Orbiter; up to two more can be carried for EVA and Space Crane support operations.)  This lead to a pattern of mass produced/mass derived payload bus hardware...which was conducive to Bluestar's main market: phone satellites...in the INSEA universe today, there would be no cellular phones...period.

Also, these prices are a la carte, and do not include dunnage, deployment, and processing services, which took until about 1996 to really catch up.  One of the more exotic ones was the growth of a family of upper stages and later reusable space tugs, to support the geosynchronous market.  Hazardous processing was moved to space stations, making payload qualification for ascent on Bluestar a whole lot easier.  Between fabrication cost savings, which hurt performance of the satellite, and the cost savings incurred directly and indirectly from Bluestar, I estimated geostationary communications would cost about half of reality's today, about $500,000 per transponder-year.  These transponders would offer about 50% greater EIRP performance and more flexibility (satellite lifespans of 10 years vs. reality's 15 allow more frequent upgrades...but require also more frequent replacement.)

INSEA Bluestar Booster:
A low wing delta with high bypass ramfans (turbofan bypass shuts down to allow the bypass afterburner to operate as a straight ramjet) and oxykerosene rocketry (two engines, two motors per engine for a total of four nozzles.)  It is equipped with Kreuger flaps and zap-flap like extensions on the elevons, allowing lower landing alphas and shortening of the landing gear.
800,000kg total (ramp weight)
150,000kg ascent fuel for airbreathing engines
130,000kg ascent propellant for oxykerosene climbout rockets
200,000kg Orbiter
320,000kg everything else (operational empty weight)
$500mln purchase cost

INSEA Bluestar Orbiter:
Alumina protected HL-10 based lifting body; composite integral hydrogen tankage dominates the structure.  Uses high performance two part nozzle oxyhydrogen engines similar in thrust to the J-2, but similar in Isp to the SSME (actually it gets 475sec as it operates in a vacuum.)  The wing loading is reduced so the Orbiter does not require columbium or RCC on its leading edges.  The L/D is crappier than on Shuttle Orbiter, leading to a 25deg glideslope; landing speed is about 140kt vs. Shuttle's 200.
200,000kg total (staging mass)
160,000kg ascent propellant for oxyhydrogen main propulsion
2,000kg orbital maneuvering propellant (tetroxide/UDMH for Mk.1, oxymethane for Mk. 2)
8,000kg payload
30,000kg everything else (entry mass)
$280mln purchase cost

The 40,000kg Bluedot expendable upper stage was intended to replace the Orbiter.  It was a failure, and I had decided that it was when I first thought of it.  The Bluestar Booster was designed around the original Orbiter's aerodynamics and structural interface.  In order to be cost effective, Bluedot had to be simpler.  In early 1994, a Bluedot failed to stage properly, and blew up shortly after staging.  The booster crew was somewhat shaken up, but the booster and crew were alright.  After consideration of Bluedot's threat to the booster and crew concluded that it was safe enough to continue testing Bluedot, another attempt was made late in the year.  The Bluedot stage experienced an unexpected boundary layer transition through the transonic portion of the climb and exploded on the booster's back.  Both booster crew members were killed.  Fearing a potential problem commen to the Bluedot and Bluestar Orbiter or Booster, the fully reusable Bluestar system was stood down for six months, devastating the commercial market. 

After determining the fullly reusable system was safe, with the help of four test flights, Bluestar commercial services were resumed in mid 1995.  Bluedot was determined to be far to expensive to operate safely, approaching the competing Lilmax in cost/kg...the expendable it was supposed to replace.  Bluedot was cancelled in 1996 after three test flights (one success in addition to the two above failures.)

A final note on INSEA Bluestar is that it was the first aircraft in the INSEA universe to require the pilots to get a type certificate "learner's permit" in the simulator before they were allowed to take the controls of a real one.

ISTS Bluestar (the "real" one) is very similar to the Bristol SpaceBus at http://www.bristolspaceplanes.com/projects/spacebus.shtml

Therefore, after ISTS numbers starting coming back saying, translated into layman's English, "uh, you're joking, right?"  I did a check of SpaceBus (the concept that gave me confidence the Saenger II type runway operated spaceplane was reusable) and found that it was 400m/s short of being able to make LEO with fairly optimistic assumptions about the rocketry Isp.  Dang.

This brings us to Ascent Bluestar:

Essentially, we are stuck with ISTS Bluestar level propulsion performance with the Booster, but there are a number of things that can be done to the Orbiter that can make it feasible without sacrificing reusability...the downside is that it needs to sacrifice runway operations.

The booster is essentially the same as the ISTS Booster, reshaped to take whatever new shape the Orbiter winds up with.  Because the Bluestar booster will still use airbreathing engines and wings for ascent lift, as well as wheel landing gear, they can't be removed without leading essentially back to Shuttle Phase B...one of the worst developmental concept disasters in history. (the existing Shuttle is quite a bit better, but still pretty bad.)  The most onerous components to the Shuttle's performance are its wings's, landing gear, and other hardware and decisions made to support gliding runway landing.  The answer is simply to drop these items from the Bluestar orbiter and use a more ballistic method of recovery:

The front-running Bluestar orbiter is now a big, unpiloted Sprint-like gumdrop...and the jury is still out on most of its details...especially since I haven't figured out how to get the girth of this thing to fit in a sleek Concorde-like booster.  The "Phase B" style approach of pad operated full reusability can be approached totally sans runway by using Greenstar's reusable first stage blocks under Ascent Bluestar's new gumdrop Orbiter, probably four blocks in a square arrangement.

So, as you can see, Bluestar, and with it the cost optimal fully reusable vehicle is not-even-close-to-literally up in the air.  The limbo will probably last the better part of a year, as Mars Challenger, Sprint, and Greenstar are all higher priority projects right now...not least because it turned out full reusability was such a pain.
42  General Projects / Mars Challenger / Mars Challenger Overview on: June 25, 2006, 08:57:25 PM
I've started the characterization of Mars Challenger, ACP's entry into the MarsDrive Consortium, but I have not selected a configuration or booster yet.  The front runners are

a) far too tentative to bother posting at this point
b) proprietary: I want to get a head start on developing my ideas before letting the competition see them

The contest rules are here: http://marsdrive.com/files/MDC_MSR_ISRU.pdf

I will say that I have selected the standard Sabatier cycle oxymethane propellant combination, the favorite of Martian ISRU concepts, but I'm also going to look up ethylene as a possible alternative fuel.  It would be a surprise if it passes muster in its production processes and performances.

Right now the MSR entry is my primary timegrabber, and the other projects are currently on hold (as much as is possible without leaving OMDP in the lurch.)

Terry Wilson
43  General Projects / Cyanstar / Cyanstar Overview on: June 25, 2006, 08:43:38 PM
Prior to STS-107 and the founding of the After Columbia Project, I worked on concepts under the still unpublished flag of the International Space Exploration Alliance, or INSEA, which is a fictional alternate space race, sort of a "what the space race should have been" rendition.  It's not perfect...I decided to make one concept that flunked detailed analysis in late 2005 into one of INSEA's mistakes...that one is Greystar, the 40,000kg payload edition of Bluestar, it suffers from the rather simple problem of being just plain too big, coming to a gross mass of 3400 tonnes and having a bunch of nettlesome area scale issues that required expensive solutions.

Anyway, this post is about Cyanstar.  The After Columbia/Ascent version of Cyanstar has not been detailed, however the INSEA version had some numbers worked up for it in late 2004.  The INSEA Cyanstar enters service in 1975, and is basically a classic aluminum constructed supersonic shoulder wing delta aircraft not too unlike the North American XB-70 Valkyrie or the Avro CF-105 Arrow, both of which enter service in the INSEA universe...XB-70 does not last very long because Cyanstar replaces it.  By 1975 the CF-105 Mk. 4 as published in the real book "Avro Aircraft and Cold War Aviation" (by Randall L. Whitcomb in 2001, no relation to '50s and '60's supersonic aerodynamicist Richard Whitcomb, BTW) is the highest performing aircraft, with ramjet speeds of up to Mach 4 and the P-13 ICBM interceptor.  The Cyanstar therefore has a healthier legacy of successful supersonic aircraft to build upon that the "real" Ascent version does not have.

The Cyanstar consists of two major components...which by quirk of my vocabulary, both wind up being defined as "booster" (i.e. INSEA "booster", a runway operated vehicle designed to begin the ascent of a multistage reusuable or partially reusable system, vs. ACP "booster", a launch vehicle including all of its stages and dunnage.)  Reverting then to INSEA terms, the booster is the supersonic aircraft which begins the flight, and the SULV (for Single Use Launch Vehicle; a term originally from Sierra On-Line's 1997 computer game Outpost 2: Divided Destiny.)

The booster is a typical aircraft of the day, constructed primarily of 2219-T6 aluminum and with smatterings of Inconel 718, 6AL-4V titanium and A-286 steel...basically the same stuff that the Shuttle Orbiter is made of, except the Cyanstar booster doesn't need tile or RCC.  It uses four afterburning turboramjets in three dimensional intakes that go critical off of wing leading edge dogteeth at Mach 2.5.  The booster's maximum speed is Mach 3.5, with the 1981 Cyanstar Mk. 2 booster capable of Mach 4.1 with materials and cooling improvements.  Its mass breakdown worked something like this:

Total: 500,000kg (Ramp weight)
Turbine Engine Fuel: 150,000kg
SULV: 100,000kg
Everything else: 250,000kg (Operational Dry Weight)

The Cyanstar Mk. 2 would be just a shade lighter.

The SULV launched a payload of 4800kg to orbit (in the case of the Union II piloted spacecraft, the ascent guidance was controlled by the spacecraft, and so could weigh 5100kg.)  The Cyanstar SULV staged from the Cyanstar at 1000m/s in a considerable climb, and proceeded to orbit using a pressure-fed oxykerosene first stage, a pressure-fed oxykerosene second stage, and a hypergolic (NTO/UDMH+Hydrazine 50%; better known as Aerozine-50) pressure-fed third stage.  The choice of pressure-fed was dictated by costs.  With turbopumps, Cyanstar would not have beaten its predecessor Kilder's cost/kg to LEO by enough to be worth doing.  The choice of three stages was dictated by performance: the Cyanstar bay imposed a rather inflexible and limited facilities environment upon it.  Because the SULV was volume limited, these performances did not improve with commercial operators on the Mk. 2, but the Mk. 2 was operated at derated capacity and saved a lot of money on the booster's maintenance and amortization (the Cyanstar was replaced in 1990 by the Bluestar, and so enjoyed a long commercial life.)  All of the SULV main structures were made of 2024-T4 and some GFRP (better known as "fiberglass".)

Military operators preferred denser solids and payloads, and could take advantage of the Mk. 2's performance.  The US Air Force was the primary military customer of the Cyanstar (only one until 1982, when USAF and RCAF began to allow exporting of a dumbed-down Mk. 1 as an export version; commercial operators were able to lease Cyanstar Mk. 1's prior to this.  The Royal Canadian Air Force was a strong presence in the INSEA universe during this period, as the public environment needed to allow INSEA to be successful would not have allowed the cancellation of the CF-105 Arrow and the fracas of the Glassco commission as early as they happened historically.  The RCAF did not use the Cyanstar, as its primarily defensive purposes didn't need an aircraft like this.)

The USAF designed an all-solid air launched ICBM to be launched from Cyanstar.  This ICBM was lighter, thus allowing more fuel to be carried for airbreathing propulsion, allowing it to remain on airborne alert.  From airborne alert, the missile's warhead could be delivered in 40 minutes, which involved 10 minutes of accelleration and climbout, and 30 minutes of missile flight time.  Scramble added about 30 minutes.  The warhead mass was in the 3000kg ballpark, for a performance similar to that of the Minuteman III (The Cyanstar missile wasn't much smaller.)  The Cyanstar remained in military service until the collapse of the Soviet Union in 1991.  It was retired because of its expense, which was unjustifiable in a conventional role (also Cyanstar was not a stealth aircraft and wouldn't have been very good at it.)

All of my INSEA concepts were adjusted to the year 2000 dollar, and so these costs come out as such:

Development: 12 billion from 1968 to 1976
A booster: $450 million
An SULV: $11 million (incl. charter to launch)
A scrub: from $500,000 (no worse than real launch vehicles)
A missile: I don't give a hoot  Cool
44  General Projects / Greenstar / Greenstar Overview on: June 25, 2006, 07:56:45 PM
Ascent: The First Step Into Space is a book under development by After Columbia (authored by Terry Wilson) with the Greenstar cost-optimized booster as its centrepiece.

Greenstar currently has the following technical decisions and assumptions behind it.  “Front runner” means that an item is possibly subject to change because it is not known for certain if it is optimal.  “Firm” means that it is certain that the item is optimal.  “Dogma” is an item not certain to be optimal, or perhaps known not to be, but is held firmly for perspective reasons (i.e. one of Greenstar’s goals is to find the pressure-fed cost-optimal configuration for a 5m diameter booster.  It is possible that a 5m turbopump-fed booster may be economical, but it is not Greenstar’s goal to determine that.)

- Pressure fed; no turbopumps (dogma)

- Pressure vessel safety factor 2.5, Structures safety factor 2.0; the rationale is to allow shipyard class tolerances rather than aerospace ones, and to reduce the need for and chances of delay during integrated ground testing. (firm)

- Material is currently implemented with a density of 7.8 g/mL and yield strength of 575MPa; assumed to be a quench and temper hardened steel.  No assumption is made on whether it is brittle at liquid oxygen temperatures.  The material has an allowable stress of 250MPa when used in tanks, and 287MPa when used in structures (front runner)

- First and Second Stages use Propulsion System, Lower (PS-L), currently operating at a tank pressure of 2500kPa (362psia) (front runner), with a motor chamber pressure of 2000kPa (290psia) (front runner); it uses segmented construction and tang and clevis joints equipped with inner/outer dual O-rings, LOX tank internal insulation (a simpler modification of the original SRB joint which would be unsuitable for SRB use for the same reasons that caused the STS-33 Challenger mishap on 28 January 1986; Greenstar provides much gentler operating conditions other than temperature.) (front runner)

- PS-L stages use a combination of angled verniers and differential throttling of both main and vernier motors for control in all three axes using moving components no bigger than globe or iris throttling valves. (firm)

- PS-L tankage takes the axial accelleration compression loads directly in the tank wall structure (firm)

- The first stage uses recovery parachutes to allow reuse (firm); at Urwumpe’s (Dennis Krenz) recommendation, they are now located in the forward skirt instead of in a piggy back pod (firm).

- The second stage is not recovered (front runner) but does possess enough thermal protection that it will impact the sea in one piece (firm).

- The third stage uses 1200kPa (174psia) tank pressure and 750kPa (109psia) motor chamber pressure (tentative) in an all-welded, post weld heat treated main structure and tanks (front runner)  This is called Propulsion System, Upper (PS-U)

- The first three stages use liquid oxygen and kerosene propellants (firm), mixed at 2.3:1 for PS-L stages (firm) and 2.2:1 for PS-U stages (front runner)

- PS-U stages use N2 cold gas RCS (firm; performance of monopropellant hydrazine sacrificed to eliminate the toxic hazard from ground safety consideration), plus assured liquid acquisition verniers using the main propellants to provide stage separation and settling thrust (front runner).  The main engine is position balanced and uses electromechanical gimbal actuation (front runner).

- The high energy top stage uses the Dunn Engineering self-pressurized hydrogen peroxide/propane propulsion system with 95% hydrogen peroxide mixed at 6.6:1 with a motor chamber pressure regulated at 750kPa (109psia) (front runner)  See www.dunnspace.com for details on self-pressurized propulsion system.  In Greenstar application, the pressure vessel will be a sphere.

- The top stage is designed with sufficient endurance to complete delivery to Geosynchronous Equatorial Orbit in anticipation of the demand for that service from lower cost satellites designed for the cheaper and friendlier launch market environment provided by Greenstar (firm, intermediate performances will still be available.)

- PS-U and PS-T both use structures (compression ring/interstage and fairing, respectively) which allow axial accelleration and aerodynamic compression forces to bypass the propellant tanks (front runner for PS-U, firm for PS-T)

- PS-L and PS-U vacuum optimized nozzles will be two part extendable (the lengths of the second stage skirting required to accomodate fixed nozzles created very unfavorable mass fractions and doubts as to whether the vehicle could handle the lateral loadings from wind sheer.)  PS-T and PS-L surface optimized nozzles are one piece. (firm)

- All stages use helium pressurization to provide a reserve pressure above saturation vapor pressure for vapor pressurized propellants (liquid oxygen and propane) to prevent line and manifold cavitation and as the primary pressurant for kerosene. (firm)

- The payload fairing integrates with the highest PS-U stage in the booster. (firm)  It is jettisoned at a deterministic thermal rate of 1000W/m^2 for trajectory/simulation studies, and with the second stage for delta-v optimizations.

- Errant Flight Termination is provided by intact abort, the FTS signal commands the Flight Termination System to do many things, only a few of which are necessary to ensure the booster is no longer capable of propelling itself to a course that will endanger the public.  It withdraws power from all valve power busses, disabling the guidance system’s ability to control the vehicle and forcing all valves, including pressure-venting valves, into a power-off configuration.  It commands closed all motor valves and a redundant set of FTS butterfly valves, which isolates all main and vernier motors from their propellant supplies.  The guidance system will continue to operate and telemetry transmitters will continue to transmit, as the information they are generating may be conducive to the failure investigation. (firm) An autonomous mode may be implemented for attitude limit and attitude rate (front runner; it may not be a good idea if launching from the Cape because of the possibility of an AFTS cut off while the vehicle’s Instantaneous Impact Point is in or too close to Africa.)  An autonomous mode which is activated by stage separation sensors and breakwires intended to detect vehicle breakup will be implemented (firm).

- As many pyrotechnic devices as possible will be eliminated from Greenstar for ground safety (and associated expense) reasons. (firm) Perhaps all of them can be eliminated (unlikely)

The Greenstar uses a blockable first stage in two lengths, the shorter of which is used only as a second stage.  The motors are common among all PS-L stages, and a PS-L stage can have one, three, or four motors, each version being equipped with an appropriate length nozzle.

The PS-U (third stage, fourth in very large configurations) stage comes in three sizes: four metre, five metre short, and five metre long.

The PS-T (fourth stage, fifth in one configuration) comes in two sizes, the smaller of which will fit under the four metre fairing.

The configuration numbering system has four digits, which are associated with the four stages of a high energy vehicle.  The first digit is the number of first stage full length blocks used (1, 3, or 5); the second digit is the size of the second stage (1 is half-length with one high expansion motor, 2 is full-length with three moderate expansion motors.)  The third digit represents the PS-U stage (’4′ is the four metre stage, ‘5′ is the short five metre stage, while ‘6′ is the long five metre stage.)  The fourth digit designates the fourth/fifth stage(s) (’0′ is no stage, and denotes a three stage low energy version of Greenstar, ‘1′ is the small PS-T stage, ‘2′ is the large PS-T stage, ‘3′ is the five metre short PS-U stage, and ‘4′ denotes a two stage stack consisting of the five metre short PS-U stage as the fourth, and the larger PS-T stage as the fifth.)

The first stage full length block is at a gross weight of 400,000kg, with the remaining stages sized around that.  It has shrunk somewhat because of the first stage areal thrust density issues of pressure fed vehicles of this size.  The rest of the stages will be sized on that assumption and these design point configurations:

1140 to a 60deg support orbit (The half length PS-L block actually contains about 1/3 the propellants of the full length block; both second and third stages respond to optimization)

3250 to a 60deg support orbit (here only the third stage responds to the optimization)

5260 to a 60deg support orbit (here only the third stage responds to the optimization)

1141 to 1500m/s GTO from a 60deg support orbit with a three maneuver mission plan

3252 to full GEO delivery from a 60deg support orbit with a four maneuver mission plan

5263 to a GTO with somewhere around 1500 to 2100m/s remaining delivery delta-v from a 60deg support orbit with a three maneuver mission plan

5264 to full GEO delivery with a four maneuver mission plan from a 60deg orbit

 These are the remaining off design point configurations, some of which will be eliminated from further consideration as development continues:

Low Energy: 1150, 3140, 3150, 3260, 5250

High Energy: 1151, 1152, 3141, 3151, 3152, 3251, 3261, 3262, 5251, 5252, 5253, 5254, 5261, 5262, 5263

Potentially Successful Wierdos: 1640, 1641, 1650, 1651, 3640, 3641, 3650, 3651, 3652 (These use the five metre long ‘6′ PS-U stage as a second stage.)

Payload envelope depth may be affected by the selection between the 5m short and 5m long PS-U stages because the compression ring for the PS-L forward skirt/interstage must begin at the top of the PS-U aft bulkhead and no further forward.  “May be” because the long PS-U stage may use a longer compression ring, but this hasn’t been determined yet.

The existing Delta II second stage provides an excellent example of a compression load bypass structure.  The compression ring interfaces forward with the fairing and aft with the first stage forward skirt.  This results in the trailing edge of the long second stage nozzle sneaking out of a very long tunnel reminiscient of the romantic undocking of the Starship Enterprise from its refit yard, only a heck of a lot faster.  Greenstar does not use this technique for 1/2 staging because there is still enough dynamic pressure to cause the first stage to rotate quickly after staging, a phenomenon counted upon to save the recovery parachutes from second stage exhaust.

45  General Projects / Sprint / Sprint Ferry Overview on: June 25, 2006, 01:41:45 PM
Not to be confused with After Columbia's original Delta Sprint at http://aftercolumbia.tripod.com/deltasprint/ ; Sprint is based on it, increasingly loosely.

The Sprint ERV Mk. 1 is currently under development by myself, Brian Heick and Seth Hollingsead at the Orbiter Mars Direct Program, which also resides at Meadville Space Center.

The Delta Sprint was intended to substitude the STS-107 stand-down by being a crew ferry for three members that could be developed on the same budget and schedule as the STS-33 stand down after the Challenger disaster (as Delta Sprint was characterized before the extent of the STS-107 stand down was known.)  The basic philosophy was to launch it on the most reliable US commercial booster that could be sought and then built to survive its failures as a commercial rated booster rather than man-rating it.  The booster selected was the Delta II.  Originally the Delta Sprint had intended to use the Fregat upper stage to assist its service module in rendezvous maneuvers, thus saving on Service Module development...then I found out that was a pretty dumb idea.  The lessons learned from Delta Sprint have been carried forward into the Sprint Ferry.

The Sprint Ferry amasses a total of about 8000kg at booster cutoff breaking down something like this (updates pending)

Orbital:
Descent Module: 5000kg
Service Module: 3000kg

Ascent:
Service Module Fairing: 1200kg
Payload Escape Stage: 1500kg

The current front runner for booster is the Greenstar 3150.  These are the other current front running details:

- Pressure Bearing Safety Factor: 2.0
- Compression Loads Safety Factor: 1.5
- X-axial Load Factor based on worst case survivable abort entry: 20g
- Service Module Main Propulsion: Diaphragm/regulated GN2 fed monopropellant hydrazine; four main motors
- Service Module RCS: same supply, 12 secondary thrusters
- Descent Module RCS: Diaphragm/blowdown GN2 fed monopropellant hydrazine; 16 thrusters linkable to SMRCS supply prior to SM separation
- Descent Module Payload Escape Stage: Four solid motors located under RCS pods in Heick 0606 model
- Service Module Backup Maneuver Motor: Solid motor, probably Star 30 class, located on the Service Module X-axis.  This motor provides backup deorbit capability to the SMMP and also provides the abort capability after the Payload Escape Stage has been jettisoned.
- Service Module constructed of 7075-T73 aluminum alloy, Descent Module structures made of 2024-T6 (machinability concerns, as DM main structure is very complicated.)
- Descent Module Thermal Protection System uses RCC for the base heatshield, and AFRSI throughout the rest of the craft.  These are backed up by ablatives to allow survival of up to an 50mm2 breach in the main material.
- The Service Module possesses a protective impact shield to guard the main heatshield of the Descent Module from debris hazards during an uncommanded abort.
- An Orbital Docking System compatible with the current Shuttle/ISS system is carried in the nose in clamshell doors.
- Recovery is by a ram-air inflated parasol (similar to parachutes and paragliders used by their respective sports, only much larger) and airbags.
- Backup is provided by water recovery, reserve parachutes and the David Clark S1035 Advanced Crew Escape System, the suit and individual parachute currently used by Shuttle astronauts during launch and entry.
- Design goal for the probability of all crew members surviving any given mission without major injury is 99.99%
- Expected probability for a mission being completed normally is 98.50%, 1.00% deficit from booster failure, the remaining 0.50% from the spacecraft and mission on orbit.

Terry Wilson
After Columbia Project
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