Fly Centaur atop Delta IV-H, use as Earth Departure Stage after LEO Rendezvous

• #60 by simcosmos on 13 May, 2010 10:51
• I have done some past brainstorm and simulation musings related with using and/or adapting something like the current Delta IV Heavy Upper Stage to achieve some kind of earlier crewed beyond LEO exploration goals (as well related with Delta IV Heavy being used for Orion CLV duties).
  
  I ask the reader for patience when reading what will write next (thanks in advance!).
  
  
  I. Slightly Heavier Lift plus more or less existing stages (single launch BEO crewed applications)
  
  
  The mentioned above brainstorms were sometimes mostly focused on integration of such stage + Orion on a single launch vehicle (based on past DIRECT  J12X / J13X vehicle design iterations but you could consider something else with equivalent conceptual capability, the important is the Earth Departure stack related stuff), please see the following links for (outdated) imagery...
  
  http://www.flickr.com/photos/simcosmos/4014041934/
  http://www.flickr.com/photos/simcosmos/3179561421/
  
  ... and the following related NSF posts for extra context:
  http://forum.nasaspaceflight.com/index.php?topic=18139.msg477552#msg477552
  http://forum.nasaspaceflight.com/index.php?topic=12379.msg323124#msg323124
  
  
  I'm only sharing these links here because they contain information that might be useful for this thread (at least given my perception of recent posts).
  
  
  The advantages of using some kind of  slightly heavier lift (than current EELV Heavy but not going as far as super heavy lift) single launch approach are:
  
  a) Near delivery or even full delivery into LEO of the departure stack composed by such stage + payload (in this case, some kind of crewed spacecraft): this means that the mission stage would have almost all its main prop. load at the start of TLI... The mission capability for such stage would depend of payload mass and boil-off mitigation assumptions...
  
  b) Another advantage of a generic slightly heavier lift capability is that something like the Delta IV Heavy Upper Stage would then perhaps not require as much modifications, at least from the perspective of boil-off mitigation: this means that the current design could almost be used - again, from this perspective - under current operational assumptions for the main burn (TLI), which would be done soon after LEO insertion (if using a parking orbit before TLI;  the mission design could also consider other slightly more 'aggressive' TLI assumptions).
  
  
  II. The relevance of all the above for this thread?
  
  The moment in which additional assumptions start to be included for further changes on a chosen existing hardware / stage (such as the case of Delta IV Heavy US on recent posts), it is the moment where such changes need to be a little more clearly accounted on things like dry mass estimations, performance calculations, simulation work (and in the implementation impact on time line / cost, etc)
  
  This also means a number of extra assumptions that need to be analyzed. If the objective is to use a dual launch of some kind of currently existing or even slightly upgraded EELV to achieve some kind of crewed beyond LEO target goal then such changes need to be accounted in the implementation / development assumptions for the mission design trade space, else the analysis - even if just a very first order analysis - might lack in detail and not be suitable for comparison with other cases.
  
  This to say that, as most people here are aware (even if just by looking at pdf links that were provided above), the idea of using existing or slightly upgraded EELV for beyond LEO crewed exploration is not new.
  
  
  Focusing in DeltaIV related hardware, there are some famous images from the pre-Constellation times where Earth Departure / Transfer Stacks for lunar missions (just an example) would be made by joining a number of modified Delta IV Heavy Upper Stages: the 'problem' is that such stages dry masses would be higher than the current operational design if needing to assume some kind of rendezvous due to lower launch lift capability... and this beyond assuming other needed adaptations related with integration for crewed duties...
  
  ... Only as a very rough example, have somewhere in my download archives, a pdf that showed a first order estimation of ~7t or so (if remembering well) for such kind of modified Delta IV Heavy US (in order to account for extra mass related with boil-off mitigation for extended orbital / mission capability and for rendezvous + docking with other stages / crewed spacecraft). Of course that such data could be first order or even outdated or I might not remembering well (might try to (re)find the pdf, if someone wishes)...
  
  
  This then brings the question about the assumption of using 'something else' a bit more optimized for the mission designs in question... Such 'something else' could be one or several of the following examples... new upper stage design (instead of modifying existing stage) and/or heavier lift capability and/or depots, etc... or, also as alternative, might lead to the reduction of mission design goals / capabilities (example: assumption of much lighter crewed spacecraft launching first... something like a modified Soyuz or Shenzhou equivalent mass class spacecraft) and rendezvous within hours with upper stage being used for departure (if such stage is a cryogenic stage with very little modifications regarding currently operational designs).
  
  
  To conclude, the topic of this thread is very interesting indeed but it might become even more interesting if thinking about introducing a little more data on the starting assumptions in order to better study mission capabilities vs implementation constraints (technology assumptions vs development times vs cost vs impact on mass / propulsion assumptions vs impact on mission design ground rules).
  
  António
• #61 by neilh on 14 May, 2010 17:56
• I don't think anybody's mentioned it yet, but the feasibility study for sending a manned Orion to a NEO uses as its lower bookend a mission consisting of a Centaur launched on a Delta IV-H which docks with an Orion:
  
  http://ti.arc.nasa.gov/projects/neo_study/pdf/NEO_feasibility.pdf
  
  Of course, if Orion isn't being man-rated for launch, you'll probably want to launch a separate crew taxi or get the crew from the ISS.
• #62 by FinalFrontier on 14 May, 2010 19:21
• Quote from: neilh on 14 May, 2010 17:56

  I don't think anybody's mentioned it yet, but the feasibility study for sending a manned Orion to a NEO uses as its lower bookend a mission consisting of a Centaur launched on a Delta IV-H which docks with an Orion:
  
  http://ti.arc.nasa.gov/projects/neo_study/pdf/NEO_feasibility.pdf
  
  Of course, if Orion isn't being man-rated for launch, you'll probably want to launch a separate crew taxi or get the crew from the ISS.
  

  This was interesting.
• #63 by simcosmos on 15 May, 2010 08:29
• Quote from: neilh on 14 May, 2010 17:56

  I don't think anybody's mentioned it yet, but the feasibility study for sending a manned Orion to a NEO uses as its lower bookend a mission consisting of a Centaur launched on a Delta IV-H which docks with an Orion:
  
  http://ti.arc.nasa.gov/projects/neo_study/pdf/NEO_feasibility.pdf
  
  Of course, if Orion isn't being man-rated for launch, you'll probably want to launch a separate crew taxi or get the crew from the ISS.
  

  
  I was (still am) trying to prepare a post where would clearly share starting assumptions about Earth Departure Stage (or stack of stages) for such kind of 'lower bookend' crewed missions, I mean, missions which would assume a minimum of two launches (up to 4 EELV 'class'): at least one vehicle to launch an adaptation of an existing EELV stage (either AtlasV Centaur or Delta IV Heavy US) plus at least another launch to deliver some kind of crewed spacecraft which would rendezvous with such departure stage(s)...
  
  
  However, for the moment, I would ask the readers to please look with a little more attention to the assumptions available on documents such as the one that was shared above versus look at the date they were produced vs think a little more about the starting assumptions...
  
  This all to say that in the special case of these 'lower bookend' missions it is enough to just be a little more conservative in some assumptions to quickly make a mission concept unfeasible (which would lead to the consideration of extra assumptions and their impact on time line + development + cost).
  
  
  To give an example of a constrained mission design (really, really, lower bookend), for the case of using something like an adapted AtlasV Centaur US + some kind of Orion while, at the same time, assuming slightly more conservative numbers for those components: I'm not being able to achieve much better than ~3950m/s up to 4150m/s dV budgets (preliminary math for total US + Orion dV). The Orion spacecraft assumed on some older documents also seems to have a dV of 1700m/s (older iteration) but have seen some official public NASA documents where the maximum dV available for more recent Orion iterations was closer to ~1500m/s or so... Not sure if the work about 'lower bookend' missions was updated to reflect details like these or not...
  
  
  ... In any case, that is why wrote my previous post: unless the starting assumptions for the mission hardware and mission design constraints are more clearly shared, it becomes a bit difficult to really brainstorm - and compare - the impact of such starting assumptions on conceptual mission profiles...
  
  ...Without such 'work' done - or without other people's work somehow 'verified' and/or updated against last known information - it is very easy to write, in these forums, about assuming a 'Delta IV Heavy Upper Stage waiting for a crewed spacecraft outside of a given ISS envelope' or writing about 'using Centaur + Orion' for these 'lower bookend' BEO missions... (slightly more challenging and time expensive is to present some math / extra assumptions to backup the thoughts and for others to review / discuss a bit further).
  
  So, on my humble opinion, it would be a lot more interesting to fully share / verify / adapt all starting assumptions so that everyone could have a common starting ground - even if just first order - to clearly follow calculations and mission design constraints (at least believe that was the spirit of this thread, please let me know if not).
  
  
  If possible, will then try to share (although not sure when will be able to do it) the range of starting assumptions that, at least (and for what is worth!) I'm using here for a modified AtlasV US,  DeltaIV Heavy US and for crewed spacecraft options versus eventual 'low bookend' conceptual mission assumptions (while trying to keep things constrained to a maximum of four EELV, minimum of two). 
  
  Thanks,
  António   
• #64 by simcosmos on 16 May, 2010 12:00
• 'Lower Bookend' Beyond Earth Orbit Crewed Missions
  
  Definition: mission designs using existing, near-existing and/or currently ongoing testing / in-development hardware to achieve,with Orion (or other  crewed spacecraft), a given beyond-LEO mission goal by using a minimum of two launches (EELV or vehicles with similar payloads range) at the earliest and most cost-effective possible opportunity.
  
  
  As noted on past posts, will try to share a few starting assumptions related with this thread’s topic so that:
  - other forum participants are able to review the data
  - some mission design and implementation constraints are shared
  - comparisons with other past or future mission assumptions, documents, etc becomes slightly easier (at least a common ground / numerical starting point is provided for a specific scenario)
  
  
  For this particular post will only focus on the following case: one launcher delivering an adapted Centaur SE which would act as Departure Stage for the rest of the mission components (crewed spacecraft and perhaps mission specific hardware would be delivered by the second launcher).
  
  _{Note: on later occasion might focus the attention on Delta IV Heavy Upper Stage being used as EDS but, for the moment, starting then by a lower bookend case…}
  
  
  Whenever possible will try to use existing or near-existing hardware and reference to source data. More aggressive assumptions are certainly possible (subjected to the impact on time line + cost, etc…). As a last introductory note, of course that readers are free to review this text but, if doing so, please try to do it with some ‘substance’.
  
  
  
  I. Atlas V Centaur US (single engine, adapted for beyond LEO missions)
  
    3500 kg : inert mass
  20830 kg : prop. mass (at lift-off)
  99200  N : thrust
    450.5s    : ISP
  
  Notes:
  - Accordingly with ‘AtlasV Users Guide 2010 (rev11, March)’ inert mass for 5X1 Centaur SE is 2247 kg and, for AtlasV Heavy, such stage mass is given as 2316 kg; on the ‘Centaur Application to Robotic and Crewed Lunar Lander Evolution’ paper, the Centaur SE is 2500 kg: I’m using this latest value together with the addition of 800 kg for an ‘Extended Duration Mission Kit’ (also from public LM/ULA materials) in order to obtain what believe to be a better representation of a Centaur SE stage adapted for operation within 7 days of being launched (equal to say, a better representation for the kind of mission profiles being referenced in this thread…)
  
  - An extra mass of ~200 kg is assumed on the stage’s ‘inert mass’ to account for a passive LIDS (similar to the hardware integrated on HST on STS-125): I’m not sure about the mass of such passive version (feel free to point for a source regarding that value)
  
  - Will also share results calculated by using 450.5s ISP / 0.25% prop. left on departure stage versus 445s ISP / 1% prop. left on the stage, to compare with a more conservative dV budget calculation (and to calculate an average dV budget).
  
  - Boil-off will be assumed at 0.2% / day of the initial LEO delivered prop. load (this assuming that the stage is fully delivered into LEO by the launch vehicle)
  
  
  Sources:
  http://www.ulalaunch.com/site/pages/Education_PublishedPapers.shtml
  http://www.ulalaunch.com/site/pages/Products_AtlasV.shtml
  http://www.ulalaunch.com/site/docs/product_cards/guides/AtlasVUsersGuide2010.pdf
  
  
   
  
  II. Crewed Spacecraft & Mission Hardware
  
  II.a) Orion
  
  Will assume that some kind of BEO Orion implementation will happen in the future. Because I’m not sure about what is the most current state of Orion mass breakout will use past information / extrapolate eventual Orion mass breakouts for the intended mission profile.
  
  _{Note: another option would be to look at this as a ‘black box’ (placeholder number) for a given total mass which could represent a generic crewed spacecraft alone (not necessarily Orion) or a crewed spacecraft + mission related module & payloads to, at least, calculate the dV provided by the modified Centaur SE.}
  
  
  Orion CM (lunar / BEO variant) has been noted, on earlier iterations, at ~8t… up to ~8.7t up to perhaps 9.3t, at lift-off, for a crew of four astronauts living inside it for periods up to ~18 days (?) (when accounting for Constellation’s Lunar Mission profile and for maximum loiter times on LEO, LOI, TEI phases + when taking in consideration the length of the journey to/from Moon).
  
  Orion SM: have seen estimations ranging from ~3.6t on very early iterations up to ~4.3t up to ~>4.5t (?)… Main prop. load of at least 8t… Will assume main engine at 326s ISP and from 0.25% up to 1% prop. left on the service Module (for FPR / Residuals, etc.)
  
  
  As mentioned above, not very easy to have a clear picture about the current Orion state… Some of its mass growth and configuration changes are/were very closely related with the specific AresI integration, other part of the spacecraft mass growth is ‘natural’ of any development process…
   
  What will present next are then some ‘options’ for several Orion mass breakouts / rough first order performance calculations (kept SM mass the same and varied SM prop for two CM cases)… This might hopefully be representative and conservative enough for a first order estimation (while assuming full LEO delivery of the vehicle)...
  
  22.3t : 9.3t (CM) + 4.5t (SM) + 8.5t (SMprop) ~1515 m/s
  21.8t : 9.3t (CM) + 4.5t (SM) + 8.0t (SMprop) ~1443 m/s
  20.2t : 9.3t (CM) + 4.5t (SM) + 6.4t (SMprop) ~1203 m/s
  
  21.7t : 8.7t (CM) + 4.5t (SM) + 8.5t (SMprop) ~1568 m/s
  21.2t : 8.7t (CM) + 4.5t (SM) + 8.0t (SMprop) ~1495 m/s
  20.2t : 8.7t (CM) + 4.5t (SM) + 7.0t (SMprop) ~1343 m/s
  
  
  Note:
  - If the total mission scenario would take ~>18 days then the crew would probably have to be reduced to three or even to two astronauts, depending of extra mission requirements (such as desired contingency margins, mission hardware, etc)…
  
  
  II.b) Dragon
  
  Not sure about what would be the mass breakout for a lunar Dragon variant: perhaps some of the eventual mass growth could be achieved with crew number considerations and by carrying less cargo on the trunk. For the moment will share a first order estimation of a Dragon mass breakout for LEO, only to provide a starting point for eventual later brainstorms (feel free to correct data):
  
  3250 kg : capsule mass
  2500 kg : 7 (?) max. crew or less crew (3?, 4?) + some combination of internal / external cargo,  etc
  1290 kg : OMS / RCS prop (270s ISP?)
    681 kg : Trunk
  ------------
  7721 kg : Dragon mass (crewed flights)
  
  Falcon9 assumes a payload of ~10450 kg into 200km, 28.8 inc. There seems to exist enough performance margin to either include a standard LAS design (released during ascent) or some kind of dual-duty abort system on the capsule itself (simulations needed!).
  
  Assuming 270s ISP and ~1% left on the integrated SM tanks, the above numbers would result on a dV budget of ~479 m/s.
  
  Only for reference, assuming a cargo only variant (2500 kg internal cargo + equipment, etc and ~2729 kg of external cargo carried on the trunk, for a total cargo of ~5.2t and a total spacecraft mass of ~10450 kg, the dV budget (again keeping 1% prop) could be ~345 m/s.
  
  This all to write that in theory, a lunar Dragon with a crew of ~2 to ~3? astronauts could perhaps have enough resources to return from EML2 (using propulsive swing-by at the Moon): extra study needed about the mass / performance impact of the changes that would be required to transform the conceptual LEO Dragon into an even more conceptual Lunar Dragon.
  
  
  
  II.c) Mission Hardware
  
  Crew size could also be strongly dependent on the amount of extra mission hardware that would need to be carried in this kind of ‘lower bookend’ missions…
  
  Back to Orion mass breakout examples,  the reader can, for example, look at the heavier crew command module and think about reduced crew size (2 astronauts?) and some of that 9.3t mass as being ‘mission hardware’ (either transported inside the CM – such as EVA suits and other tools - or attached to the SM exterior…).
  
  Another option would be to assume an extra mission module… However, such extra module (plus its internal contents / external payloads) could well reach a mass from ~3t up to ~>4t (depending of mission needs)…
  
  … Always depending of extra assumptions – and focusing for the moment only on Orion - I’m not sure for which mission design such extra module would be feasibly integrated IF assuming only two launches of existing EELV / EELV Heavy (or equivalent payload vehicles) and without assuming something else beyond the AtlasV Centaur SE modifications mentioned above… Yes, extra lift capability under the form of EELV upgrades could also be assumed but then all that starts adding to the timeline (I’m already assuming several adaptations to Centaur SE upper stage, its eventual integration on a launcher that might not be an AtlasV,  development of Orion for BEO, etc)…
  
   
  
  
  III. Mission Design Constraints (Launch)
  
  The mission would start with the launch of the crewed spacecraft or even the departure stage… A good number of extra simulation and related considerations would be needed to decide which payload would be launched first and to also better study injection targets for each payload vs a better definition of rendezvous procedures vs Earth Departure setup requirements…
  
  
  If adding a mission module of ~3t up to ~>4t, such module could perhaps be launched (depending of mass and launch vehicle choice) together with a non crewed Orion vehicle: in such case, the delivery requirement could be something like ~>25.25t total spacecraft mass (not counting with adapters) into ~250 km altitude…
  
  Will however focus on the ‘no-mission module case’ for this specific scenario where the departure stage would be a modified AtlasV SE US: as roughly represented above, the Orion could then have a lift-off mass of ~22t for an average dV of ~1500m/s or so...
  
  The AtlasV Centaur SE based departure stage could have a total mass of ~24.33t…
  
  
  Both these payload targets seem to be feasible (need to double-check / simulation work needed) for Delta IV Heavy (which in this time frame would use RS-68A): it would perhaps require dual launch within up to ~5 days (extra study needed, have somewhere a pdf about AresI+V regarding this issue of launch order vs lunar mission). If using dual Delta IV Heavy, one of them would need Human Rated Kit (as well related facilities adaptations).
  
  Another Heavy vehicle option (perhaps AtlasV Heavy… or Falcon9 Heavy!, or AresI!... although AresI injection would be sub-orbital and require a few extra considerations, etc) or some other EELV upgraded iteration could also be assumed  (which, depending of specific assumption, would also add to development and/or timeline and/or cost)…
  
  Please remember that I’m still only talking here about an entry level for lower bookend type of missions. This all needs to be taken in consideration…
  
  
  Yet another ‘nearer-time’ alternative could be to assume Delta IV Heavy launching the AtlasV upper stage (which poses its own integration challenges), something like an AtlasV 552 launching ~21t Orion with no crew (the dual Centaur would have to be fielded…) and ‘something else’ (not many options left, really...) launching the crew… although this would become a 3 launch scenario with things starting to complicate… near-ISS space could also be used as a stationary / assembly point for the crew to wait for the Earth Departure Stack docking but that would introduce yet another layer of complications for this type of ‘lower bookend’ missions... better to avoid that for this specific case.
  
  
  Given all the constraints, not sure when even this kind of lower bookend mission would be possible… maybe I’m being pessimist…
  
  
  Summing up, would perhaps then baseline two launches of the heaviest EELV (or similar class) possible vehicles available at the date: one perhaps launching EDS alone, the other – human rated - launching the crewed spacecraft. Considerations about which launch would be first or about eventual mission module would need a much better definition of mission requirements, mission hardware and specific simulation work.
  
  
  
  
  III. Mission Design Constraints (dV budgets)
  
  Regarding the dV budget, under what have assumed:
  
    3500 kg : AtlasV Centaur SE US + Kit + Passive LIDS
  20538 kg : Centaur main prop. load (after 7 days @ ~0.2% boil-off / day)
  20900 kg : Orion: 9.3t (CM+mission hardware?) + 4.5t (SM) + 7.1t (SMprop, after rendezvous, initial load would be ~>8t, simulation needed) 
  --------------
  44938 kg : total mass at start of departure burn
  
  Depending of how math is made:
  
  dV1 (Centaur): ~2688 m/s (450.5s ISP / ~0.25% prop. left)
  dV1 (Centaur): ~2661 m/s (450.5s ISP / ~1.00% prop. left)
  dV1 (Centaur): ~2656 m/s (445s ISP / ~0.25% prop. left)
  dV1 (Centaur): ~2629 m/s (445 ISP / ~1.00% prop. left)
  
  dv2 (Orion): ~1311 m/s to 1322 m/s (depending of FPR, etc)
  
  This all gives about ~3975 m/s total dV budget (in average) for the above Centaur SE + Orion assumptions (two launches).
  
  
  III.a) Performance Comments
  
  - Such kind of average dV estimation and ~18 day journey duration could be compatible with a 3 astronauts crew delivery into EML2 and return from there:
  
  ~1 x 3148.5 m/s : TLI (shared by Centaur SE + Orion)
  +2 x   184.1 m/s : propulsive swing-by at the Moon (In/Out)
  +2 x   147.5 m/s : EML2 (In/Out)
  (with ~163 m/s available for MCC, station keeping)…
  
  
  _{Source: http://forum.nasaspaceflight.com/index.php?topic=1337.msg18213#msg18213
  (RE: An Alternative Lunar Architecture)}
  
  It would be better if the crew could have ‘something’ waiting for them at EML2, else they would arrive there only to return on the next hours (and if that would be the mission objective perhaps a simpler lunar fly-by profile would be better)
  
  - Such kind of ~3975 m/s dV and more than 18 day mission having EML2 (or EML1, also using swing-by, see pdf below) could perhaps also be achieved by a reduced crew of just two astronauts while, at the same time, perhaps also carrying a little of extra mission specific payload (IF the goal of the mission would be to service something at EML2, for example): not sure however how much payload would be feasible (depending of assumptions for the crewed spacecraft, etc).
  
  
  On my humble opinion and under my clumsy first order considerations, little else could perhaps be assumed (do not see great opportunities for reaching a BEO object) unless going for more capable launch vehicles and/or more capable propulsion stage assumptions and/or optimized spacecraft assumptions and/or depots and/or using specific date geometries together with extra trajectory tweaks for intended mission objectives…
  
  … Each or several groups of those extra assumptions would need to be taken in account in order to better compare implementation constraints, but all that would add to the time line of the ‘simplest’ case that have tried to very preliminarily study and share here (extra study, assumptions refinement, simulation work and more specific design assumptions would be needed).
  
  So, summing up, have tried to share some considerations about:
  
  -   AtlasV Centaur SE (with kit + passive LIDS) fully delivered into LEO
  -   Orion (also fully delivered into LEO) launched within ~5 days or so and rendezvous, within ~2 days, with departure stage
  -   departure dV provided by upper stage burn + Orion undocking and making the final commitment for such burn
  -   Orion making rest of middle course correction / main burns
  -   mission duration of ~18 days for crew of 3?
  -   mission duration of ~>18 days for crew of 2? (+ X? payload?) 
  
  
  
  III.b) Related readings and final comments:
  
  http://servicingstudy.gsfc.nasa.gov/presentations_final/day2/Harley_Thronson/Thronson_servicing_workshop.pdf
  
  (http://servicingstudy.gsfc.nasa.gov/workshop_1_presentations.htm)
  
  
  DOES THE NASA CONSTELLATION ARCHITECTURE OFFER OPPORTUNITIES TO ACHIEVE SPACE SCIENCE GOALS IN SPACE?
  http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080032732_2008031420.pdf
  (IAC-08-A5.3.6)
  
  A few differences on the IAC-08-A5.3.6 pdf (versus what I have written above) are:
  
  - No boil-off assumed? (absence of long duration kit on Centaur?)
  
  - Quick (next orbit?) back-to-back launch of dual AresI from both KSC pads 39 (instead of assuming other launch vehicles) + rendezvous within hours (instead of allowing for up to ~5 days difference between launches, from the moment the Centaur would be launched?)
  
  Comment: If using aggressive mission components delivery, the crewed spacecraft (perhaps launched first into a lower orbit) could rendezvous within hours with the departure stage but that does not leave much margin for unexpected events…
  
  … On another hand if using slightly less aggressive rendezvous assumptions, the departure stage would need addition of extra boil-off reduction kit (probably similar to what I have assumed) and rendezvous time line of mission components could be a little more relaxed although the kit  assumption would have impact on how early such boil-off reduction kit would be available for operational use… as far as I’m aware, an early sun shield demo is expected for ~2011 (that would be only one component of the kit).
  
  
  - Inert mass for Centaur SE of ~1.8t (I have used ~2.5t for a modified Centaur SE + 800 kg Kit)
  
  - Lower Orion mass breakouts (CM + SM) than what I have assumed: it would be really interesting to have an update regarding Orion status, despite all the indefinition we are experimenting nowadays…
  
  - ~>2.7t airlock module (the differences on Centaur assumptions + CM + SM almost cover the minimum module mass range): on the IAC-08-A5.3.6 paper the AresI also seems to be defined as a ~>26t payload vehicle
  
  - Different dV calculation methods and trajectory assumptions: haven’t fully verified those. On my calculations tried to keep some FPR / Residuals on departure stage, SM, etc, not sure if that was the case on the pdf. This post is already long enough (!!!): feel free to define ground rules, compare and refine estimations and methods.
  
  
  Real Final Comments:
  
  Quick note about assuming the crewed spacecraft with a total mass of ~15t instead: this could allow the Centaur to make full TLI. Such crewed spacecraft could look like an Orion CM with a smaller SM or like a Shenzhou with larger SM or with an extra mini-stage (or, more relevant, Lunar Dragon with some kind of SM connected to the trunk: mission capability would depend of assumptions).
  
  
  To end, could also write a similar preliminary analysis / references about a possible brainstorm 'next level' (using / adapting DeltaIV Heavy Upper Stage, based on some pdf / past notes / simulation work / math) but that text could probably reach a point where it would be 'simply' better to start assuming or a new stage development or heavier lift capability than the one expected for Delta IV Heavy (or similar EELV-class) or both those changes and/or even several other extra assumptions...
  
  So, will save forum space and, at least for now, post just this ‘lower bookend’ case where the Departure Stage would be an adapted AtlasV Centaur SE... At the same time provided a few related links / references and some numbers about what could be the mass breakout of such adapted stage, Orion, Dragon, etc which might provide a good starting point if someone else wishes to further refine calculations and mission design / implementation constraints for this kind of lower bookend cases.
  
  Thanks,
  António
  
  
  _{PS: this was a loooong post: my apologies in advance for any less clear text / math / errors and/or typos (wrote this in pieces and then did a bit of intensive final copy+paste)}
• #65 by sdsds on 18 May, 2010 01:14
• Another reason to select Centaur rather than DCUS is because over the program's history Centaur vehicles have been exposed to (and survived) a wide variety of launch environments.  DCUS has flown about a dozen times total, on Delta III and Delta IV vehicles only.
  
  And speaking of launch environments, what are the chances a Centaur (or DCUS) could be flown inverted, i.e. nose-to-nose with the second stage of the launch vehicle?  This might substantially reduce the mass of the interstage hardware.
• #66 by simcosmos on 18 May, 2010 10:35
• Quote from: sdsds on 18 May, 2010 01:14

  Another reason to select Centaur rather than DCUS is because over the program's history Centaur vehicles have been exposed to (and survived) a wide variety of launch environments.  DCUS has flown about a dozen times total, on Delta III and Delta IV vehicles only.
  
  And speaking of launch environments, what are the chances a Centaur (or DCUS) could be flown inverted, i.e. nose-to-nose with the second stage of the launch vehicle?  This might substantially reduce the mass of the interstage hardware.
  

  
  AtlasV Centaur vs DeltaIV Heavy US vs longer space endurance kits
  
  On my opinion, something like the assumption of modifying an existing Centaur (vs DeltaIV Heavy US) could make sense (for the specific case of lower bookend missions while using existing or near-existing launchers) given the following reasons:
  
  - on-going work about the topic (longer endurance kits) seems to be already happening and using Centaur as a test platform (at least in a more active way than on Delta IV Heavy US case)
  
  - the total mass of such modified Centaur could apparently still be within the maximum payload range for a good and stable enough LEO insertion by an Heavy EELV (or equivalent ‘class’ vehicle)
  
  
  If wishing to take the brainstorm to a ‘next level’ and extrapolate about modifications done to the Delta IV Heavy Upper Stage… Well, the current stage is ~3490 kg for a propellant load of ~27.2t, thrust of ~110 kN, vac. ISP of ~462s (459s if wishing to be conservative on calculations)… 
  
  … The above means that the total mass of the stage, *without modifications* is ~3.49t + 27.2t = 30.69t… This could be OK if assuming a heavier launch capability to single-launch the departure stack (crewed spacecraft + eventual mission related hardware + departure stage) and with such stage making departure burn soon after reaching orbit (in a similar way to what have mentioned above on DIRECT related brainstorms): such single-launch heavier lift capability could at least workaround the issue of not using the stage much beyond its current endurance limits (and/or minimize the modifications needed)…
  
  
  However, if assuming instead something like current EELV / Delta IV Heavy (or even something like AtlasV Heavy, Falcon Heavy, etc)… Well, a modified Delta IV Heavy Upper Stage *with extra kits for longer space endurance* and also with extra mass related with mission hardware integrations (docking / integration with mission payload or extra stage, etc) could eventually weight more than 31t, depending of estimations for the modifications vs conceptual mission profiles… The mass of the stage could jump from those ~3.49t up to something closer to ~5t or even greater (have seen an older pdf, where a probably conservative brainstorm (?) assumed ~7t, including stuff such as densified prop. for a load of ~28.2t, boil-off reduction kit, docking hardware capability with other Delta IV Heavy US… if remembering well, the two-stage stack would allow for a TLI payload of ~35t or so).
  
  
  Summing up, if not assuming heavier launch vehicles (between ~50t up to ~80t or so, being that such heavier vehicles could also make good use of longer endurance stages althought not requiring it for a first phase) and/or several other things (such as depots, etc), we could end-up by talking about a longer endurance modified Delta IV Heavy Upper Stage resulting in a total stage mass ranging from ~32t up to ~36t… Depending of extra assumptions, this could require a launch vehicle capable of at least ~35t up to 40t or so delivery into at least 240 km up to 300 km 29 inc Earth orbit… this, if the objective would still be to deliver a fully loaded departure stage – in this case based on Delta IV Heavy design - to rendezvous with mission payload (and/or with other stage to form a departure stack)… 
  
  
  So, from this perspective, not sure if considerations about flying the departure stage upside down in order to eventually save on interstage mass would be relevant because:
  
  a) longer endurance kit additions to current Centaur + interstage, etc seems to still be in the range of something like Delta IV Heavy (or AtlasV Heavy or Falcon9 Heavy, if any of those would be fielded one day…); extra refinement / study / simulation work needed... 
  
  
  b) if talking instead about longer endurance kits for Delta IV Heavy US we could probably end-up by talking on requirements for larger delivery capabilities (= extra modifications) of the launch vehicles I mentioned above… If such would be the case, the vehicles would need to be modified (or any of their evolution phases implemented and/or something like at least an ‘entry-level’ SDLV or clean sheet design, etc) in order to meet lift requirements… Or else… some other changes regarding technology and/or mission design, etc would have to be assumed… (articulated with better assumptions for BEO crewed spacecraft configuration...)
  
  … That is why wrote somewhere on the previous posts that this could quickly lead to considerations about new upper stage designs, stages which would be built and optimised - from day one - with larger endurance applications in mind… (versus talking about adapting current upper stages).
  
  
  
  Extra Comment about AtlasV Centaur vs Lower Bookend Missions
  
  Having written the above, that is why tried to focus the previous looooong post only on lower bookend missions using Centaur as a tesbed for slightly longer endurance applications (than what happens nowadays): anything else would probably have impacts on the implementation timeline as well need the assumption of extra development / monetary / technology stuff… In which case it would probably be better to leave behind the ‘lower bookend’ mission concepts / assumptions and go instead for something slightly more optimised and adapted to mission needs (in terms of hardware design assumptions) and less constrained in terms of mission capabilities…   
  
  
  Would also like to remember the readers that on my previous post, haven’t really taken in account things such as gravity losses: with a T/W of ~0.23 for something like the modified Centaur SE + Orion described on such post, that could mean gravity losses in the order of ~75 m/s or so (this looking at a graphic for a 400 km departure, from ESAS report)… This could be in part mitigated by using a much lower departure altitude and that is why also mentioned on such post that extra assumptions refinements (mass breakouts, for example) and a good amount of extra simulation work would be needed about LEO injection targets for the modified Centaur and for the crewed spacecraft vs order of launch vs rendezvous procedures  vs departure burn setup, etc, etc
  
  António
• #67 by Jim on 18 May, 2010 12:37
• Quote from: sdsds on 18 May, 2010 01:14

  
  And speaking of launch environments, what are the chances a Centaur (or DCUS) could be flown inverted, i.e. nose-to-nose with the second stage of the launch vehicle?  This might substantially reduce the mass of the interstage hardware.
  

  
  They are designed to take loads from a certain direction
• #68 by Robotbeat on 03 Dec, 2010 00:45
• This topic is being brought up in another thread, so I'm bumping this thread.
  
  Another thought: Dragon on Delta IV Heavy could go straight to EML1/2. I think Dragon would have plenty of delta-v capability (~500-700m/s) to handle EML1/2 insertion and departure (total of 350 m/s?).
• #69 by simcosmos on 19 Feb, 2015 13:48
• Hello all,
  
  I'm trying to (slooowly) return to some addon development work for Orbiter Space Simulator and, when looking at 'simcosmos' archives, I have decided to unfreeze / revamp an old 2006 Centaur 3D model in order to transform it into an Earth Departure Stage for an Orion spacecraft.
  
  This is part of an 'what if' conceptual side-exercise about Constellation Program and is also a kind of a follow-up to the brainstorms I have written in this thread (please see previous posts).
  
  It may also be more or less relevant for current days (Orion MPCV), because I'm assuming an heavier CxP Orion (~24.5t CSM at liftoff). This kind of Lower Bookend Mission does not give space for wild mission planning (at least not with an heavy assumption for the spacecraft mass)... As noted, with these heavy constraints, it will be all about a lunar flyby (unless doing extra tweaks on several assumptions)... As also referenced, a crewed spacecraft with a total mass between 6t up to 8t or so would obviously be much better (Orion's CM alone is already above those values).
  
  To conclude, my apologies for bumping the old topic but, independently of something like this (Centaur modification for EDS duty on a crewed mission) ever having a chance or not of becoming reality (in particular with a lighter crewed spacecraft), at very least I think that this is a right place to complete my previous participation on this thread with (the attached below) 'work-in-progress' contribute to the 'eye-candy' department.
  
  Cheers,
  António Maia
  
  PS: Anyone knows if there is still any intention to demonstrate a Centaur Sun shield in an actual flight? Thanks!
  

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