totm dec 2023 Artemis Discussion Thread

[jinnantonix]

13 hours ago, sevenperforce said:

Right. Plus, Earth orbit assembly allows a single TLI burn, a single lunar powered flyby, and a single LOP-G rendezvous. So it's more mass-efficient by definition, because you have less dry mass going to LOP-G.

If you build the tanks larger, then that's more dry mass wasted at LOP-G, and conceivably more dry mass being hauled down to LLO and even further.

 

When you mention building use the Canadarm, are you assuming using the existing facilities at ISS?

The problem I see with this is the size of the second stage of the LV that is needed to launch the fully assembled lunar lander from LEO to TLI.  There will also need to be a 800m/s burn from TLI to LOP-G rendezvous, and assume that there will be small prop drop tanks included so these  can be jettisoned prior to LOP-G to LLO transit?

Quote

A downthrottled vacuum-expanded Superdraco can get no lower than 12 tonnes lunar thrust, so it is just too thrusty...

BlueMoon+LockMart ascent vehicle a dry mass of 2.5 tonnes, but I'll be a touch more generous and say 2.8 tonnes (assuming 319 s Isp, 2.6 km/s from lunar surface to LOP-G, 6.5 tonnes on the surface). Those 3.7 tonnes of props will require a tankage dry mass of 630 kg and of course 100 kg of AJ-10 engine, leaving the crew vehicle plus control thrusters at a mass of about 2,070 kg.

These numbers seem very low.  At lunar landing, the DV will still have some fuel in the mostly empty tanks, plus lander legs and decouplers. I calculate approx 3.2 tonnes for that, we agree on that.  Then there is static payload (science equipment, rovers, etc) and 2 weeks of life support, what mass are we assuming here?  How about the airlock?  Is all this included in the 2.8 tonnes you mention above? 

My calculations are higher than yours.  Assuming crew taxi is 4.0t AV dry mass including payload, LS etc, adding AV props and tank and DV components, I am getting 15.1t wet mass immediately prior to landing on the surface.  One down-throttled SuperDRACO is fine for that.  What am I missing here?

Edited October 24, 2019 by jinnantonix

[tater]

Blue Origin has said that the "stretch" version of their lander for crew lands 6.5 tons of cargo on the surface (not counting their lander stage). So you have 6500kg to play with for a crew compartment, engine, and props to get from the surface to Gateway.

 

[sevenperforce]

13 hours ago, jinnantonix said:

When you mention building use the Canadarm, are you assuming using the existing facilities at ISS?

This is one of the bridges; we would need a Canadarm added to LOP-G. Though all it has to do is move a single replaceable tank from a mount on the ascent vehicle and attach it to the reusable crew taxi while both are docked. No major assembly.

13 hours ago, jinnantonix said:

The problem I see with this is the size of the second stage of the LV that is needed to launch the fully assembled lunar lander from LEO to TLI.  There will also need to be a 800m/s burn from TLI to LOP-G rendezvous, and assume that there will be small prop drop tanks included so these  can be jettisoned prior to LOP-G to LLO transit?

Well, to begin with, the cost to get from TLI to LOP-G is not 800 m/s; that's the round-trip cost for Orion. If you are making a one-way trip from TLI to LOP-G, you only need a 183 m/s powered lunar flyby and a 215 m/s NRHO insertion burn, for a total of ~400 m/s. See page 5 of this paper. I was allowing 430 m/s as this is the value cited in some other papers.

But in any event, there are multiple ways to do it:

• Launch the landing/ascent module (LAM) as a monolith to TLI with drop tanks so it can use its AJ-10 to do the powered flyby and NRHO insertion. In this configuration, the LAM needs two drop tanks, one for the trip to LOP-G and one for the trip to the lunar surface, and masses around 20 tonnes. You would need distributed launch, with the monolith going to LEO on one commercial launch vehicle and mating to a separately-launched docking ring on a Centaur or FHUS that would push it to TLI.
• You can also launch the LAM with only one drop tank (for the trip from LOP-G to the lunar surface) and keep it attached to the TLI stage so that the TLI stage provides the 400 m/s required for the trip to NRHO. With this config, the monolithic LAM masses 17-18 tonnes so you would still need distributed launch, but you could probably do it a lot cheaper because the TLI stage will be more efficient.
• Finally, you can launch the LAM with no drop tank at all and use the TLI stage (maybe ACES?) for NRHO insertion AND transfer to LLO and finally to the lunar surface, but that requires a lot of restarts. This is by far the most efficient, though. The LAM masses under 7.6 tonnes.

13 hours ago, jinnantonix said:

These numbers seem very low.  At lunar landing, the DV will still have some fuel in the mostly empty tanks, plus lander legs and decouplers. I calculate approx 3.2 tonnes for that, we agree on that.  Then there is static payload (science equipment, rovers, etc) and 2 weeks of life support, what mass are we assuming here?  How about the airlock?  Is all this included in the 2.8 tonnes you mention above? 

My calculations are higher than yours.  Assuming crew taxi is 4.0t AV dry mass including payload, LS etc, adding AV props and tank and DV components, I am getting 15.1t wet mass immediately prior to landing on the surface.  One down-throttled SuperDRACO is fine for that.  What am I missing here?

Two things. First, your AV mass seems high. As @tater points out, the stretched Blue Moon lander can put no more than 6.5 tonnes on the lunar surface, so that's the upper limit on LockMart's reusable AV wet mass. With hypergols, it needs to burn 56% of its weight to go from the lunar surface to LOP-G. Factoring in dry mass for the AJ-10 and tankage, the LockMart AV has only 2,070 kg for capsule, LS, crew, return cargo, airlock, and everything else you mentioned. Not a lot, but keep in mind that the dry mass of the Apollo LM AV was only 2,050 kg including tanks, so this is reasonable.

Additionally, I'm dispensing with the descent vehicle altogether. Wet mass on the surface is under 7 tonnes and is merely the reusable crew taxi, the ascent motor+tanks, and landing legs/ladders that are jettisoned at takeoff. The descent phase is either performed by the transfer vehicle as a crasher stage, or it is performed using drop tanks on the ascent vehicle. It makes no sense to have a large landing structure to support dry mass you no longer need.

EDIT: To @tater's point, here's the quote:

Quote

At an invitation-only media event in Washington DC on Thursday, Blue Origin founder and CEO Jeff Bezos showed off a full-scale mock-up of the Blue Moon lander, which he says is capable of delivering 3.6 tonnes of cargo to the lunar surface and will be the basis for a larger manned version capable of carrying 6.5 tonnes.

From here.

[sevenperforce]

By my numbers, the mass of the cabin and associated structure for the Apollo LM AV (no prop tanks or engine) was 1,600 kg, plus crew and cargo. Mass growth of only 30% (for the LockMart reusable AV concept) seems pretty minimal.

[Barzon]

Boeing lander appears to be methalox 

"Houston-based Intuitive Machines was selected to build, test and deliver prototype main stage and reaction control (RCS) engines for Boeing’s Human Lander System (HLS)" 
https://www.intuitivemachines.com/post/intuitive-machines-selected-to-build-engines-for-boeing-s-human-lander-system-technology-development

 

[55delta]

1 hour ago, sevenperforce said:

we would need a Canadarm added to LOP-G.

I'd heard a while back that the Canadian Space Agency has already committed to a 'Canadarm III' for LOP-G.

[jadebenn]

3 minutes ago, 55delta said:

I'd heard a while back that the Canadian Space Agency has already committed to a 'Canadarm III' for LOP-G.

They have. They were the first country to sign-on. Japan was the second. We're still waiting on the ESA and Roscosmos, but both are almost guaranteed to participate.

[sevenperforce]

Just now, jadebenn said:

4 minutes ago, 55delta said:

I'd heard a while back that the Canadian Space Agency has already committed to a 'Canadarm III' for LOP-G.

They have. They were the first country to sign-on. Japan was the second. We're still waiting on the ESA and Roscosmos, but both are almost guaranteed to participate.

Yes, a WHILE back, but NASA's most recent request for submissions for a HLS stated that a Canadarm would not be installed at LOP-G initially.

[jadebenn]

6 minutes ago, sevenperforce said:

Yes, a WHILE back, but NASA's most recent request for submissions for a HLS stated that a Canadarm would not be installed at LOP-G initially.

Only because it wouldn't be ready in-time. They're definitely providing one, just not by 2024.

[sevenperforce]

One final configuration that may prove to be the most efficient of all: you can build the LAM with a single, smaller drop tank, enough to give it around 1400 m/s of dV...around 4.3 tonnes of propellant. Wet mass to LEO is just under ten tonnes. You then pair it in LEO with a Centaur or FHUS, which pushes it to TLI and performs the powered lunar flyby and NRHO insertion. The manned crew taxi is fueled and mated, then that same high-performance upper stage performs the transfer to LLO and initiates the first 150 m/s of descent before breaking free.

This is probably the most mass-efficient approach, though I would have to run the numbers. It does require five restarts of the upper-stage engine. However, it's also extremely safe. If for some reason the LAM engine fails after the initiation of descent or the upper stage fails to separate from the LAM, the crew taxi has ample dV and plenty of time to separate, cancel descent, and return to LOP-G.

[mikegarrison]

 

"Artemis: because the governor needs a new jet."

Edited October 24, 2019 by mikegarrison

[jinnantonix]

8 hours ago, sevenperforce said:

Well, to begin with, the cost to get from TLI to LOP-G is not 800 m/s; that's the round-trip cost for Orion. If you are making a one-way trip from TLI to LOP-G, you only need a 183 m/s powered lunar flyby and a 215 m/s NRHO insertion burn, for a total of ~400 m/s. See page 5 of this paper. I was allowing 430 m/s as this is the value cited in some other papers.

Thanks for the link.  When I did the station building simulation, I did the rendezvous many times, and without the use of a supercomputer to accurately calculate the optimum path, I was consuming about 800m/s.

Quote

First, your AV mass seems high. As @tater points out, the stretched Blue Moon lander can put no more than 6.5 tonnes on the lunar surface

This is more like the mass of the AV I used in my simulation.  I was thinking this was wrong (too light), but now you have confirmed that my original calcs were accurate.  This definitely rules out using an Orion pressure vessel (too massive), but I reckon a smaller, lighter vessel could be constructed using the same proven methods, and use Orion's avionics and LS facilities.

One issue I am facing with the ascent vehicle:  at 6.5t wet mass, a single AJ10-180 at 11kN thrust does not get the AV off the lunar surface efficiently. I prefer the AJ10 Space Shuttle engine for it's proven reliability and throttling capability.  My simulation utilised two AJ10-190 engines mounted each side of the AV capsule.  This works from a TWR viewpoint, but would mean disaster if one of the engines failed during ascent.  

A single central SuperDRACO is fine for launching the AV, but as you say, cannot be throttled down to hover when landing the AV alone (after crashing the DV stage).  The single SuperDRACO throttled down works fine when landing the DV and AV together, and the latter configuration, while less fuel efficient during landing, allows leaving behind the landing legs at AV launch. 

In order to transit from LLO to lunar surface three engines are needed ~100kN thrust.  The two additional engines could be attached to the DV and left on the lunar surface.  Alternatively three inline SuperDRACOs on the AV provides redundancy during the ascent phase.

 

 

Edited October 25, 2019 by jinnantonix

[sevenperforce]

1 hour ago, jinnantonix said:

Thanks for the link.  When I did the station building simulation, I did the rendezvous many times, and without the use of a supercomputer to accurately calculate the optimum path, I was consuming about 800m/s.

Were you using a single insertion burn or were you doing the powered flyby? I think that doing a single insertion burn at NRHO is more along the lines of 600-700 m/s.

1 hour ago, jinnantonix said:

This is more like the mass of the AV I used in my simulation.  I was thinking this was wrong (too light), but now you have confirmed that my original calcs were accurate.  This definitely rules out using an Orion pressure vessel (too massive), but I reckon a smaller, lighter vessel could be constructed using the same proven methods, and use Orion's avionics and LS facilities.

It's worth noting that the original NASA HLS RFP we were reviewing earlier definitely had a higher quoted mass for the ascent vehicle:

Spoiler

9-12 tonnes is significantly larger than the 6.5 tonnes that the stretched Blue Moon can land, so that's odd. Using a 9-12 tonne AV rather than the 6.5 presumed for the LockMart reusable AV, we'd be seeing a crew capsule that's 2.9-3.8 tonnes apart from engines and tankage. That's much better margin, but not SO much more that the 6.5-tonne version is utterly implausible.

1 hour ago, jinnantonix said:

One issue I am facing with the ascent vehicle:  at 6.5t wet mass, a single AJ10-180 at 11kN thrust does not get the AV off the lunar surface efficiently. I prefer the AJ10 Space Shuttle engine for it's proven reliability and throttling capability.  My simulation utilised two AJ10-190 engines mounted each side of the AV capsule.  This works from a TWR viewpoint, but would mean disaster if one of the engines failed during ascent.  

I certainly wouldn't be choosing an 11 kN engine. I was baselining with the AJ-118K, at 44 kN max thrust, hover at 25% throttle, 100 kg mass. A Shuttle OME would also do the trick, though not as spry at liftoff.

1 hour ago, jinnantonix said:

In order to transit from LLO to lunar surface three engines are needed ~100kN thrust.  The two additional engines could be attached to the DV and left on the lunar surface.  Alternatively three inline SuperDRACOs on the AV provides redundancy during the ascent phase.

Why do you need multiple gees of thrust during descent?

[sevenperforce]

There's a high sensitivity to things like mass ratio, too. Transtage had a 16.995% dry mass fraction (not counting the engine) while the Delta-K and Delta-P had a 12.4% dry mass fraction. So that makes a significant difference.

[jinnantonix]

9 minutes ago, sevenperforce said:

Were you using a single insertion burn or were you doing the powered flyby? I think that doing a single insertion burn at NRHO is more along the lines of 600-700 m/s.

Yes, that's what I was doing.  Just couldn't eyeball the angles right with the powered deceleration.

 

Quote

I certainly wouldn't be choosing an 11 kN engine. I was baselining with the AJ-118K, at 44 kN max thrust, hover at 25% throttle, 100 kg mass. A Shuttle OME would also do the trick, though not as spry at liftoff.

I thought of using the AJ10-118, but that's a beast of an engine, weight 450kg with a huge engine bell.  The OMS engine at 29kN thrust would be a good choice and should have just enough thrust at lift-off.  Still need to add two engines for the LLO to surface burn though.  

[jinnantonix]

Simulation of the 6.5t AV completed

Spoiler

Dropping the transit tanks

Hovering

Stable on slope

Lift off

Thrusters only from LLO to NRHO

 

My model assumes a smaller lighter Orion-type pressure vessel,  the AV has a wet mass of 6.8t, slightly more than the BO assumption to accommodate three engines at AV lift off.  The central engine is throttled down for hovering/landing, but assuming it can be throttled up to full in case of an outer engine failure.  I still like the 3 inline design for redundancy for ascent vehicle lift off.  Since 3 engines are required for descent anyway, it is a trade-off between mass and redundancy (I choose safety first).  A single Shuttle OMS or a AJ10-190 as a central engine with two additional outer engines (could be SuperDRACOs or AJ10-190s).

The re-usable AV (wet mass 2.7t) , with a transit attachment, a drop tank and single engine or multiple thrusters for insertion into NRHO and rendezvous with the LOP-G .  I calculate the extra props, drop tank and engine adds 1.0t to the craft.  This may be delivered to the LOP-G initially on a Falcon 9 or Vulcan.

The DV + transit drop tanks stage has a wet mass of 27.5t at Earth launch, plus 5.0t additional drop tank for the deceleration and rendezvous at the LOP-G.  It is assumed that this is launched to LEO on a Falcon Heavy, thence TLI from any one of the options that @sevenperforce has suggested.  Note during NRHO insertion, the engine burns on the three inline engines can be interchanged between central and outer pair to reduce the total number of firings per engine.

The craft tests like a dream, in particular the wide landing base reduces the risks around lateral movement at touchdown and landing on a slope.

 

[sevenperforce]

9 hours ago, jinnantonix said:

Yes, that's what I was doing.  Just couldn't eyeball the angles right with the powered deceleration.

That's Oberth for you. The moon might not have much gravity but it definitely gives you an advantage.

Quote

I thought of using the AJ10-118, but that's a beast of an engine, weight 450kg with a huge engine bell.  The OMS engine at 29kN thrust would be a good choice and should have just enough thrust at lift-off. 

Ah, I see my mistake. I was following Wikipedia, which listed the heaviest AJ10 variant at 100 kg. Just took a look at astronautix and I see there is a huge array of different mass variants. Gonna have to run some new numbers to account for that.

Since we are talking about liftoff thrust, it would be worth factoring in some of the information from here about ascent profiles. For example, the Apollo LM AV had a ten-second vertical burn and then a total pitchover. That corresponds to 16.2 m/s of gravity drag losses; if you have more thrust, you can reduce that.

 

Edited October 25, 2019 by sevenperforce

[sevenperforce]

9 hours ago, jinnantonix said:

Still need to add two engines for the LLO to surface burn though.  

Why? Gravity drag on descent is fairly low if you come in at a sharp angle.

After a lot of iteration, I'm fairly convinced that the ideal configuration is going to have the LAM launched to LEO with a tank for the crew taxi and a single drop tank to handle descent only. It loiters in LEO while SLS sends Orion to LOP-G, then a Centaur and docking ring is launched as soon as possible thereafter to push the LAM to TLI and perform the rendezvous. The vehicle stacks at LOP-G and Centaur performs the transfer to LLO and initiation of descent; the single engine on the LAM uses its drop tank props to complete all but the last 300 m/s of descent with a better T/W ratio.

9 hours ago, jinnantonix said:

Since 3 engines are required for descent anyway, it is a trade-off between mass and redundancy (I choose safety first).  A single Shuttle OMS or a AJ10-190 as a central engine with two additional outer engines (could be SuperDRACOs or AJ10-190s).

The primary reason to go with a pressure-fed engine in the first place is safety first so that you never need engine-out capability. If you are going for multiple engines for engine-out, then you should go with pump-fed engines and balloon tanks, which saves way more dry mass and cranks up specific impulse.

[sevenperforce]

 I went ahead and threw together some references numbers for what kind of residuals can be expected if someone wanted to use a "naked" HLV upper stage to perform TLI or any number of BLEO ops. To calculate, I've taken published or estimated GTO performance figures for various vehicle configurations and done the math in reverse to calculate staging velocity, then "removed" the payload and calculated forward to get to LEO. As such, these estimates contain some degree of conservatism; the absence of a payload during primary ascent will improve performance marginally. Buoyed by this, I've also ignored the necessary addition of a docking ring: this is likely a safe assumption given that the upper stage will need no payload adapter or decoupling mechanism.

By the time Artemis is realized, Atlas V will likely no longer be flying, but I included it for reference anyway.

Assumptions were 2270 m/s from LEO to GTO, 3200 m/s from LEO to TLI, 430 m/s from TLI to NRHO. I estimated the propellant load of Centaur V by comparing the volume of Centaur V to the volume of the Atlas V Centaur; I estimated the dry mass of Centaur V by subtracting the mass of the RL-10 from the dry mass of the Atlas V Centaur, increasing mass in proportion to fractional surface area, and adding two RL-10s back in.

Edited October 25, 2019 by sevenperforce

[Barzon]

Boeing Lander

[tater]

12 minutes ago, Barzon said:

Boeing Lander

They have been talking about this one for years,

https://arc.aiaa.org/doi/abs/10.2514/6.2017-5147

[ThatGuyWithALongUsername]

4 minutes ago, tater said:

They have been talking about this one for years,

https://arc.aiaa.org/doi/abs/10.2514/6.2017-5147

Looks like it got a bit taller, interestingly. Very unique design- interesting open area in the descent module.

[Barzon]

Damn you @tater!

[tater]

As soon as I saw that lander the other day I remembered I had read about it, then posted images in the SSTU thread, so it was easy to find my old links.

LOL, I actually posted that ^^^ October 2017. Think it was presented at the same AIAA meeting, and I went through their PDFs.

[sevenperforce]

2 hours ago, Barzon said:

Boeing Lander

Upon review, that looks like four AJ-10s on the ascent stage and 6-7 on the descent stage.

If you can refuel the ascent stage, great, but you need to master microgravity propellant transfer. And pressurant transfer, if you're using pressure-fed engines, which is challenging at best.

I maintain that using thrusty engines for the ascent stage is a waste because a quarter of your dV requirement is not thrust-dependent.