View Poll Results: Which Orion is better? A (smallSM) "CorkScrew Orion" or a (bigSM) "SwissKnife Orion&q

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Thread: Which Orion is better? A (smallSM) "CorkScrew-Orion" or a (bigSM) "SwissKnife-Orion"?

  1. #31
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    Quote Originally Posted by gaetanomarano View Post
    the image is from same days of ESAS report
    about the data in your post, 23 mT less 3 mT is 20 mT, just half ton more than my evaluation
    don't forget that (both) my ESAS table and your ESAS data refer to the old 5.5 m. CEV and small 4 mT LAS launched with the 27 mT SSME/CLV
    now the max payload of the J-2x Ares-I is 21-22 mT and the (real) LAS has +2 mT more than planned
    if the CLV lift less tons, the Orion has a known weight and the LAS has two extra tons... the only object we can/must reduce to close the equation is the SM

    Ö

    probably the author of the article has (simply) allocated the mass reduction in proportion with the original mass of the single parts
    his evaluation is clearly wrong since we know (now) that the LAS is not reduced but incresed in weight of 2.1 mT, the CM has only 1 mT of reduction, etc.
    then (again) the only weight we can/must resize to match the final (reduced) Orioin/SM mass (still) is the SM
    unfortunately, LM don't give (so far) any data about SM weight, so, we can only wait to know if the Orion/SM will weigh 19.5 or 21 or 21.5 mT
    I am through discussing this with you. You are simply being obtuse and there is no point rehashing the same thing over and over with someone entrenched in a demonstrability wrong position.

    And by the way, based on the current poll results, you havenít convinced a single person that youíre right. Perhaps that should give you some indication regarding the validity of your arguments.

  2. #32
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    Quote Originally Posted by Bob B. View Post
    ...no magical proportion...
    its' not a "magical proportion" but only the LEM specs NASA engineers have calculated from lunar gravity, engines, etc. (that can't change so much with the LSAM)
    ...it doesnít work that way...
    again... it can be a little more or less than 24 mT but not 30, 40, 50% less
    the AS/DS propellant a lunar lander need STILL is (about) 2/3 of the full weight
    then the mass of LOI propellant is 20.38 mT...
    first version not 100% correct, then edited:
    the Apollo+SM dry mass at TEI was about 12 mT, then, if (from your calculations) the 11 mT Orion dry mass needs 9.5 mT of propellant for TEI, the Apollo value was about 10.4 mT
    then, since the SM propellant (total) mass was 18.4 mT, the Apollo SM used (about) 8 mT of propellant to brake the 45 mT Apollo/LEM convoy to LOI
    the weight of the Orion/SM/LSAM convoy at LOI is only 45% more than Apollo/LEM, then, it can't need over 20 mT of the (more efficient!) LOX/LH2 propellant!
    to brake (only) 45% of (Orion/LSAM) extra mass, the right value is 8 mT (Apollo SM propellant for LOI) x 1.45 = 11.6 mT (that is very close to the amount of propellant of my evaluation!)
    and, since the Apollo LOI was 100% successful with these data... probably your calculations are wrong...
    ...a propellant's burning system...
    a "fire system" to start the LOX/LH2 burning (while hypergolic burns at contact of fuel and oxidizer)
    ...We now know the LOI propellant must be 29.53% of the convoy mass, or 20.38 mT based on a 69 mT convoy...
    no... YOU "now know", not "we"... (and the lunar-convoy is around 66 mT, not 69 mT)
    ...the 44 mT number was for a fully fueled Orion performing a mission in convoy with a LSAM...
    well... 66 mT of total Orion/SM/LSAM mass - 44 mT for the Orion/SM (from your calculation) = 22 mT LSAM - 10.8 mT of the ascent stage = 11.2 mT for the descent stage weight - 2.5 mT of LSAM cargo (rovers, etc. that are in the descent stage) = 8.7 mT - the weight of the tanks, DS engines, landing pads, LSAM structure, etc. (maybe 5 mT) = 3.7 mT !!!!
    if you're right about the LOI propellant, the consequence is that the LSAM must land on the moon with 3.7 mT of propellant!!!
    it's not a landing!
    it's a miracle!
    ...donít see what benefits are coming from all these autonomous Orion flights you keep promoting...
    read again my article and posts
    ...we want to land on the Moon...
    read again my post
    it's the best solution for the "political risks" you claim
    Last edited by gaetanomarano; 2006-Sep-21 at 01:17 AM. Reason: grammar

  3. #33
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    Quote Originally Posted by Bob B. View Post
    I am through discussing this with you. You are simply being obtuse and there is no point rehashing the same thing over and over with someone entrenched in a demonstrability wrong position.

    And by the way, based on the current poll results, you haven’t convinced a single person that you’re right. Perhaps that should give you some indication regarding the validity of your arguments.
    Thanks for all your effort Bob B. Very good info. Nice website too, what with all the real numbers and science and stuff. So many other space websites are so phantasmal.

  4. #34
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    Quote Originally Posted by Boxes View Post
    Thanks for all your effort Bob B. Very good info. Nice website too, what with all the real numbers and science and stuff. So many other space websites are so phantasmal.
    Thanks, Boxes. Even though a dialog with gaetanomarano sometimes feels more like I'm banging my head against a wall, it does make me think about some new things and I learn from it, even if gaetanomarano doesn't. And hopefully other people learn something as well.

    Let me add one more comment for the record. When I wrote that gaetanomarano is "entrenched in a demonstrability wrong position", I didn't necessarily mean his opinion about the big SM versus the small SM. I tend not to agree with him, but he certainly has a right to his opinion and it is an issue without a definitive right or wrong answer. I was referring to the numbers he keeps throwing around. Things like the amount of propellant required for LOI can be calculated, or at least closely estimated with a small margin of error. There are definite right and wrong answers to these questions and are not subject to opinion. In most cases gaetanomarano's numbers are simply wrong.

    The part I don't understand about gaetanomarano's stubbornness is that he could easily use corrected facts and figures and still make his argument for the use of the big SM, though perhaps some modification might be necessary. Doing so would make his argument far more creditable, but instead he clings to fantasy numbers. He is digging in his heels over the wrong stuff and losing all credibility in the process.

  5. #35
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    Quote Originally Posted by Bob B. View Post
    And hopefully other people learn something as well.
    Yes, both in rocket technology and some amazing insights in human (well, Italian actually) psychology! I really enjoy following these discussions and I admire the amounts of time and energy you guys put into this. Both of you, thanks!

  6. #36
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    Quote Originally Posted by gaetanomarano View Post
    well... 66 mT of total Orion/SM/LSAM mass - 44 mT for the Orion/SM (from your calculation) = 22 mT LSAM - 10.8 mT of the ascent stage = 11.2 mT for the descent stage weight - 2.5 mT of LSAM cargo (rovers, etc. that are in the descent stage) = 8.7 mT - the weight of the tanks, DS engines, landing pads, LSAM structure, etc. (maybe 5 mT) = 3.7 mT !!!!
    if you're right about the LOI propellant, the consequence is that the LSAM must land on the moon with 3.7 mT of propellant!!!
    it's not a landing!
    it's a miracle!
    And this is the problem with your plan; the entire LSAM has to be downsized and reproportioned (ascent and descent stages). The Apollo LM landed on the moon with a total starting mass of only 15 mT, so we know it can be accomplished with 22 mT. However, making the LSAM smaller may mean a smaller crew, less experiments, fewer tools, etc. You have seriously curtailed the capabilites of each mission, which is why your plan is a bad one.

    And by the way, my 44 mT figure was revised to 43.4 mT and was based on a 69 mT convoy. If the convoy is only to 66 mT, then my estimated mass for the CEV drops to about 42.5 mT. This makes the LSAM 23.5 mT, not 22 mT.

  7. #37
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    For the benefit of those who may listen, Iíd like to further explain the situation regarding the heavier Launch Abort System (LAS) and how it affects the amount of payload that can be carried. Let me start with an analogyÖ

    Letís say a truck will be carrying a load of cargo between two cities. The truck has just enough fuel to make the trip and there are no refueling stations along the route. Suppose at the last minute a piece of critical cargo shows up that has to be added to the truck. Since adding weight will decrease fuel mileage, something has to be removed to make room for the new cargo or else the truck wonít have enough fuel to make it to the destination. Letís say the new cargo weighs two tons, we therefore have to remove two tons of less critical cargo so the gross weight of the truck remains the same.

    Letís now say the last-minute cargo wonít be going all the way to the final destination; it will be dropped off at an outpost halfway along the route. Had we removed a full two tons of payload to make room for the new cargo, we will now be traveling two tons light during the last half of the trip. This will cause an increase in fuel mileage and we will arrive at our destination with excess fuel, which could have been use to haul more payload.

    The proper course of action is to remove only as much payload as necessary to allow the truck to reach its final destination using the last of its fuel. Letís say we off loaded one ton of payload. We will now be traveling one ton heavy during the first half of the trip and one ton light during the last half of the trip. The fuel mileage we lose on the first half of the trip will be made up for on the last half of the trip.

    The situation with the Ares I is very similar. The payload performance to low Earth orbit is 22.0 mT with a 4 mT LAS. When the LAS mass is increased to 6 mT something must be done to compensate. If everything else remains the same, the payload will have to be lightened. If the LAS where carried all the way to orbit, the payload would have to be reduced by 2 mT to offset the 2 mT increase in the LAS. However the LAS is carried along for only 3 minutes of the 8 minute launch. The amount of payload reduction required to offset the heavier LAS is much less than the magnitude of the LAS increase.

    Consider now that a launch vehicle must produce a certain amount of velocity to place a payload in orbit. Velocity is the product of acceleration and time, and acceleration is force divided by mass. If we add mass to the rocket we decrease the acceleration, which will result in a slower velocity.

    During the early part of the launch, while the LAS is attached, the rocket will be running slightly heavy because of the larger LAS. During the last part of the launch, after LAS jettison, the rocket will be running a little light because of the small decrease in payload. At three minutes into the flight the rocket will be traveling a little slower than initially designed (using the lighter LAS), however it will gain back this velocity during the last five minutes.

    Now consider that during those first few minutes of launch the rocket is at its biggest and heaviest. The extra 2 mT we are carrying because of the larger LAS (less the reduction in payload) is only a very small fraction of the total vehicle mass. Since acceleration is inversely proportional to mass, a small fractional increase in mass results in a proportionately small decrease in acceleration.

    During the latter part of the launch the rocket has consumed most of its propellant and has jettisoned the first stage and LAS. It is now many times lighter than it was during the early launch phase. It is during this time we need to gain back the velocity we lost during the first few minutes. We do this by increasing acceleration, but since the rocket is now much less massive, we can achieve the required acceleration with a very small decrease in mass. The decrease in acceleration caused by the extra 2 mT early in the flight can now be offset with a decrease of only a couple hundred kilograms of mass.

    Assuming all other things remain equal, my calculations indicate that the necessary decrease in Ares I payload resulting from the heavier LAS is not more than 200 kg. The Ares I should therefore have no problem inserting a 21.8 mT payload into a 220 nautical mile, 28.5-degree orbit.

  8. #38
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    Quote Originally Posted by Bob B. View Post
    ...the entire LSAM has to be downsized and reproportioned...
    why?

    the LSAM is designed with sufficient propellant for lunar landing/departure + LOI of the full lunar convoy
    if you trasfer the LOI propellant weight (with some adjustments) from LSAM tanks to SM tanks (and reduce the LSAM tanks) the lander module become a double-sized LEM with sufficient fuel for its mission without any change of the size and a weight around 33-35 mT

    the LSAM can't land because you have moved too much propellant weight from LSAM to SM

  9. #39
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    Quote Originally Posted by gaetanomarano View Post
    why?

    the LSAM is designed with sufficient propellant for lunar landing/departure + LOI of the full lunar convoy
    if you trasfer the LOI propellant weight (with some adjustments) from LSAM tanks to SM tanks (and reduce the LSAM tanks) the lander module become a double-sized LEM with sufficient fuel for its mission without any change of the size and a weight around 33-35 mT
    We've been through this a dozen times; this will be the last because I won't explain it again.


    The total convoy is 66 mT.

    Orion is about 21 mT.

    Therefore the LSAM is, 66-21 = 45 mT.

    The maximum LOI delta-v is 1,100 m/s

    The amount of LSAM propellant required to produce 1,100 m/s is, 66*0.2200 = 14.5 mT.
    ** The 0.2200 factor was derived in post #29

    Deleting this propellant from the LSAM allows us to downsize the tanks, call it a 3 mT savings.

    Therefore the revised LSAM mass is, 45-14.5-3 = 27.5 mT.

    The amount of Orion propellant required to produce 1,100 m/s is, 66*.2953 = 19.5 mT.
    ** The 0.2953 factor was derived in post #29

    Adding this propellant to Orion requires enlarging the tanks, call it a 2 mT add.

    Therefore the revised Orion mass is, 21+19.5+2 = 42.5 mT.

    The mass of the convoy is now, 27.5+42.5 = 70 mT

    Since the maximum mass injected by the Ares V is 66 mT, we are 4 mT overweight.

    We have to deduct 4 mT from the LSAM to get within the total mass limit.

    The final LSAM mass is, 27.5-4 = 23.5 mT


    Quote Originally Posted by gaetanomarano View Post
    the LSAM can't land because you have moved too much propellant weight from LSAM to SM
    I didn't do that, you did. I've only calculated what must result from the decisions you've made.

  10. #40
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    Quote Originally Posted by Bob B. View Post
    ...During the early part of the launch, while the LAS is attached, the rocket will be running slightly heavy because of the larger LAS. During the last part of the launch, after LAS jettison, the rocket will be running a little light because of the small decrease in payload. At three minutes into the flight the rocket will be traveling a little slower than initially designed (using the lighter LAS), however it will gain back this velocity during the last five minutes.
    your example appears logical ...with trucks... but the problem is that NASA engineers have calculated the J-2x CLV performaces with only (22+4) 26 mT at lift-off and with the acceleration given with that mass
    if you change the flight profile the rocket may not work properly or (simply) reach a lower orbit
    you say that (from your calculations) the rocket will have the right performances also with two extra tons at lift-off
    but, if that is true, why NASA has not written "24 mT" (+4) of MAX payload (instead of only 22 mT) in the ESAS CLV table???
    maybe, you are right... but I prefer to wait the real NASA specs or (better) the real TEST flight data...
    before new NASA data or REAL tests, I still think that (if NASA specs are 22+4 mT) with a 2 mT bigger LAS the Orion/SM mass must be 20 mT max to have the same performances of NASA evaluations

  11. #41
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    Quote Originally Posted by Bob B. View Post
    For the benefit of those who may listen, Iíd like to further explain the situation regarding the heavier Launch Abort System (LAS) and how it affects the amount of payload that can be carried. Let me start with an analogyÖ
    I snipped the analogy (seemed redundant to post it again). Excellent job, though - very clear and easy to understand.

    Thanks.

  12. #42
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    Quote Originally Posted by Bob B. View Post
    We've been through this a dozen times; this will be the last because I won't explain it again.
    you're not in condradiction with me but with NASA on (both) the (planned) LSAM and the REAL Apollo/LEM flights

    now the LSAM...

    these are the LSAM data you've posted: ascent stage mass of 10,809 kg. descent stage mass of 35,055 kg. (45,864 kg. total)

    well, just imagine we have a LEM-like LSAM in lunar orbit (no matter how it is arrived there) that (like a LEM) can/must perform ONLY the descent/ascent flights and has NO extra-propellant for LOI

    how much that LEM-like LSAM must weigh?

    clearly, the AS/DS proportion will be not (exactly) "the same" of the LEM, but may be very close to it, since the LSAM has more efficient DS propellant but (also) bigger structure and more cargo weight (2.5 mT planned) then, since we don't have (nor NASA has now) full data and details of the (unexisting) real LSAM, we can assume that more weight and more efficient fuel will compensate and the AS/DS proportion will be similar to the Apollo LEM

    since the LSAM/AS is 10.8 mT, the new DS mass of the (LEM-like-lunar-only-no-LOI) LSAM may be around 24 mT and the total AS/DS weight may be around 34.8 mT

    that means NASA has allocated (45.8 - 34.8 =) 11 mT of LOX/LH2 propellant (and extra-tanks) for LOI

    then, we must move 10.5 mT of propellant and 0.5 mT of extra-tank weight (your evaluation of tank's weight seems excessive) from the LSAM to the SM adding the extra-fuel mass for the less efficient hypergolic (about +30%) and another 0.5 mT of extra SM tanks weight (3 mT is the dry mass of the full SM, then, can't be the extra weight of a simple tanks' resizing!) for a total of 10.5 x 1.3 + 0.5 = 14.15 mT (14 mT rounded)

    if we add the extra weight to your evaluation of the standard Orion+SM we have 21 mT + 14 mT = 35 mT

    if we add it to my evaluation of the base vehicle we have 19.5 mT + 14 mT = 33.5 mT

    not 44 mT

    .
    Last edited by gaetanomarano; 2006-Sep-22 at 02:36 AM.

  13. #43
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    Quote Originally Posted by gaetanomarano View Post
    (your evaluation of tank's weight seems excessive)
    You are right about that. I just read in the ESAS report that the LSAM will use a pump-fed system. This means its tanks will be lighter than I estimated. The Orion system will be pressure-fed, therefore the tanks will be relatively heavy. I revise my numbers as follows:


    "The total convoy is 66 mT.

    Orion is about 21 mT.

    Therefore the LSAM is, 66-21 = 45 mT.

    The maximum LOI delta-v is 1,100 m/s

    The amount of LSAM propellant required to produce 1,100 m/s is, 66*0.2200 = 14.5 mT.

    Deleting this propellant from the LSAM allows us to downsize the tanks, call it a 1.5 mT savings.

    Therefore the revised LSAM mass is, 45-14.5-1.5 = 29 mT.

    The amount of Orion propellant required to produce 1,100 m/s is, 66*.2953 = 19.5 mT.

    Adding this propellant to Orion requires enlarging the tanks, call it a 2 mT add.

    Therefore the revised Orion mass is, 21+19.5+2 = 42.5 mT.

    The mass of the convoy is now, 29+42.5 = 71.5 mT

    Since the maximum mass injected by the Ares V is 66 mT, we are 5.5 mT overweight.

    We have to deduct 5.5 mT from the LSAM to get within the total mass limit.

    The final LSAM mass is, 29-5.5 = 23.5 mT"



    The tank estimates may still be a little on the high side, but it makes little difference in the final analysis.


    Quote Originally Posted by gaetanomarano View Post
    that means NASA has allocated (45.8-34.8) 11 mT of LOX/LH2 propellant (and extra-tanks) for LOI
    No, they have not.

    The ESAS report says NASA is basing the LSAM descent stage on the R-10 family of engines. There are several variants of this engine so I donít know exactly what the specific impulse will be, but the RL-10 is a high specific impulse pump-fed LOX/LH2 engine. The 451.5 sec Isp used in my previous calculations is right in line with this engine's expected performance.

    The ESAS report also says 1,100 m/s is the maximum LOI delta-v and 1,900 m/s delta-v has been budgeted for descent. When the LSAM performs LOI, 22% of the initial convoy mass must be LOI propellant (see post #29), thus

    Propellant for LOI = 66,000*0.2200 = 14,520 kg

    This makes the remaining mass of the LSAM,

    45,000 Ė 14,520 = 30,480 kg

    The amount of propellant required for descent is therefore,

    Mf = Mo*e^-(dV/(Isp*g))
    Mf = 30,480*e^-(1,900/(451.5*9.80665))
    Mf = 19,845 kg

    Propellant for descent = 30,480 Ė 19,845 = 10,635

    Therefore the total propellant in the descent stage for both LOI and lunar descent is,

    Descent stage propellant = 14,520 + 10,635 = 25,155 kg


    So just how good is my estimate? Table 4-23, Page 168 of the November-2005 Final ESAS Report lists the descent stage propellant as 25,105 kg.


    If I'm so wrong, as you seem to think, why is my number within 0.2% of NASA's calculations?
    _

  14. #44
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    and where are G's similar numbers?
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    Quote Originally Posted by captain swoop View Post
    and where are G's similar numbers?
    The ESAS report doesn't give G's for the LSAM, or at least not that I've seen, but we can easily calculate it. The report says the descent stage will have four 66.7-kN (15,000 lbf) engines. That's a total thrust of 266.8 kN, or 27,200 kgf.

    For LOI the mass varies from 66,000 kg to 51,480 kg (my numbers), therefore

    G @ start of burn = 27,200/66,000 = 0.41 g

    G @ end of burn = 27,200/51,480 = 0.53 g

    During descent the mass varies from 30,480 kg to 19,845 kg, therefore

    G @ start of burn = 27,200/30,480 = 0.89 g

    G @ end of burn = 27,200/19,845 = 1.37 g (100% thrust)

    Of course the engines won't be firing at 100% thrust near the end of the descent burn because they need only to counteract gravity so the LSAM can hover and set down gently. The "weight" of the LSAM in lunar gravity at touchdown will probably not be less than about 33 kN (unless we use gaetanomarano's downsized LSAM). Hovering will therefore require 1/8th thrust, or two engines at 25% thrust.

  16. #46
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    Quote Originally Posted by Bob B. View Post
    If I'm so wrong,
    it's too easy to give different evaluations about the LSAM since it don't exist now and all data about it are hypotetical

    clearly, not even NASA knows the real LSAM data since we know (now) that (both) Orion and Ares-I real data are (or will be) different from early ESAS plan evaluated specs

    so, I try here another approach to the problem from REAL Apollo missions (that have successful accomplished their moon missions!) using their TRUE specs (that can't be changed or modified or "evaluated" or "calculated" again)

    these are the Apollo specs:

    Apollo CSM mass: 30.3 mT
    Apollo LM mass: 14.7 mT
    Apollo CSM+LM mass: 45 mT

    Apollo CM mass: 5.8 mT
    Apollo SM mass: 24.5 mT
    SM propellants mass: 18.4 mT
    SM propellants: N2O4/UDMH

    you know that with the Apollo missions (both) LOI (of the entire CSM+LM convoy) and TEI (of the CSM alone) was performed ONLY with the SM propellant and that propellants mass was 18.4 mT (a value that can't be changed or evaluated or "calculated" again)

    the total CSM+LM mass at LOI was 45 mT while the CSM mass at TEI was 30.3 mT, LESS the propellants used for LOI

    you have evaluated the Orion+SM propellants mass for TEI at 9.5 mT and the Apollo CSM dry mass was 5-10% (edited figure) more than the (expected) Orion+SM dry mass

    however, since we don't know (exactly) how much propellants was used for Apollo LOI and TEI we can divide it this way:

    SM propellants mass for LOI: 10 mT
    SM propellants mass for TEI: 8.4 mT

    we can't add more to the LOI propellants' figure since the TEI propellants' figure already is too little compared with the TEI propellants mass for the (lighter) Orion TEI

    then, the REAL Apollo SM used (about) 10 mT of hypergolic propellants to brake a 45 mT convoy to LOI

    since the 66 mT Orion+SM+LSAM convoy has +45% extra mass than Apollo CSM+SM, the total (hypergolic, not LOX/LH2) propellants' mass it needs for LOI cant' be more than 14.5 mT (+0.5 mT of extra tanks weight)

    I think that Orion-SM engines will be more efficient than old Apollo engines (so, we must cut 5-10% from propellants' mass value) however, also with that extra propellants' mass, the bigOrion may be around 36 mT (from your 21 mT evaluation of the standard Orion+SM) or 34.5 mT (from my 19.5 mT evaluation)

    not 44 mT (you put 8-10 mT in excess on the bigOrion)

    .
    Last edited by gaetanomarano; 2006-Sep-22 at 02:20 PM.

  17. #47
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    Quote Originally Posted by gaetanomarano View Post
    ...
    however, since we don't know (exactly) how much propellants was used for Apollo LOI and TEI we can divide it this way:
    ...
    WE? Do you have a mouse in your pocket?

    Apparently you don't wish to do any research. I am not an expert, and out of curiosity, I decided to see if this information is available. So, by doing a quick google search on some related words I found this , which includes a budget of both propellents at each stage of the A-15 trip in detail, and weight losses during various maneuvers.

    If I were familiar with the technical terms, I'm sure I could come up with a whole lot more details, and zero in on the exact numbers that your arguments require. This is something that others on this board can do very well, and have demonstrated time, and time again.

    You have lost my interest long ago, mainly because I don't have the technical background to debate you. But; as a non-expert, it is still clear to me that you are blowing smoke with your "opinions". You are blurring the distinction between known facts, estimates, and opinions. And each time a fact comes out, you start yelling "It's my opinion", and start to morph that opinion back into what you call a fact, until someone brings out another fact, and the cycle starts all over again.

    I am very suprised I have not seen moderator comments in some of these threads. I have seen many direct questions directed toward you that remain unanswered, or answered in such a vague way as to being almost useless.

    I'm sorry if this sounds harsh, but the steam has been building since my last post, and I thought I could hold it in.

  18. #48
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    Quote Originally Posted by gaetanomarano View Post
    however, since we don't know (exactly) how much propellants was used for Apollo LOI and TEI we can divide it this way:

    SM propellants mass for LOI: 10 mT
    SM propellants mass for TEI: 8.4 mT
    No, you canít divide it that way because the Apollo SPS (service propulsion system) performed more maneuvers than just LOI and TEI. There were lunar orbit altitude changes, orbit plane changes, and course corrections. Also the mass of the CSM at TEI was much less than the CSM+LM at LOI, therefore much less propellant is required for TEI than for LOI. Just doing a quick calculation, I estimate the breakdown is closer to:

    SM propellants mass for LOI: 11.0 mT
    SM propellants mass for TEI: 4.5 mT
    SM propellants mass for other: 2.9 mT

    I have in the past performed the calculations for Apollo using the same equations and methods Iíve been using in this thread, and the results match very well with the actual Apollo data. When I have more time Iíll reproduce the analysis here. It will confirm the correctness of my method.

    Quote Originally Posted by gaetanomarano View Post
    then, the REAL Apollo SM used (about) 10 mT of hypergolic propellants to brake a 45 mT convoy to LOI

    since the 66 mT Orion+SM+LSAM convoy has +45% extra mass than Apollo CSM+SM, the total (hypergolic, not LOX/LH2) propellants' mass it needs for LOI cant' be more than 14.5 mT (+0.5 mT of extra tanks weight)
    You are failing to consider the magnitude of the velocity change at LOI. ESAS is designed to allow access to the Moonís polar regions. Inserting into an orbit with the inclination needed to reach these areas requires a high delta-v. The LSAM will be designed for this worse-case scenario -- about 1,100 m/s delta-v -- and that is what Iíve been basing my calculations on. The minimum delta-v for insertion into a low inclination orbit is about 845 m/s (per the ESAS report). The Apollo orbits were all near the equator, thus required smaller delta-v than the LSAM will be required to produce.

    Less delta-v means less propellant. If I recalculate the Orion+LSAM LOI propellant requirement for the minimum scenario, then only about 10.8 mT is required (LOI performed by LSAM). Since the LSAMís propellant load will be partly reduced due to the lower delta-v requirement, the total initial convoy mass is about 62.3 mT. The fraction of LOI propellant to convoy mass is then,

    Apollo SPS: 11.0/45.0 = 0.24
    LSAM, minimum = 10.8/62.3 = 0.17
    LSAM, maximum = 14.5/66.0 = 0.22

    Apollo required about 50% more propellant than the LSAM to insert the same amount of mass into the same orbit.

  19. #49
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    Quote Originally Posted by NEOWatcher View Post
    ...WE?...
    thank you very much for the link!

    I've searched the right data on Google but I've not found your document (probably I've not used the right keywords)

    clearly, the LOI/TEI propellants' weights was not "unknown" to NASA, but, since I've not found them, I've supposed this data as not available on internet (that the reason I've used "we")

    from page 5-25 table (below) the SM propellants' mass to brake the 45 mT Apollo convoy to LOI was 11 mT (that's very close to my "10 mT" evaluation!)

    that means the (bigSM) hypergolic propellants' mass to brake the Orion/LSAM convoy to LOI may be around 14.3 mT (pretty close to my evaluation!!!)

    table.jpg

    I am very suprised I have not seen moderator comments in some of these threads.
    why they must comment my posts?
    I've not the Bob.B's enginneers' skill (so I can't post similar calculations) but I explain (the best way possible) my proposals and (very important) I don't post insults and PERSONAL ATTACKS against anyone!

    .

  20. #50
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    Quote Originally Posted by gaetanomarano View Post
    from page 5-25 table (below) the SM propellants' mass to brake the 45 mT Apollo convoy to LOI was 11 mT (that's very close to my "10 mT" evaluation!)
    Hmm... that is EXACTLY my evaluation. Also note that the propellant usage for TEI was 4.5 mT. Funny how that happened.

    Quote Originally Posted by Bob B. View Post
    Just doing a quick calculation, I estimate the breakdown is closer to:

    SM propellants mass for LOI: 11.0 mT
    SM propellants mass for TEI: 4.5 mT
    SM propellants mass for other: 2.9 mT

  21. #51
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    Quote Originally Posted by Bob B. View Post
    ...the mass of the CSM at TEI was much less than the CSM+LM at LOI, therefore much less propellant is required for TEI than for LOI...
    you're right on that, I've thought the same thing (the CSM at TEI was half the full convoy weight at LOI) and I've started to write a new post (with less TEI propellants' mass and more LOI propellants' mass) but NEOWatcher has anticipated my post with his excellent and very useful link (that confirms my evaluations about the Apollo hypergolic propellants' mass used for LOI)

    now we know (from the document's table) that the LOI propellants' mass was 11 mT and the TEI propellants' mass only 4.5 mT

    since the full Apollo SM propellants' mass was 18,4 mT (less the LOI/TEI mass) the residual propellants' mass was 2.9 mT but only (about) 0.8 mT was used "for other" since the propellants' mass after TEI was 2.1 mT

    maybe, part of this (2.1 mT) residual propellants was used to correct the CSM trajectory in the moon-earth flight, but great part of it was (probably) only an extra-mass loaded to give some redundancy

    the FIRST consequence of these exact (and REAL) Apollo's data is that a +45% bigger (66 mT) Orion/LSAM convoy needs only 14.3 mT of hypergolic propellants' mass for LOI

    the SECOND consequence is that (both) your and my evaluations about the SM propellants' mass for Orion's TEI was WRONG

    the Apollo CSM dry mass at TEI was 11.9 mT while the Orion+SM dry mass is around (8.5 mT+3 mT) 11.5 mT (or 12 mT with 0.5 mT of extra tanks' weight)

    like with Apollo, we can add to the Orion's SM some extra-propellants (for redundancy and maneuverings) so, the total (standard Orion's) SM propellants' mass (for TEI) may be around 6.5 mT

    that means the standard (TEI-only) Orion+SM mass will be 11.5 + 6.5 = 18 mT (not 19.5 mT of my evaluation nor 21+ mT of your evaluation) while the (LOI+TEI) bigOrion mass may be around 11.5 mT + 0.5 mT + 6.5 mT + 14.3 mT = 32.8 mT (pretty close to may evaluations!) and the full Orion/LSAM convoy mass around 67.5 mT

    also, the 18 mT (Orion/SM) + 0.6 (2nd stage adapter) + 6.2 mT (LAS) = 24.8 mT perfectly matches the MAX (22+4 mT) mass the J-2x Ares-I can lift from launch pad ground!

    ...ESAS is designed to allow access to the Moonís polar regions...
    1st... the polar option was scrapped from the ESAS plan months ago (if I find again the articles I'll post their links here)

    2nd... you can't know things not even NASA knows now (like, where the moon missions will really land)

    3rd... we are talking (and evaluated/calculated) the vehicles' weights and propellants' mass for an Apollo-like Orion/LSAM mission (with a small SM or a big SM) and the (extrapolated) REAL data are close to my evaluations!

    Apollo required about 50% more propellant than the LSAM to insert the same amount of mass into the same orbit.
    if that's true... that means the smallOrion and ("my") bigOrion may be (both) much lighter than we have calculated/evaluated so far!

    .

  22. #52
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    Quote Originally Posted by gaetanomarano View Post
    that means the (bigSM) hypergolic propellants' mass to brake the Orion/LSAM convoy to LOI may be around 14.3 mT (pretty close to my evaluation!!!)
    How do you figure that?

    11/45 = X/66

    X = 66*11/45 = 16.1, not 14.3.

    For a low delta-v LOI this is about right, though a little on the high side because the Apollo SPS was slightly less efficient than the AJ10-118K. The actual number for the minumum delta-v, if performed by the Orion CEV, is

    LOI propellant = Mo - Mo*e^-(dV/(Isp*g))

    LOI propellant = 66 - 66*e^-(845/(320.5*9.80665)) = 15.56 mT

  23. #53
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    Quote Originally Posted by Bob B. View Post
    Hmm... that is EXACTLY my evaluation. Also note that the propellant usage for TEI was 4.5 mT. Funny how that happened.
    ------------------
    Quote:
    Originally Posted by Bob B.
    Just doing a quick calculation, I estimate the breakdown is closer to:

    SM propellants mass for LOI: 11.0 mT
    SM propellants mass for TEI: 4.5 mT
    SM propellants mass for other: 2.9 mT
    -------------------

    NO!!!!

    in your previous posts you have evaluated at 9.5 mT the Orion propelllants' mass for TEI

    the 4.5 mT value you posted now don't come from your calculations but from a table of the document NEOWAtcher linked here (that's is also the source of my recent posts)

    also, if you have (really...) evaluated the Apollo SM hypergolic propellants' mass for LOI at 11 mT... the hypergolic propellants' mass for LOI of the Orion convoy must be 14 mT, not 21+ mT (as you've calculated and posted so far)

    .
    Last edited by gaetanomarano; 2006-Sep-22 at 05:03 PM. Reason: grammar

  24. #54
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    Quote Originally Posted by Bob B. View Post
    How do you figure that?
    I've multiplied by 1.45 the Apollo mass, but, right, 16 mT is more correct, so, the bigOrion mass may be around 34.5 mT (not 44 mT)

    .

  25. #55
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    Quote Originally Posted by gaetanomarano View Post
    3rd... we are talking (and evaluated/calculated) the vehicles' weights and propellants' mass for an Apollo-like Orion/LSAM mission (with a small SM or a big SM) and the (extrapolated) REAL data are close to my evaluations!
    I haven't been talking about that. I've been basing my evaluations on the worse-case requirements as outlined in the Nov-2005 ESAS Report. I've made no attempt to compare Orion/LSAM to Apollo until you brought it up; and then I told you that they were not comparable because the ESAS report has more demanding requirements.

    If I rerun my calculations based on an Apollo-like mission then our numbers won't be nearly as far apart as they currently are. Reducing the LOI delta-v will have a significant effect on the results. However, Apollo missions were limited in how much lunar surface they could cover. I can't believe we'll go back to the Moon with the same limitations. Expanding the coverage area has got to be a ESAS objective.

  26. #56
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    Quote Originally Posted by Bob B. View Post
    ...basing my evaluations on the worse-case requirements...
    when?
    I don't remember any claim (like that) from you before your 9.5/21 mT calculations!
    that seems me a "last-minute excuse" for your wrong calculations... but you can link here the posts where you talk of "worst-case" for your over-sized calculations and results

    and, about the ESAS plan... probably you know it's changed very much in latest months...

    .

  27. #57
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    Quote Originally Posted by gaetanomarano View Post
    NO!!!!

    in your previous posts you have evaluated at 9.5 mT the Orion propelllants' mass for TEI

    the 4.5 mT value you posted now don't come from your calculations but from a table of the document NEOWAtcher linked here (that's is also the source of my recent posts)

    also, if you have (really...) evaluated the Apollo SM hypergolic propellants' mass for LOI at 11 mT... the hypergolic propellants' mass for LOI of the Orion convoy must be 14 mT, not 21+ mT (as you've calculated and posted so far).
    Everything I've posted prior to now was for ORION, this is the first time I've calculated or posted anything about APOLLO. What I posted previously about Orion is irrelevant. Furthermore, I hadn't even seen NEOWAtcher post until after I posted my Apollo propellant calculations. NEOWAtcher posted while I was writing my response and performing my calculations.

    You better not be calling me a liar!!!

  28. #58
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    Quote Originally Posted by Bob B. View Post
    ...was for ORION, this is the first time I've calculated or posted anything about APOLLO...
    the hypergolic propellants' mass to brake a given mass to LOI doesn't change so much between Apollo and Orion, then, the REAL Apollo data (from REAL and successful flights!) are very good to know if some "calculated" data (about the propellants' mass to brake a bigger vehicle) are right or wrong

    in your previous posts, you've calculated the TEI (not Polar!) hypergolic propellants' mass for the (11.5 mT dry mass) Orion/SM at 9.5 mT and the hypergolic propellants' mass to brake the "66 mT" Orion/LSAM convoy (no matter if the LSAM will be able to go Polar or not!) to over 21 mT (and the extra-tanks increase/reduction to 2 mT and 3 mT, more than the SM total mass!) calculations that (now) are (clearly) WRONG under every point of view!

    I haven't called you a "liar" since (clearly) I can't know when you have written your (new) "4.5" post... then, I'm sorry ...but my mistake was due to the fact "your" new data was so close to the (not easy to find) data of the NEOWatcher's link (posted 30 minutes before your post...)

    .
    Last edited by gaetanomarano; 2006-Sep-22 at 05:25 PM. Reason: grammar

  29. #59
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    Quote Originally Posted by gaetanomarano View Post
    when?
    I don't remember any claim (like that) from you before your 9.5/21 mT calculations!
    that seems me a "last-minute excuse" for your wrong calculations... but you can link here the posts where you talk of "worst-case" for your over-sized calculations and results

    and, about the ESAS plan... probably you know it's changed very much in latest months...
    LOL! You sure haven't been paying much attention, have you?


    Quote Originally Posted by Bob B. View Post
    The ESAS report gives a LSAM ascent stage mass of 10,809 kg (page 166) and descent stage mass of 35,055 kg (page 168). The total CEV+LSAM mass as given in the ESAS report is therefore 69 mT. The LOI delta-v is given as 845 m/s minimum to 1,100 m/s maximum (page 167), depending on the location of the landing site.

    Slowing an initial mass of 69 mT by 1,100 m/s using hypergolic propellant will require about 21 mT of propellant. This mass has to be added to the CEV plus another couple tonnes for larger tanks. This puts the total mass of the CEV at around 44 mT.
    Quote Originally Posted by Bob B. View Post
    All that we need to determine the ratio of the LOI propellant to the total convoy mass is the change in velocity and the specific impulse of the engines. We know from the ESAS report that the maximum expected delta-v is 1,100 m/s. We donít know the specific impulse of the LSAM engines, but we know the propellant is LOX/LH2. LOX/LH2 engines have vacuum specific impulses well over of 400 seconds. For instance, the J-2X is listed as having an Isp of 451.5 sec. Letís use this value for our calculation.
    Quote Originally Posted by Bob B. View Post
    The total convoy is 66 mT.

    Orion is about 21 mT.

    Therefore the LSAM is, 66-21 = 45 mT.

    The maximum LOI delta-v is 1,100 m/s

    The amount of LSAM propellant required to produce 1,100 m/s is, 66*0.2200 = 14.5 mT.

  30. #60
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    Quote Originally Posted by Bob B. View Post
    ...Furthermore, I hadn't even seen NEOWAtcher post until after I posted my Apollo propellant calculations. NEOWAtcher posted while I was writing my response and performing my calculations...
    Which would only be possible if Bob were pulling numbers out of thin air. But, I've been keeping track of the players.

    Although, this has all been very aggrevating, I've been learning alot.

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