ESAS draft report is out!

[krrr]

I get a slightly different result using this:<br /><br />Orbital velocity of a 200 km circular orbit: 7784.3 m/s<br />Orbital velocity of a 1200 km circular orbit: 7252.5 m/s<br />Perigee velocity of a 200x1200 orbit: 8054.5 m/s<br />Apogee velocitiy of a 200x1200 orbit: 6991.7 m/s<br /><br />Hence total delta-v for a Hohmann manueuver = (8054.5-7784.3) + (7252.5-6991.7) = 531 m/s which is almost (but not quite) the same as the difference in orbital velocities.<br /><br />v_p = sqrt(2*GM*r_a/(r_p*(r_p+r_a))) <br /><br />v_a = sqrt(2*GM*r_p/(r_a*(r_p+r_a))) <br /><br /> <br /> <br />

 

[JonClarke]

"Current concepts of NTRs are also bi-modal. Which means the NTR not only propels the ship it also supplies the ship's electrical power thereby eliminating the mass of a more typical power system. "<br /><br />Yes, it's called a Brayton cycle generator. However it does not come for nothing. For example the 3 X 6.8 tonne thrust engines I described earlier will have a Baryton cycle generator massing 1.35 tonnes. The generator itself is bigger than needed because it has to run both the crew module and the systems the NTR stage needs like a propellant refrigeration unit.<br /><br />Remember if you commit to a Brayton cycle you are committed to lugging the reactor and all associated systems including largely empty tanks through every manouver. These means more manouvering propellant, RCS, etc.<br /><br />Remember that by the time the Mars mission is ready to go there already will be a fully developed chemical departure stage of proven reliability available at unit cost.<br /><br />Jon <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[JonClarke]

So far we have only a very brief summary is available of the proposed ESAS Mars mission architecture at http://www.spaceref.com/news/viewsr.html?pid=19066 .<br /><br />It may be worth noting two previous studies.<br /><br />The architecture of this proposal resembles the Mars “Dual Lander” mission proposed by B. G. Drake in an internal NASA memo in 1999. The most detailed summary I have found is at http://advlifesupport.jsc.nasa.gov/documents/JSC-39502A.pdf <br /> although this article deals mainly with life support issues. It too consists of the same prepositioned hab and DAV, and a crew MTV. The main differences is that it had an electric propulsion system for the two cargo missions and that all three missions are sent at the same opportunity.<br /><br />The spacrcraft designs closely resemble the BNTR/LANTR missions of Borowski et al., the artwork in this current summary looks very much liked they are reduced from the Borowski paper http://gltrs.grc.nasa.gov/reports/2002/TM-1998-208834-REV1.pdf . However the Borowski proposal is packaged into orbitally assembled units sized for the much smaller Magnum system and with a somewhat larger overal mission mass.<br /><br />Jon <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[krrr]

<font color="yellow">If the SDLV can carry 125 tonnes to a 400 km orbit, what can it carry to a 200 km and a 1200 km orbit, respectively?</font><br /><br />Roughly, 128.5 tonnes and 113 tonnes.<br /><br />Assuming a delta-v difference of 116 m/s between 200 and 400 km and of 416 m/s between 400 and 1200 km (and an Isp of 430 sec).

 

[gunsandrockets]

"but it is more efficient to "burn" propellant in low orbit as quickly as possible... That is why the parking orbit of spacecraft on interplanetary missions are as low as possible,"<br /><br />By interplanetary missions do you mean manned ships assembled in Earth-orbit at a 200 km altitude? If you have a link to any such plan out there I would like to see it.<br /><br />And it occurs to me that a low 200 km parking orbit would quickly decay and require even more stationkeeping propellent than a more typical 400 km orbit. I have heard the ISS expends 7 tonnes of propellent per year stationkeeping. In a higher orbit, such as the Soviet RORSATS used, fighting orbital decay for all practical purposes uses zero propellent.

 

[JonClarke]

Thanks! This makes single SDLV launches to Mars even with NTR and advanced materials very tight for a 6 person crew.<br /><br />Jon <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[mlorrey]

ISS needs that propellant because it is so huge with little density. There are plenty of upper stage sized vehicles in LEO that see little orbital decay, and a trans-martial upper stage is going to very dense, being full of fuel. It should see no orbital decay of any significance for a number of months.

 

[gunsandrockets]

"The generator itself is bigger than needed because it has to run both the crew module and the systems the NTR stage needs like a propellant refrigeration unit...you are committed to lugging the reactor and all associated systems including largely empty tanks through every manouver..."<br /><br />Tanks used for TMI are jettisoned when empty and don't need refrigeration even if they weren't jettisoned. Propellent refrigeration is only needed for the full propellent tanks carried to Mars. The very same conditions that applies to any chemical rocket engine that uses cryogenic propellent. <br /><br />And since the less efficient chemical rocket would need even bigger tanks than NTR, the chemical rocket would have higher power needs for refrigeration than NTR. Of course the chemical rocket could use non-cryogenic propellents but those are less efficient and would probably impose even greater mass penalties than a chemical rocket already suffers compared to NTR.<br /><br />

 

[JonClarke]

if you are using LH2 hydrogen boil off is a critical issue regardless of how it is used. So you are going to spend as little time as possible in LEO. Borowski allows for a 32 day loiter in 400 km earth orbit (reference already given). Two hundred km should be stable enough for this period, I would have thought. Version 1 of the DRM used a 220 km orbit. <br /><br />If you are in parking orbit for less time you can go lower. pollo missions used 190 km orbits (http://nssdc.gsfc.nasa.gov/database/MasterCatalog?sc=1968-118A). Some Venera missions used 280 km (www.astronautix.com/craft/venra1va.htm). Proton/Breeze parting orbits for GTO missions are as low as 160 km (www.space.com/missionlaunches/proton_launch_020330.htm), but they are only there briefly.<br /><br />Incidently, on the subject of save disposal orbits, Borowski et al. say that up to 2.5 km/s is need to for each NTR stage to put into a disposal orbit. So the total velocity penality for an NTR mission, including both higher starting and disposal orbits is ~3 km/s. <br /><br />Jon <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[gunsandrockets]

"ISS needs that propellant because it is so huge with little density. There are plenty of upper stage sized vehicles in LEO that see little orbital decay,"<br /><br />Empty balloon-tank upper-stages don't have much density themselves. Do you have any examples of stages at the same altitude as the ISS which linger?

 

[gunsandrockets]

"Version 1 of the DRM used a 220 km orbit."<br /><br />Version 1 of the DRM? Are you refering to Borowski? If you are refering to something else, link please.<br /><br />

 

[gunsandrockets]

"So the total velocity penality for an NTR mission, including both higher starting and disposal orbits is ~3 km/s."<br /><br /><br />No, since Borokowski assembles his rocket at 400 km, not at 1200 km as you insist must be the case. <br /><br /><br /><br />

 

[JonClarke]

DRM version 1 http://ares.jsc.nasa.gov/HumanExplore/Exploration/EXLibrary/docs/MarsRef/contents.htm - note that it only became known as version 1.0 on release of later iternations.<br /><br />Jon <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[mlorrey]

uh, yeah, a quick scan of Astronautix.com should turn up lots of examples. Looking at the Agena upper stage, for instance, which reached 400-550 km orbits, kept those orbits for 12-14 years.<br /><br />Here is an example of one in a lower orbit:<br />1962 Mar 7 - 22:10 GMT - Launch Site: Point Arguello . Launch Complex: LC1-2. Launch Pad: LC1-2. Launch Vehicle: Atlas LV-3A / Agena B. Model: Atlas Agena B. LV Configuration: Atlas Agena B 112D / Agena B 2204. <br />Samos 6 Nation: USA. Payload: Samos E-5 no. 3. Mass: 1,860 kg. Class: Surveillance. Type: Military. Spacecraft: Samos. Agency: U.S. Air Force. Perigee: 236 km. Apogee: 686 km. Inclination: 90.9 deg. Period: 93.9 min. COSPAR: 1962-Eta-3. USAF Sat Cat: 259. Decay Date: 7 June 1963. <br />First generation photo surveillance; return of camera and film by capsule; SAMOS type satellite. Failed to return camera and film. Samos film return project cancelled; remaining 4 cameras placed in warehouse and later used on KH-6 Lanyard. References: 1 , 2 , 5 , 6 , 278 . <br /><br />http://www.stormingmedia.us/87/8788/A878833.html<br />Abstract: By now the explosions of the upper stages were supposed to give the most essential contribution to the space debris production. Thus, investigation of the breakup of the upper stages is of great importance for further development of space contamination models. It was only in 1981 when Don Kessler, NASA-JSC was able to correlate space debris from satellite breakups recorded by NORAD/ADCOM to upper stages of rocket carriers left on orbit after completion of their mission 1. Since 1969 up to 1981, ten cases of breakup of Delta second stages left in orbit after mission took place 2. The duration of stay of the vehicles in orbit before the explosion varied from 1 day up to 5 years. <br /><br />http://ast.faa.gov/files/pdf/q22002.pdf

 

[JonClarke]

I don't insist on a 1200 km orbit - it is what is specified in the ESAS summary. This carries a 0.5 km/s penalty over a 200 km orbit, regardless of propulsion mode. It would seem that the ESAS people prefer to keep their live reactors at 1200 km, whereas Borowski is happy to have them at than 400 km. I know which the public will want.<br /><br />Jon<br /><br /> <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[JonClarke]

Certainly the more massive the spacecraft the less they will be effected by air resistance. The ballistic cross section will also be low.<br /><br />Small spacecraft decay more quickly of course. <br /><br />Some figures from some datellites stranded in low orbits:<br /><br />Cosmos 21 with a perigee of 192 km decayed after 3 days.<br /><br />Cosmos 29 with a perigee of 197 km decayed in two days.<br /><br />Cosmos 96 perigee 222 km, decayed 22 days<br /><br />Cosmos 167 perigee 211 km, decayed 8 days<br /><br />Rule of thumb seems to be that for small satellites <200 km you have a few days, />200 km you have a couple of weeks.<br /><br />For orbital assembly 300-400 km is a safer bet (depensding on how many spaceraft you want to dock. if all you are after is a parking orbit, than 200 km is fine.<br /><br />Jon<br /> <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[gunsandrockets]

"the Borowski paper http://gltrs.grc.nasa.gov/reports/2002/TM-1998-208834-REV1.pdf"<br /><br />Just finished skimming through most of that report. Very interesting. Thanx for the link.

 

[gunsandrockets]

"Incidently, on the subject of save disposal orbits, Borowski et al. say that up to 2.5 km/s is need to for each NTR stage to put into a disposal orbit."<br /><br />After reading the Borowski report for myself (pg 20 Table 5 and pg 24) I see you neglected to mention that as little as 100 m/s is needed for safe disposal of the NTR stages.

 

[gunsandrockets]

"I don't insist on a 1200 km orbit - it is what is specified in the ESAS summary."<br /><br />No, it says a 800-1200 km orbit...<br /><br />http://www.spaceref.com/news/viewsr.html?pid=19067<br /><br /><br />"The CEV is first injected into a 55x296-km altitude orbit while the MTV loiters in a circular orbit of 800- to 1,200-km altitude. It then takes the CEV up to 2 days to perform orbit-raising maneuvers to close on the MTV, conducting a standard ISS-type rendezvous and docking approach to the MTV."<br /><br />You are exaggerating the delta V penalty. There is no reason to believe NASA must put it's NTR MTV as high as 1,200 km. Or that a CEV would EOR as low as 200 km. (The DRM link you provided did not mention any EOR) A more realistic penality is about half of the 0.5 km/s you estimate.

 

[gunsandrockets]

In the ESAS document the manned Mars-Transit-Vehicle (that carries crew from Earth to Mars and back again) is very similar to the spacecraft described on page 34 of the report (and illustrated on page 39) that Jon linked to...<br /><br /> http://gltrs.grc.nasa.gov/reports/2002/TM-1998-208834-REV1.pdf<br /><br />It's a 139.6 tonne spacecraft with 3 small 15,000 lb thrust bi-modal nuclear-thermal-rockets. The nuclear engines together also generate 50,000 watts of power.<br /><br />(page numbers corrected) <br /><br />

 

[JonClarke]

Glad you like the Borowski et al. report. He is one of the leading US researchers in the field. I am still mining that paper for information.<br /><br />The NTR design they propose (which incidentally is the result a joint US/Russian expertise) has the extremely impressive Isp of 950. All theoretical of course, it has never been built.<br /><br />Jon <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[JonClarke]

"After reading the Borowski report for myself (pg 20 Table 5 and pg 24) I see you neglected to mention that as little as 100 m/s is needed for safe disposal of the NTR stages."<br /><br />It is good planning pratice to take the more conservative figure. Especially when we are extrapolating out to safety requirements 25 years from now.<br /><br />Jon <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[JonClarke]

Several things here.<br /><br />Do you know the difference in m/s between a 1200 km circular orbit and a 800-1200 ellipitical one?<br /><br />NASA is the one that has specified the high orbit not me. Therefore there is every reason "to believe NASA must put it's NTR MTV as high as 1,200 km" (or, if you prefer, 800-1200 km.<br /><br />I never said that EOR would take place at 200 km. I said (or tried to) was that with a chemical TMI stage you can use this orbit this low for departure, achieving the maximum payload and propulsive efficiency. Remember with the ESAS architecture proposed you don't need EOR for 2 of the three launches. So your parking orbit could be even lower than 200 km, as it was with Apollo, if necessary. <br /><br />Remember regardless of propulsion mode you have a limited window to get to Mars from a orbital perspective. Plus if you are doing EOR you are restricted by the hydrogen boil off. For both say 30 days. The orbit has to be stable for this time. While 200 km is a bit low for this, 300 km should be fine.<br /><br />If you are going to do EOR for Moon or Mars missions, you want to do it as low as possible to get maximum advantage. Unfortunately for NTR low as possible means a 800-1200 km orbit.<br /><br />However, remember, we are talking 25 years down the track. I would be be surprised if, when the missions are actually in preparation, that serious consideration is not given to launching the crew with the Mars bound spacecraft, eliminating the need for EOR, at least for crew transfer.<br /><br />Jon <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[JonClarke]

I'll go to try and steer the discussion back on track (the ESAS Mars architecture) and away secondary issues of the issues that surround the use of NTR. <br /><br />I take as given:<br /><br />1) The use of a modified the Drake "Dual lander" architecture. This means:<br /><br />Three launches, two unmanned ~two years before the crew<br /><br />Use of inflatable Mars hab<br /><br />Use of a dual descent/ascent vehicle (DAV)<br /><br />No ISRU<br /><br />Two MORs between the Mars transfer vehicle (MTV) before and after landing.<br /><br />2) The use of a BNTR similar to that of Borowski, rather than the SEP of the original Drake mission. This means:<br /><br />Slow (240 day) cargo transfer)<br /><br />Fast (180 day) crew transfer<br /><br />All propulsive mission, except for earth and Mars entry.<br /><br />3) Use of the ESAS SDLV and CLV philosophy. This means:<br /><br />Payloads of 118 and 27 tonnes to a 1200 km orbit, respectively. Total mission mass (3 X SDLV and 1 X CLV) will be 381 tonnes.<br /><br />EOR to transfer crew from the CEV to the MTV<br /><br />This raises a number of interesting questions we might like to discuss:<br /><br />A - Why no ISRU? Almost every mission proposal to NASA since 1990 has used it. It means that most of the DAV will be taken up in both mass and volume by propellants with only enough supplies for the 30 day initial mission.<br /><br />B - Is the mission mass too small? Borowski is probably the tightest NASA-related study to date, with 396 tonnes in LEO, ESAS, using current performance . This is 4% smaller on on already very tight mass budget.<br /><br />C - What modifications can be made to the inherited ESAS elements? A modest 4 increase in SDLV payload would make the total mass equivalent to that of Borowski. The EOR could be eliminated by having the crew ride the MTV into orbit onboard the SDLV (surely this is will be an option after 10 years of SDLV operations). <br /><br />Jon <div class="Discussion_UserSignature"> <p><em>Whether we become a multi-planet species with unlimited horizons, or are forever confined to Earth will be decided in the twenty-first century amid the vast plains, rugged canyons and lofty mountains of Mars</em>  Arthur Clarke</p> </div>

 

[josh_simonson]

I saw a ISRU lunar 'module' that was part of their envisioned lunar base. If lunar ISRU is used or at least developed to the prototype stage, Mars ISRU will be a slam dunk. The mars mission probably would require fuel for the return journey though in the event of a no-go on landing for some reason.