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

[TrAI]

Heh... First time looking over these posts I got a bit confused trying to figure out what all these wierd inductance measurements was about("T" being the SI symbol for tesla and "m" the prefix for mili-)...

[Bob B.]

Heh... First time looking over these posts I got a bit confused trying to figure out what all these wierd inductance measurements was about("T" being the SI symbol for tesla and "m" the prefix for mili-)...

mT = metric ton

I don't know if that is an internationally recognized abbreviation, but it is the one NASA uses in its ESAS report.

[TrAI]

mT = metric ton

I don't know if that is an internationally recognized abbreviation, but it is the one NASA uses in its ESAS report.

heh, I did figure it out, it was just confusing on the first reading.

The SI unit for the metric ton is a lowercase "t". Of course, over here, when someone says "ton", you can safely assume they are refering to the unit for one megagram, but I see how this could be confusing if there is more than one unit named "ton" in use.

[Bob B.]

the latest figure about the MAX AresV payload to LEO is 130 mT (no more, no less!)

Correct, but that number doesn’t represent what you think it does.

NOR the LSAM + "your" (122.5 mT) EDS since the 167.5 mT total mass EXCEEDS very much the AresV max payload!

It is not “my” number; it comes directly from NASA.

NASA describes their assumptions for the 1.5-launch concept as follows (From NASA-TM-2005-214062, Page 426):

In case 2, LV 27.3 was assumed to have a 44.9-mT LSAM attached to the EDS at launch. The EDS propellant load was also reoptimized for this case. The EDS with LSAM attached then rendezvoused with a CEV on orbit at 160 nmi circular orbit that weighed 20.6 mT, for a total cargo stack mass of 65.5 mT in orbit. The EDS with LSAM and CEV then performed a TLI burn with the remaining EDS propellant. The TLI net payload capability for this case was determined to be 66.9 mT, which is 1.4 mT greater than the required delivery mass of 65.5 mT

NASA then gives the numbers for this concept as follows (From NASA-TM-2005-214062, Page 432):

Figure 6-43. Case 2: LV 27.3 SP + EDS (44.9 mT LSAM)

Vehicle Concept Characteristics
EDS+PL Gross @ Liftoff: 640,281 lbm
EDS Gross @ Liftoff: 541,294 lbm

EDS Stage
Propellants: LOX/LH2
Useable Propellant @ Liftoff: 490,744 lbm
Useable Propellant @ 160 nmi cir.: 219,443 lbm
Stage PMS: 0.9066
Dry Mass: 44,118 lbm
Burnout Mass: 50,494 lbm
# Engines / Type: 2 / J-2S+
Engine Thurst (100%): 274,500 lbf @ Vac
Engine Isp (100%): 451.5 sec @ Vac
Mission Power Level: 100.0%

TLI Delivery
CEV @ Liftoff: 48,061 lbm

LSAM Payload: 98,988 lbm
CEV Payload: 45,415 lbm
Margin Payload: 19,500 lbm
Gross Total Payload: 163,903 lbm

Net Payload: 147,513 lbm
Net Allowable CEV Mass: 48,525 lbm

I call your attention to two numbers:

Useable Propellant @ 160 nmi cir.: 219,443 lbm
Burnout Mass: 50,494 lbm

The burnout mass plus the useable propellant is the total EDS mass. Note that this is the amount of propellant in a 160 nmi circular orbit. The total mass is therefore,

EDS Mass @ 160 nmi cir. = 219,443 + 50,494 = 269,937 lbm, or 122,441 kg

Add this to the LSAM mass and we have the total mass inserted into orbit by the CaLV,

Inserted Mass @ 160 nmi cir. = 269,937 + 98,988 = 368,925 lbm, or 167,342 kg

Add the CEV and we have the total mass prior to TLI (less propellant boil-off),

Mass @ TLI (maximum) = 368,925 + 45,415 = 414,340 lbm, or 187,941 kg

The mass after TLI is the EDS burnout mass plus the LSAM and CEV,

Mass after TLI = 50,494 + 98,988 + 45,415 = 194,897 lbm, or 88,404 kg

The amount of TLI propellant needed can be confirmed by Tsiolkovsky's rocket equation. We already know the final mass and the specific impulse, so all we need is the delta-v, which NASA provides for us (From NASA-TM-2005-214062, Page 151):

The following assumptions were used for calculating global access delta-V costs:

• A TLI delta-V limit of 3,150 m/s is imposed.

Propellant = Mf*e^(dV/(Isp*g)) – Mf

TLI Propellant = 88,404*e^(3,150/(451.5*9.80665)) – 88,404 = 91,666 kg

There MUST be at least this much propellant because it is scientifically impossible for there to be any less. This number matches nicely the on-orbit propellant mass of 99,538 kg (219,443 lbm) given in the ESAS report. The difference is easily explained by boil-off and residuals.


So if the Ares V can lift over 167 mT into orbit, why is its LEO payload performance listed at 130 mT? Before answering this question let's first consider something else.

Based on the above, the CaLV (Ares V) can deliver about 167 mT when hoisting a 44.9 mT LSAM all the way from the ground. When combined with a 20.6 mT CEV launched on a separate rocket, the EDS can deliver 65.5 mT to the Moon. However, this 1.5-launch architecture is not the only mission the CaLV is intended to perform. It is also designed for a 1-launch concept in which it can lift a payload and inject it into a translunar trajectory without the on-orbit rendezvous. In this case the payload lifted from the ground is heavier but the TLI mass is less. The EDS burns more propellant during ascent to deliver the heavier payload to orbit, but then requires less propellant during TLI to inject a mass smaller than the combined LSAM+CEV.

NASA describes this as follows (From NASA-TM-2005-214062, Page 426):

The net payload capability of LV 27.3 plus EDS to TLI is 54.6 mT for the maximum TLI payload carried from liftoff (no orbital rendezvous).

We know from previous numbers the burnout mass of the EDS is 22,904 kg (50,494 lbm), therefore the final mass after TLI for the 1-launch concept is the payload mass plus the EDS burnout mass,

Mass after TLI = 54,600 + 22,904 = 77,504 kg

Thus the amount of TLI propellant required is,

TLI Propellant = 77,504*e^(3,150/(451.5*9.80665)) – 77,504 = 80,364 kg

Therefore the pre-TLI mass, or that inserted into orbit is,

Inserted Mass = 77,504 + 80,364 = 157,868 kg

This number is less than mass injected for the 1.5-launch scenario because the gross liftoff weight is more with the heavier payload (54.6 mT vs. 44.9 mT). Hauling that extra mass all the way from the ground requires more propellant to be burned during ascent, thus resulting a smaller injected mass.


Let’s say now that instead of launching a 54.6 mT payload along with propellant for TLI, we want to deliver one large payload to LEO. This might be the case if we were, for example, delivering a large space station component. In this case the size of the payload we can deliver is approximately equal to 54.6 mT plus the mass of the TLI propellant. Since the TLI propellant is not needed in this case, its mass can be allocated to carrying more payload (note that the spent EDS stage is not considered payload),

Approximate LEO payload = 54,600 + 80,364 = 134,964 kg

Note that this is only approximate because we must also consider the larger payload will require a bigger payload adapter, larger nose shroud, etc. We should therefore expect a somewhat smaller payload.

This is where the 130 mT LEO payload performance comes from. 130 mT is not the maximum amount of mass the Ares V can insert on a lunar mission, it is the maximum amount of payload it can deliver to LEO in a single launch (payload being exclusive of the EDS).

[Bob B.]

The SI unit for the metric ton is a lowercase "t". Of course, over here, when someone says "ton", you can safely assume they are refering to the unit for one megagram, but I see how this could be confusing if there is more than one unit named "ton" in use.

In America when someone says "ton" they generally mean 2,000 pounds. Whenever I have a dialogue with an international group, like this forum, I try to specify "metric ton" or "tonne" so there's less chance for confusion. If I just say "ton" an American reader may misinterpret.

Another interest of mine is warships. Here the word "ton", when referring to ship displacements, generally means "long tons" or 2,240 pounds (1,016 kg). Ton is one of those words that just seems to have too many meanings.

[gaetanomarano]

This is where the 130 mT LEO payload performance comes from. 130 mT is not the maximum amount of mass the Ares V can insert on a lunar mission, it is the maximum amount of payload it can deliver to LEO in a single launch (payload being exclusive of the EDS).

your evaluation appears logic and matches the ESAS data you quote, but, comparing your results with the (real) Apollo missions' data, the EDS propellants' mass for TLI seems too much
according to astronautix the max payload mass (to a 185 km Orbit. 28.00 degrees) of the SaturnV was 118 mT and the total mass injected to TLI 47 mT (CSM+LM+2 mT) + 13.3 mT (3rd stage dry mass) = 60.3 mT leaving only (118-47-13.3=) 57.7 mT of propellants for TLI burn
.

[JMV]

your evaluation appears logic and matches the ESAS data you quote, but, comparing your results with the (real) Apollo missions' data, the EDS propellants' mass for TLI seems too much
according to astronautix the max payload mass (to a 185 km Orbit. 28.00 degrees) of the SaturnV was 118 mT and the total mass injected to TLI 47 mT (CSM+LM+2 mT) + 13.3 mT (3rd stage dry mass) = 60.3 mT leaving only (118-47-13.3=) 57.7 mT of propellants for TLI burn
.

I don't think the S-IVB dry mass is part of the payload as it's part of the launch vehicle. According to Apollo By The Numbers from 70 to 74 tonnes of propelant was used for TLI.

[Bob B.]

your evaluation appears logic and matches the ESAS data you quote, but, comparing your results with the (real) Apollo missions' data, the EDS propellants' mass for TLI seems too much
according to astronautix the max payload mass (to a 185 km Orbit. 28.00 degrees) of the SaturnV was 118 mT and the total mass injected to TLI 47 mT (CSM+LM+2 mT) + 13.3 mT (3rd stage dry mass) = 60.3 mT leaving only (118-47-13.3=) 57.7 mT of propellants for TLI burn.

According to Apollo by the Numbers, the TLI propellant mass for each of the Apollo missions was:

S-IVB 2nd Burn
Apollo 8: 149,510 lbm (67,817 kg)
Apollo 10: 155,136 lbm (70,369 kg)
Apollo 11: 156,679 lbm (71,068 kg)
Apollo 12: 158,653 lbm (71,964 kg)
Apollo 13: 156,097 lbm (70,804 kg)
Apollo 14: 158,495 lbm (71,892 kg)
Apollo 15: 163,469 lbm (74,148 kg)
Apollo 16: 162,441 lbm (73,682 kg)
Apollo 17: 163,498 lbm (74,161 kg)

If you add up the masses of all the individual components inserted into orbit along with the propellant remaining after ascent, you'll find that the Saturn V inserted 128 mT for Apollo 8, 134 to 137 mT for Apollo 10-14, and 140 to 141 mT for Apollo 15-17.

Without investigating further, it looks like the quoted LEO performance of 118 mT for the Saturn V must be a similar situation to the 130 mT performance of the Ares V, that is, it is payload to LEO exclusive of the third stage.

EDITED TO ADD:

Given the improved numbers, the propellant fraction is almost identical for Apollo and ESAS.

Apollo: 74 mT propellant / 141 mT total = 0.525

ESAS: 99.5 mT propellant / 188 mT total = 0.529

[gaetanomarano]

I don't think the S-IVB dry mass is part of the payload as it's part of the launch vehicle.

with the Apollo and ESAS 2nd stage that doubles as EDS we don't konw if we can call it "payload" or only part of the injiected mass
if it's not included in the "payload" definition, we must add 13.3 mT to the propellants' mass used for TLI that grows to 71 mT

[gaetanomarano]

Apollo: 74 mT propellant / 141 mT total = 0.525

the mistake comes from adding the EDS dry mass in the payload count, then, recalculating, your previous evaluation of the bigOrion+miniEDS mass was correct (74 mT in my evaluation, since the miniEDS has the same engine of the bigEDS)
then:
LOI+TEI bigOrion mass: 26 mT
mini EDS + J-2x dry mass: 8 mT
TLI propellants' mass: 40 mT
full miniEDS mass: 48 mT
full bigOrion+miniEDS mass: 74 mT
however, the full mass is not "over three times the Ares-I" but only a +180% more than its (26 mT) max lifted upperstage mass
.

[Bob B.]

with the Apollo and ESAS 2nd stage that doubles as EDS we don't konw if we can call it "payload" or only part of the injiected mass
if it's not included in the "payload" definition, we must add 13.3 mT to the propellants' mass used for TLI that grows to 71 mT

118 mT is definitely not the total inserted mass on an Apollo mission. The most mass ever inserted into orbit was Apollo 15, broken down as follows:

Ground Ignition Weights
S-IVB stage, total: 266,315 lbm (includes 239,462 lbm propellant)
Instrument unit: 4,487 lbm
Spacecraft/LM adapter: 3,964 lbm
Lunar Module: 36,238 lbm
Command/Service Module: 66,925 lbm
TOTAL: 377,929 lbm

Source: http://history.nasa.gov/SP-4029/Apol...on_Weights.htm

S-IVB First Burn
Propellant at start: 239,462 lbm
Propellant at end: 172,709 lbm

Source: http://history.nasa.gov/SP-4029/Apol...ellant_Use.htm

Therefore the total mass at the end of the first S-IVB burn, i.e. at orbit insertion, was

Inserted mass = 377,929 – 239,462 + 172,709 = 311,176 lbm (141,147 kg)

[gaetanomarano]

.

the bigOrion + Ares-II design allows the LOI+TEI of the bigOrion alone with mid-sized tanks but also the 35-37 mT superbigOrion with large tanks and LOI propellants for (both) LSAM and bigOrion, since, in a joined LSAM/bigOrion/bigEDS mission, the miniEDS must lift the bigOrion only to earth orbit and not to TLI

also, since the max payload to earth orbit of the Ares-II (without TLI) is higher than (only) the SM extra propellants, the Ares-II can lifth (also) a (bigger) LSAM ascent stage, while, the AresV can lift a bigger LSAM descent stage and a bigger EDS with more propellants

this way a joined superbigOrion+bigLSAM+superbigEDS can inject a bigger mass to TLI ...maybe, around +20 mT to have a 65 mT LSAM (with propellants for LOI) and add 5 mT (or more) extra life support and lunar exploration hardware to have three-four weeks missions for the price of 1.2-1.5 (standard ESAS) missions

this option is (also) a good point for the "Reinventing the ESAS" thread...
.

[gaetanomarano]

.

- larger tanks for better missions at (near) zero extra R&D and manufacturing costs

- a vehicle able to accomplish cargo, cargo-return, crew and rescue missions alone

- a very modest increase of the SM dry mass that allows the launch with (both) a standard Ares-I or an Ares-II

- the option for better and (3-4 times) longer moon missions with the same 1.5 launch architecture and bigger vehicles

....

it's INCREDIBLE that NASA and LockMart have NOT implemented this simple, easy, powerful and low cost option in their Orion design!

and is (also) incredible that my porposal has so many critics and hostility against it (reflected in the pro-corkscrew poll result!)
.

[JMV]

Inserted mass = 377,929 – 239,462 + 172,709 = 311,176 lbm (141,147 kg)

Perhaps not exactly at that accuracy. Two ullage motors used at staging were jettisoned after use. Probably included in the
S-IVB stage, other: 1,655 lbm. I don't know about boil-off venting or attitude control propellant usage in the APS modules during the first burn. These corrections are pretty insignificant to your point though. Just nitpicking, carry on.

[Bob B.]

Perhaps not exactly at that accuracy. Two ullage motors used at staging were jettisoned after use. Probably included in the
S-IVB stage, other: 1,655 lbm. I don't know about boil-off venting or attitude control propellant usage in the APS modules during the first burn. These corrections are pretty insignificant to your point though. Just nitpicking, carry on.

Yeah... I considered mentioning some of this stuff but decided to remain silent. I didn't think it significant enough to over complicate the point I was trying to make. Nonetheless, the things you mention are certainly true and do reduce the inserted mass a little bit.

[Bob B.]

your previous evaluation of the bigOrion+miniEDS mass was correct (74 mT in my evaluation, since the miniEDS has the same engine of the bigEDS)
then:
LOI+TEI bigOrion mass: 26 mT
mini EDS + J-2x dry mass: 8 mT
TLI propellants' mass: 40 mT
full miniEDS mass: 48 mT
full bigOrion+miniEDS mass: 74 mT
however, the full mass is not "over three times the Ares-I" but only a +180% more than its (26 mT) max lifted upperstage mass.

Those numbers are fairly close to what I calculated. Assuming enough EDS propellant is burned during launch to lower the propellant mass fraction to 0.80, then

Orion CM & SM, with 2400 m/s delta-v: 26 mT
mini EDS dry mass: 9.1 mT
TLI propellant mass: 36.4 mT
Total mass @ TLI: 71.5 mT

Of course there is no way of knowing what the propellant mass fraction will actually be without a preliminary design of the entire launch system. The 0.80 number is based on the mass budgets for the large EDS as shown in the ESAS report.

By the way, according a September interview with Marshall Space Flight Center's Danny Davis, manager of the Ares I Upper Stage Element Office, the target performance for Ares I is about 25 mT injected mass.

“Our current requirement is to deliver the Orion to a 30x100 nmi injection orbit. The performance capability is approximately 55klbm.” -- Danny Davis

[gaetanomarano]

.

from post #154 my evaluation for the Orion mass was:

standard Orion SM mass: 3-3.5 mT
standard Orion propellant mass: 7.2-7.7 mT
standard Orion full mass: 19.2 mT

while the real figures from a recent NASA document are:

standard Orion SM mass: 3.85 mT
standard Orion propellant mass: 8.15 mT

the total Orion+SM+fuel mass is 21.5 mT ...but NASA has increased the Orion CM mass from 8.5 mT to 9.5 mT

.

[Bob B.]

Interesting, it looks like they must have decreased in delta-v requirement. That obviously voids all the calculations that were based on previously published data.

[gaetanomarano]

...have decreased in delta-v requirement...

the Jan. 2007 document don't have any change about this point vs. the Aug. 2006 Orion specs (probably) since that change was already done over a year ago when NASA resized the Orion dimensions and specs from early ESAS plan
.

[Bob B.]

the Jan. 2007 document don't have any change about this point vs. the Aug. 2006 Orion specs

They don't have to say it to know a change must have been made because it is implicit from the other data.

[cjl]

Exactly. This is precisely the same point that we spent months arguing with you on about the 5 segment SRB's. Just because data is not published does NOT mean that it cannot be inferred from other published data. Because we know certain rules that the vehicles must follow, we know certain things that MUST be true in order for the published data to be correct. One of these things is the delta v, which must be different than previously published to achieve the numbers that are now being released.

[gaetanomarano]

Exactly.Just because data is not published does NOT mean that it cannot be inferred from other published data.

no, not exactly
the 5-seg.SRB and Ares-I specs are not "data" (of real objects that perform real flights) but (so far) only "design goals" that need (hard) work to be achieved (and that may happen, or not, or only in part)
.

[cjl]

However, they are not going to publish design goals that go against known laws of physics. You will never see a design goal that says "nullify gravity and float to the moon." The goals are realistic, at least as far as being physically possible.

[gaetanomarano]

...they are not going to publish design goals that go against known laws of physics...

surely not, but (you know) often some (too optimistic) design goals may fail (full or part) their targets when faces with real materials, real technologies, real tests, etc.
.

[cjl]

Yes, but it is physically impossible for the delta v to be the same with those ones that you posted, regardless of the possibility of failure. So, if you agree that the known laws of physics would not be violated by the design goals, you must agree that the delta v has changed. There are no other options.

[gaetanomarano]

There are no other options.

not true, actually there is a design option that may change the Orion specs without violate any physics' law: the SM's "spare propellent" ...the Apollo SM was fueled with LOI, TEI, maneuverings and 2 mT (IIRC) of "spare" propellent (for redundancy) while the Orion will have only the TEI, maneuvering and "spare" propellant ...changing that value the Orion's GLOW may vary ...also, if NASA is (quietly) shifting (as I believe) to an ArianeX-like M.L.A. (without any earth orbit Orion-LSAM docking) the SM may save fuel vs. the original ESAS plan
.

[gaetanomarano]

.

a possible solution of the Orion's "overweight" is a Truncated Biconic Shape Orion ...it is less complex than a "BigelowOrion" and less "exotic" than an "EggCEV" but may be reduce the Orion weight (if necessary)

.

[Bob B.]

a possible solution of the Orion's "overweight" is a Truncated Biconic Shape Orion it is less complex than a "BigelowOrion" and less "exotic" than an "EggCEV" but may be reduce the Orion weight (if necessary)

Achieving the required L/D ratio has proven to be very problematic with such a shape.

[Nicolas]

Some problems that might arise:

*in ascent, the cp is located higher, which is less stable
*the heat shield area is smaller; can it stand up to the re-entry? (same goes for L/D as Bob B points out)
*apparently you didn't take the increased size of the sides (due to the cornered double surface) into account
*your cg will be higher; this likely will make the craft less stable during descend
*I wouldn't be surprised if your ascent aerodynamic drag would increase
*your capsule is more prone to collapsing with this shape than a cone
*is the heated flow detached enough from the lower sides of the capsule? (I'd need more data to have an idea on what's optimal there, as the flow closest to the skin tends to get cooler again at equilibrium because it becomes more or less stationary and hence doesn't generate that much friction anymore. The kinetic-static conversion heat still is there though, as is expansion cooling and other effects).

All these things, and others, need to be taken into acount before concluding it is a better, even feasible option. Can you also present more detailed calculations on your weight saving figures?

[Bob B.]

So, if you agree that the known laws of physics would not be violated by the design goals, you must agree that the delta v has changed. There are no other options.

There is one other option: a change in propellant. Using the numbers provided in the linked article...

CM mass: 21,000 lbm
SM mass: 8,500 lbm
SM propellant: 17,956 lbm

...we can conclude a couple things. First, using hypergolic propellants the spacecraft delta-v will be around 1,500 m/s. Second, the delta-v figure of 1,855 m/s from Aug-06 can be achieved with a specific impulse of around 400 s, which is too high for hypergolics.