You start off with a false comparison. The SM has squat to do with mission capability, which is a function of the lander, so you've already dismantled the credibility of the argument.
Gaetanomarano, you are still insisting the small SM weighs only 10 mT and has only 7 mT of propellant. The equations indicate this is impossible, which I have attempted to show you on more than one occasion in another thread. Orion simply cannot produce the published delta-v out of only 7 mT of propellant. You are certainly entitled to your opinion and are free to write as many articles as you want, but please stop repeating figures that demonstrably wrong. Doing so only make you look dishonest. A more reasonable estimate is 13 mT total mass with 10 mT propellant.
however, the point of my comparison is NOT between 7 mT or 10 mT of propellant, but between "TEI-only propellant SM" (the "corks screw Orion") and "TEI+LOI propellant SM" (the "swiss knife Orion")
then, also if I accept your evaluation (10 mT of SM propellant) that doesn't change nothing of the question explained in my article (smallSM or bigSM?)
Based on the most recent information published by NASA, we know the following:
CM dry mass: 7,891 kg
CM launch mass: 8,485 kg
SM propellant: N2O4/MMH
SM delta-v: 1,855 m/s
This is enough information to make a fairly good estimate of the SM propellant load, which Mark Wade and I have already done. Those calculations show that 9.5 to 10 mT is a reasonable estimate. Your figure of 7 mT is simply not creditable.
- LOI of the Orion+SM alone
- LOI of the Orion+SM+LSAM
- TEI of the Orion+SM
- earth/lunar orbit maneuvering
- ISS/LSS orbital reboost
- earth/lunar orbit change
- earth reentry speed braking (optional)
smallOrion and bigOrion are like two airplane with the same number of seats, but, the first, with small fuel tanks, to fly only from city to city, and, the latter, with three times larger tanks to fly between continents
which do you buy for your own airline company?
Last edited by Bob B.; 2006-Sep-19 at 05:14 PM. Reason: spelling
if a standard Ares-I + (cargo or crew) Orion launch will cost (e.g.) $800M and a bigSM Orion + enhanced-Ares-I launch cost $1.2B, the money saved will be OVER $3.6 BILLION PER LAUNCH !!! ...at EVERY (cargo/crew combo) launch!!!
as I've explained in this thread and in the SM thread, only the "propellant weight per mission" will be less efficient (only a small amount) NOT the number of missions and the total payload for the given funds!...will add up in the long run, resulting in fewer launches to deliver the same mass to the lunar surface. Therefore, the small SM decreases the cost per kilogram of payload delivered to both LEO and the Moon...
that thanks to a smaller AresV and less AresV/bigEDS/LSAM launches for crew rotation
however, the autonomous Orion has so many advantages that can't be compared with a few tons of extra cargo!
that's true, but this is only an example you can't read "literally" (like your Shuttle's example in another thread) however, the full SM propellant (not used for Mars orbital insertion) can be used to brake the (very high) Mars reentry speed or (if sufficient) for an earth orbit insertion instead of an (extremely critic and dangerous) direct Mars to Earth reentry!...no need for a large SM in this limited role...
Gaetanomarano, the CEV you have advocated building will be on the order of 40 mT. This means the only vehicle capable of launching it to the Moon is an Ares V. Every time you launch a CEV to the Moon it will cost an Ares V. This is an underutilization the Ares V’s payload capability, therefore making a solo CEV launch more expensive than it needs to be. To make the cost of a solo CEV launch more economical would require a downsizing of the Ares V; however this means it can no longer launch a CEV+LSAM, which is its reason for being in the first place. We need the large payload capacity of the Ares V for normal CEV+LSAM missions.
The only other alternative to launch the CEV to the Moon is another launcher smaller than the Ares V but still about 2/3rds the size (let’s call it the Ares IV). Clearly building another big launcher so close to the capacity of the Ares V is tremendous waste – we simply do not need two heavy launch vehicles in that class.
Now, what do we do about LEO missions? Although the CEV may weigh 40 mT for lunar missions, propellant can be off-loaded for ISS flights. It is likely we can get its mass within the 25 mT performance of the Ares I for ISS missions. However, we still need to get the bigger 40 mT payload into LEO for mating with the Ares V/LSAM during lunar landing missions. This means we need a launcher about 5/3rds the size of the Ares I, or about 1/3rd the size of the Ares V (let’s call it the Ares II).
So what are we left with? We certainly can’t afford to build four different launch vehicles customized for each mission type – Ares I for ISS missions, Ares II for launching heavy CEV to LEO, Ares IV for launching heavy CEV to the Moon, and Ares V for launching CEV+LSAM to the Moon. We definitely need the Ares V for full-up lunar landings. And with the bigger CEV, we need the Ares II to get this fully fueled vehicle to LEO. Since this is likely all we can afford to build, we are left with launching ISS missions on a vehicle too big for the task, and likewise, launching solo CEV lunar missions on a vehicle too big for the task.
I see great waste in the plan.
I've accepted your data about the SM weight and propellant only to focus the discussion on the advantages of a bigOrion
but you use your (wrong and not NASA/LM official) evaluations as a base to demonstrate that your opinons are true
well, below, you can see a NASA table available when the (original) ESAS plan was published
you can see that the (deleted) 4-segments SRB + SSME rocket was able to lift up to 27 mT to LEO at 28.5 deg.
the new Ares-I (with a 5-segments SRB and a J-2S derived J-2x) will be able to lift only 22 mT to LEO
this is the reason why the CEV was resized from 5.5 to 5 m. and its final (TOTAL) weight will be 6500 lbs. less than its first version (exactly, 22.6 mT for the Block 2 lunar-CEV)
if the TOTAL WEIGHT of the (original ESAS plan) lunar-CEV was 22.6 mT, HOW the (-6500 lbs. reduced) Orion can weigh 25 mT ?
consider, also, that the original ESAS plan extimated the LAS around 4.1 mT while the real LAS will be 6.2 mT that means the Orion must weigh 2 mT LESS than planned (exactly like the -6500 lbs. reduced Orion)
then, the TOTAL weight of the Orion+SM will be around 19.5 mT and the "new Ares-I" will be (probably) able to launch 25 mT at lift-off INCLUDING the 6.2 mT LAS
and, since the Orion mass is 8.5 mT, the Service Module CAN'T WEIGH more than 11 mT with (about) 8 mT of propellant (now I change again the SM data in my article since MY previous evaluation was, clearly, closer to reality!)
the consequence of these data is that a TEI+LOI Orion+SM can't be the 40 mT monster you claim!
the choice of a bigOrion doesn't affect the normal (Orion+LSAM) moon mission, because I only suggest to transfer the LOI propellant (maybe, around 10 mT) from the LSAM to the Orion
the LSAM-light can be launched with a smaller AresV (maybe, to-day's design, but with 4-seg.SRBs) while the Orion needs an "enhanced" Ares-I, but the total weight of the Orion+SM+LSAM+EDS convoy at TLI will remain THE SAME (around 150 mT) of to-day's plan
the autnomous Orion needs LESS propellant (since it don't need to brake the LSAM to LOI) and can be launched with two (enhanced) Ares-I (one for the 27 mT Orion+SM and one for the 33 mT micro-EDS) or (best) a single 65 mT (half the AresV payload) Ares-II (the latter may be very useful to launch ISS modules/resupply in combo cargo/crew missions)
the smallOrion is NOT the BEST Orion but ONLY the Orion that (for reasons I don't know) you (absolutely!) WANT
Last edited by gaetanomarano; 2006-Sep-19 at 09:31 PM. Reason: grammar
recent rumors say of a possible shift to LOX/LH2 for the LSAM's ascent stage... if that change come true... it's the 1st step for a REFUELABLE/REUSABLE lunar lander... and a reusable lander (absolutely) needs a small LSS and an autonomous Orion for crew rotation...
EDIT: Do you have similar information on the Ares V?
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.
EDIT: My original 40 mT estimate was based on a total CEV+LSAM mass of about 55 mT, which appears too low according to the information in the ESAS report.
Last edited by Bob B.; 2006-Sep-21 at 01:56 PM. Reason: additions are noted
so far, the ONLY true data is the ATK test that give a +5 sec. burning time and +9% of thrust
however, the (-5 mT payload) lower performance of the new Ares-I is due to the shift from SSME to J-2x
I've no further data about the AresV, only the known (130 mT) max payload
I've no doubts about this point, since I've posted the table that confirms it!...a mass of around 21 mT and Mark Wade calculated about 21.5 mT...
the image below is an extract from the ESAS table with the early 5.5 m./22.6 mT Block 2 lunar-CEV and this is the article that reported the -6500 lbs. weight reduction of the new 5 m. CEV...ESAS report I’ve seen gives a mass for the Block 2 Lunar CEV of 23,153 kg...
unfortunately, the article don't specify if the 3 mT reduction was from the early 5.5 m. (22.6 mT) CEV or from the 5 m. CEV (reduced nine months ago) that (already) may give a 1-1.5 mT reduction from the 5.5 m. CEV/SM mass
22.6 mT - 3 mT = 19.6 mT (or 18-18.5 mT if the 3 mT reduction is from the, already reduced, 5 m. CEV) and, since the Orion mass is 8.5 mT, the SM mass will be around 11 mT
in my next post I reply to the other points of your posts
not true...2 mT increase in LAS mass does not require a 2 mT decrease in payload...
the original (SSME) CLV was planned to launch 26.9 mT of payload + 4.1 mT of LAS at lift-off (31 mT total) but the (J-2x) Ares-I is able to lift 5 (NASA ev.) or 5.5 (Wade ev.) or 6 (your ev.) mT LESS than an SSME/CLV (that's only 25-26 mT MAX Orion/SM/LAS weight at lift-off) and, since the real LAS' weight will be 6.2 mT, the MAX Orion+SM mass at lift-off CAN'T BE more than 19 mT (from your ev.) or 19.5 mT (from Wade ev.) or 20 mT (from NASA ev.)
no, according to NASA specs "it looks like Mark an you are wrong"...it looks like Mark’s and my estimate of 21-21.5 mT is right...
since the weights are those I've posted, if your calculations are right (probably) NASA engineers are wrong......the required 1,855 m/s will require about more propellant...
last answers in my next post
since physics' laws and lunar gravity aren't changed from Apollo missions, the descent stage must STILL weigh 69% of the total, then, with a 10.8 mT ascent stage, the weight of the descent stage (excluding the LOI fuel) can't be less than 24 mT
the difference between the LSAM descent stage (planned) weight (35 mT) and the (minimum) descent stage (propellant and structure) weight to land on the moon (24 mT) is 11 mt (not 21 mT) that is the MAX weight of the LOX/LH2 propellant used for LOI of the full Orion/SM/LSAM convoy
if you want to say that "hypergolic is less efficient" etc... think that LOX/LH2 (less dense and cryogenic) needs bigger and heavier tanks + a propellant's burning system and that extra-weight compensate (in part) the weight saved from the less efficient hypergolic
also, don't forget the extra-weight (maybe, around 1 mT) of the descent stage tanks to store the LOI (11 mT) propellant, so, the reall LOX/LH2 propellant weight for LOI may be around 10 mT, NOT 16 mT or 21 mT
if we consider the extra weight of the (less efficient) hypergolic propellant (compensated in part by smaller tanks and no burning system needed) the extra propellant for an Orion LOI with a (resized!) LSAM may be around 12-13 mT, while, the extra propellant for LOI of an Orion alone (without LSAM) may be around 9.5 mT
then, the max weight of an autonomous Orion+SM (from my 19.5 mT evaluation of the ESAS Orion+SM) will be around 32 mT (or 29 mT for the Orion LOI without LSAM) NOT 44 mT !
true, but not 5 mT... maybe around 2 mT ...that is an INCREDIBLY SMALL SACRIFICE compared with the GIANT (safety and operational) advantages of an autonomous Orion!!!...therefore have to steal 5 mT from something else, thus decreasing the payload and the mission capabilities...
only the LSAM will be resized, the EDS must remain the same (and, I'm sure, will be able to launch two extra mTs at TLI)...you still have to inject the same mass to the Moon so the reduction is limited...
true, it may cost more for hardware and (about) the same for R&D but, thanks to the autonomous Orion and the Ares-II, dozens (very expensive) AresV/EDS/LSAM launches can be saved...need something that can launch twice the payload to LEO as the current Ares I design...
the finished ISS will need crew rotations and more resupply (for six astronauts) so, a combo-launch of an Ares-II with one Orion and a cargo module (with 5+ times the Progress payload) will incredibly efficient from the operational (ONE launch insted of SIX) and economical (about $3.6B saved compared with six cargo or crew Orion launches) points of view...then what are you going to use to launch ISS missions...
also, after Shuttles' retirement, the (double-payload) Ares-II can launch new big ISS modules and new hardware (that a standard Ares-I can't launch)
the "lost payload" is only about 2mT, also, 150 mT is the total weight of the last year's ESAS plan lunar-convoy, with a smaller CEV/SM (19.5 mT instead of 22.6 mT) the (2 mT) "lost payload" is (clearly) compensated...if you are satisfied losing 5 mT of lunar orbit payload...
but this is EXACTLY the BEST reason to PRETEND a BIGGER and AUTONOMOUS Orion!!!!!!!!!...consider the possibility Orion may never be more than an ISS shuttle. In that case the small SM is clearly the right choice...
if the Orion will be not able (at least!) to perform (alone) a lunar orbit mission, will be very easy for a future President/Congress to delete the entire moon plan!
while, with an autonomous Orion and (maybe) some lunar orbit infrastructures (best WITH an international cooperation) NASA will be able to launch some moon missions (one per year/two years) with its own annual manned missions' funds, now used to launch the Shuttle!
then, the bigOrion is the BEST choice also from a political-plan-deletion-risk point o view!
Block 2 Lunar Crew
Crew Size: 4
LAS Required: 4,218
Cargo Capability (kg): Minimal
CM (kg): 9,506
SM (kg): 13,647
Service Propulsion System delta-V (m/s): 1,724
EOR-LOR 5.5-m Total Mass (kg): 23,153
SM: 3,000 lbm
CM: 2,000 lbm
LAS: 1,200 lbm
Adapter: 500 lbm
Total: 6,400 lbm
Obviously the individual numbers have been rounded up because the sum of the numbers is greater than the total. Because of the uncertainty in rounding, it appears the CM+SM mass reduction is no more than 5,000 lbm and possibly as little as 4,700 lbm. If we deduct 5,000 lbm from the Block 2 Lunar CEV mass published in NASA’s Final Report dated Nov-2005, we get
CM: 9,506 – 2,000/2.205 = 8,599 kg
SM : 13,647 – 3,000/2.205 = 12,286 kg
Total: 20,885 kg
This is very close to my 21 mT estimate, so the evidence supporting my calculations is starting to mount.
Something else from the article is very useful, it says the SM main engine will be a AJ10-118K. That particular engine has a specific impulse of 320.5 seconds. Knowing this removes much of the uncertainty from the calculations.
If we assume the 20,885 kg total mass is correct, we can now calculate the dry mass of the SM,
dV = Isp*g*LN(Mo/Mf)
1,855 = 320.5*9.807*LN(20,885/(8,485+SM dry mass))
SM dry mass = 3,090 kg
This mass is very close to what we’ve both been assuming. The propellant mass is therefore,
SM propellant mass = 20,885 – 8,485 – 3,090 = 9,310 kg
This is a little less than I previously calculated because the AJ10-118K specific impulse is a little more than I had assumed.
Last edited by Bob B.; 2006-Sep-20 at 08:25 PM.
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
probably the author of the article has (simply) allocated the mass reduction in proportion with the original mass of the single partsInterestingly though, the article breaks down the reductions as follows:
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 don't know this engine (I've read it is derived from DeltaIV 2nd stage) but LM seems want to use a Shuttle OMS derived engine...the SM main engine will be a...
that's exactly what I've said in my postit was a reduction from a vehicle that had grown in mass...
the (6500 lbs.) reduction announced in july is (clearly) the final result from the original ESAS weights including the variations due to diameter and propellant changes
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 don’t 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. Let’s use this value for our calculation. Tsiolkovsky's rocket equation is,
dV = Isp*g*LN(Mo/Mf)
therefore the mass ratio is,
Mo/Mf = e^(dV/(Isp*g))
we then have,
Mo/Mf = e^(1,100/(451.5*9.80665))
Mo/Mf = 1.2820
The inverse of the mass ratio tells us the fraction of initial mass remaining at the end of the burn.
Mf/Mo = 1/1.2820 = 0.7800
We therefore know that 22.00% of the total mass prior to LOI must be LOI propellant. If the total convoy weighs 69 mT, then 15.18 mT must be LOI propellant. If the specific impulse is lower, then more propellant is needed. I estimated 16 mT earlier because I was being a little conservative on the specific impulse, using 425 sec.
We now know that the specific impulse of Orion’s AJ10-118K engine is 320.5 seconds. We can therefore calculate exactly how much propellant is needed at LOI without have to guess or estimate anything.
Mo/Mf = e^(1,100/(320.5*9.80665))
Mo/Mf = 1.4190
Mf/Mo = 0.7047
Therefore the amount of LOI propellant MUST be 29.53% of the total convoy mass. This is an undeniable fact and your whining about it will not change physics. So, if the convoy weighs 69 mT, then the mass of LOI propellant is 20.38 mT.
Let’s assume we have pressurized tanks and double the typical tank mass to account for this. The mass increase and decrease for the CEV and LSAM is then approximately,
CEV: +20.38*1.1 = +22.4 mT
LSAM: –15.18*1.2 = –18.2 mT
The difference is 4.2 mT, which has to be robbed from somewhere else. Since Orion has already been reduced about as much as it can, the 4.2 mT will probably have to come from the LSAM. Of course for every kilogram of mass you take away from the LSAM, less descent propellant will be needed to land on the Moon. Approximately 2.7 mT of the 4.2 mT will be actual payload, with the rest being descent propellant.
Mo = 23*1.4190 = 32.6 mT.
Therefore the amount of LOI propellant is 9.6 mT, so your estimate is pretty close in this case. Of course in reality Orion would carry only enough propellant to perform its required mission. In this case we’ve given Orion a total delta-v of about 2,750 m/s. If less delta-v is required for a specific mission, an appropriate amount of propellant can be off-loaded.
Last edited by Bob B.; 2006-Sep-21 at 03:24 AM. Reason: spelling
both (SSME and J-2x) rockets have a finished amount of upperstages mass they can lift from ground level
if the total upperstages mass the rocket can lift is reduced 5-6 mT and (part) of the upperstages mass (the LAS) is increased 2 mT, the rest of the vehicle (Orion and SM) must weigh less
the LAS is jettisoned after 3 minutes, but, before that event, it is a 6.2 mT weight the rocket must lift from ground, exactly like the Orion and the SM
I've explained that in detail in my post (but probably my english is not so clear)