## Block 1 capabilities

The Block 1 payload to Low Earth Orbit is usually described as 70 metric tons. However, if you have paid close attention to the project and have an eye for rocketry, you might have noticed that this isn't true. Initially, SLS was to have four different variants: Block 0, Block 1, Block 2 and Block 3. Block 0 was going to have 4 segment boosters, an External Tank length core and three RS-25 engines. This version would lift 70 tons to LEO. The Block 1 described here would have a stretched core, 5 RS-25 engines and 5 segment boosters and would be able to lift 100 metric tons to LEO. (More info on these variants, as well as other SLS alternatives, can be found here, on spacelaunchreport.com). When NASA later started to refine the project, they found that Block 0 would only complicate development since the core would need to be completely redesigned for Block 1 and later, so they decided to take it out. Block 1, however, remained mostly unchanged, except that one of the engines was taken out because of being unnecessary. The payload decreased slightly due to this, from 100 tons to an estimated 90-95 tons. However, due to a variety of political reasons that I don't want to get involved with, NASA still has to claim 70 tons for the initial variant. If you don't believe me, this NASA document also estimates Block 1 capacity at 90 metric tons. While it's probably slightly more than that, I'll stay at 90 metric tons.

The Boeing proposed SLS third Stage |

ce the basic 25 metric tons is not enough. For this role, I think the scaled up DCSS proposed by Boeing in their lunar exploration architecture is sufficient. It has a wet mass of 45.6 metric tons and an empty mass of 4.7 metric tons, with a specific impulse of 462 seconds. If we assume the delta V required for TLI at 3150 m/s I get 36.13 metric tons to TLI. And this is conservative, since SLS would be able to launch the stack into a somewhat higher orbit because of it's higher lift capacity, and the TLI delta V is usually slightly lower; 3100m/s gave me 37 tons, for example. So let's assume Block 1+ASCS (Advanced SLS Cryogenic Stage, which is how I shall call it from now on) TLI capacity at 37 metric tons.

## The mission architecture

The most important part of the mission is, of course, the mission architecture, which is how you will actually execute the mission. The main goal for this mission is to get people on the surface of the moon in a single launch on the basic Block 1 SLS. The primary components will be the crew vehicle and the lunar lander. For the crew vehicle, we will of course assume Orion. The lander will be of my own design. First thing the ship will need to do is enter orbit around the moon. For this task, the service module of Orion should be sufficient. Instead of entering a low lunar orbit, like Apollo did, the vehicle will enter a highly elliptical orbit to save delta V during both orbit injection and return to Earth. This will require a bigger lander, of course, but I will adjust for that later. Lander mass is 15.1 tons, as I will describe later, and I'll assume a 500 m/s injection burn and return burn.

For Orion, finding hard numbers is kind of difficult. Wikipedia has them but doesn't give a source, NASA has a fact sheet on their site but those numbers don't really match up with a realistic specific impulse for the engine. It claims a 22.78 ton full mass on orbit, 1500 m/s delta V, and 7.91 tons of propellant. If all of these values were correct the Orion main engine would need an isp of 358.5 seconds, which is impossible with current hypergolic propellants. A specific impulse of 320 seconds if far more realistic. 1500m/s for Orion is a requirement, so that means either propellant mass or total mass is incorrect. We will assume that total mass is incorrect here, though increasing propellant mass relative to the total should result in similar capabilities in the end. More recent sources indicate a total mass of 21.250 kilograms of 21.25 tons. This gives a specific impulse of 327.5, which is also what NASA claims in a SLS ConOps documents (which I can't cite here due to paywall). Still a little optimistic but certainly possible in the range of hypergolic propellants. Now, assuming a specific impulse of 327.5, a propellant mass of 7.91 tons and total mass of 21.25 tons, we can start doing some math. Total mass of Orion+lander will be 36.35 tons. Using the rocket equation, we get the following:

MassRatio=e^500/(327.5*9.81)= 1.1683923.

If we divide the total system's mass by this we get 31.11 tons, meaning Orion's SM has used a total of 5.24 tons of propellant. This leaves us with Orion having a mass in lunar orbit of 16.01 tons. With an empty mass of 13.34 tons, we get the following delta V left in Orion:

delta V= 327.5*9.81*ln(16.01/13.34) = 586.16 m/s!

This means that even if Orion has to do the LOI burn for both itself and the lander into and elliptical orbit, it still has enough delta V left to return, including a margin of over 86 m/s! That is definitely enough to enter an elliptical lunar orbit and be able to return afterward.

Next, the lunar lander. Like Apollo, it is composed of two main parts, and ascent module and a descent module. In order to increase the delta V of the system but keep the mass close to 15 tons, I used a staged approach for the Ascent Module and more efficient cryogenic fuels for the Descent Module.

For the Ascent module, we can't use efficient propellants like hydrogen. These are very volume inefficient, leading to high empty masses and poor mass ratios. Instead, I decided to use storable propellants and a dual stage approach to make up for the poor efficiency. The habitable module for a crew of two, including an engine and empty propellant tanks would have a mass of 2000 kg, and would be capable of carrying 1000 kg of propellant. The engine we used for this would be derived from Aerojet R-4D, which Boeing claims could have their Isp increased to +320 seconds easily.

The empty mass of 2000 kg is a 15% decrease from the Apollo LM, which is very much attainable due to three factors:

Lunar Lander Delta V overview:

For Orion, finding hard numbers is kind of difficult. Wikipedia has them but doesn't give a source, NASA has a fact sheet on their site but those numbers don't really match up with a realistic specific impulse for the engine. It claims a 22.78 ton full mass on orbit, 1500 m/s delta V, and 7.91 tons of propellant. If all of these values were correct the Orion main engine would need an isp of 358.5 seconds, which is impossible with current hypergolic propellants. A specific impulse of 320 seconds if far more realistic. 1500m/s for Orion is a requirement, so that means either propellant mass or total mass is incorrect. We will assume that total mass is incorrect here, though increasing propellant mass relative to the total should result in similar capabilities in the end. More recent sources indicate a total mass of 21.250 kilograms of 21.25 tons. This gives a specific impulse of 327.5, which is also what NASA claims in a SLS ConOps documents (which I can't cite here due to paywall). Still a little optimistic but certainly possible in the range of hypergolic propellants. Now, assuming a specific impulse of 327.5, a propellant mass of 7.91 tons and total mass of 21.25 tons, we can start doing some math. Total mass of Orion+lander will be 36.35 tons. Using the rocket equation, we get the following:

MassRatio=e^500/(327.5*9.81)= 1.1683923.

If we divide the total system's mass by this we get 31.11 tons, meaning Orion's SM has used a total of 5.24 tons of propellant. This leaves us with Orion having a mass in lunar orbit of 16.01 tons. With an empty mass of 13.34 tons, we get the following delta V left in Orion:

delta V= 327.5*9.81*ln(16.01/13.34) = 586.16 m/s!

This means that even if Orion has to do the LOI burn for both itself and the lander into and elliptical orbit, it still has enough delta V left to return, including a margin of over 86 m/s! That is definitely enough to enter an elliptical lunar orbit and be able to return afterward.

Next, the lunar lander. Like Apollo, it is composed of two main parts, and ascent module and a descent module. In order to increase the delta V of the system but keep the mass close to 15 tons, I used a staged approach for the Ascent Module and more efficient cryogenic fuels for the Descent Module.

For the Ascent module, we can't use efficient propellants like hydrogen. These are very volume inefficient, leading to high empty masses and poor mass ratios. Instead, I decided to use storable propellants and a dual stage approach to make up for the poor efficiency. The habitable module for a crew of two, including an engine and empty propellant tanks would have a mass of 2000 kg, and would be capable of carrying 1000 kg of propellant. The engine we used for this would be derived from Aerojet R-4D, which Boeing claims could have their Isp increased to +320 seconds easily.

The empty mass of 2000 kg is a 15% decrease from the Apollo LM, which is very much attainable due to three factors:

- The miniaturization of electronics means a computer weighing 100 kg during that time can fit in a pocket now. They also take up less space now, meaning that the cabin can be smaller but just as spacious for the crew.
- The stage has to hold less propellant, meaning that the tanks are smaller and therefore lighter
- Modern manufacturing technologies and lighter materials like plastics and carbon composites mean that even a direct design copy of the LM built using newer alloys and composite materials could be a lot lighter

So I'd say 15% is very attainable if not a little pessimistic. One of the Golden Spike designs for a lunar lander is basically a plastic ball, and it's pressurized and can hold two people for several hours with a mass of roughly 1500 kg, also including empty tanks and engines, so 2000 kg should be enough. It might even be capable to cram four people in there and give them a surface stay of a few hours, which is enough for visiting a base on the surface.

The STAR 48 booster |

The AM also includes a booster stage. For this, I assumed the STAR 48 solid rocket motor. This rocket motor might be solid fuel, which can sound scaring to some people, but it's actually a very sensible choice. On smaller systems, solid rockets are very reliable due to few mechanical parts, they have extremely good mass ratios and they are very long storable, unlike cryogenic fuels. The Star 48 is also a flown, proven and affordable system, which makes the lander simpler and safer. It has a fueled mass of 2146 kg and an empty mass of 117 kilograms, with a specific impulse of 291. They are not capable of thrust vectoring control though, so the reaction control thrusters on the AM will have to take care controlling the craft during ascent.

For the Descent Module, I again took the Apollo LM as an example. The Apollo DM had a wet mass of 10344 kilograms and an empty mass of 2144. Then I decreased the total mass to 10000 and increased the wet mass to 2600 kilograms to make up for the extra insulation of the cryogenic fuel, as well as a light fairing to cover up the Star 48 during landing. This gives a mass ratio of 3.85, which is plausible for a hydrogen powered stage. The engine used is the RL-10 with a specific impulse of 462 seconds.

Ascent Module, Habitat Module = 1273 m/s

Ascent Module, Booster = 1431 m/s

Ascent Module, Total = 2704 m/s

Descent Module Total = 3039 m/s

Lunar Lander Total = 5743 m/s

Lunar Lander Total Mass = 15146 kg

The published figures for the Apollo LM are 2500 m/s for the descent and 2200 m/s for the ascent. Because of our choice of parking orbit, an extra 400m/s is needed to go between the parking orbit and the standard Low Lunar Orbit. This means 2900 m/s and 2600 m/s are required respectively, which gives us 139 m/s of margin during descent and 104 m/s of margin during ascent.

## Mission Summary

So, to recap, the mission is divided into the following steps:

- SLS Block 1 launches an ASCS, a lander and an Orion into LEO
- The ASCS performs the TLI burn after a short on-orbit checkout
- Orion separates from the stack, docks with the lander and pulls it away from the ASCS
- Orion performs the LOI burn of 500 m/s
- The crew transfers to the lander and separates from Orion
- Lander performs a 400m/s burn to circularize into a LLO
- Lander descends to the surface and lands
- Crew Performs mission, either a 1-2 day sortie or a 27 day stay in a pre-landed habitat.
- Crew lifts off in the AM
- STAR 48 burns out and is jettisoned
- AM liquid engine performs burn to LLO
- AM engine does a 400 m/s burn to rendezvous with Orion
- AM docks with Orion, crew transfers
- Lander separates from Orion (could use spare dV to crash itself into the moon)
- Orion performs TEI burn of 500 m/s
- Orion re-enters the atmosphere and lands in the ocean.

While the use of an elliptical parking orbit limits the possible surface stays either to a few hours or days or to a full lunar orbit of 27 days, it greatly reduces the mission mass, complexity and cost. While this architecture is notional and certainly not anything that could pass NASA without any significant changes, it is an architecture that proves that even the basic Block 1 SLS has the capability to do lunar missions in the same style as Apollo. I will later write an article describing a possible cargo lander, to allow this architecture to be used to set up a whole lunar base.

Some sources:

http://spirit.as.utexas.edu/~fiso/telecon/Post-Donahue_9-7-11/Post-Donahue_9-7-2011.pdf

http://www.space.com/19292-nasa-orion-space-capsule-explained-infographic.html

http://www.nasa.gov/sites/default/files/617408main_fs_2011-12-058-jsc_orion_quickfacts.pdf

http://www.nasaspaceflight.com/l2/

http://goldenspikecompany.com/wp-content/uploads/2012/02/French-et-al.-Architecture-Paper-in-AIAA-Journal-of-Spacecraft-and-Rockets.pdf

Thanks for that. I also concluded that the Block 1 SLS will have significantly more than a 70 metric ton payload capability. I discuss on my blog getting a lunar mission using all cryogenic in-space stages.

BeantwoordenVerwijderenIf you assume the payload capability of the SLS is at 90 mT or above then you can do it with hypergolic stages. To save on development costs, I favor adapting existing stages. For instance the Ariane 5 has a hypergolic upper stage that has a very good Isp and mass ratio. Two or three of these should suffice, with a hydrogen-fueled Earth departure stage.

Again to save on development costs I suggest using NASA's SEV at about a 3 mT dry mass as the lunar crew module, since NASA wants to develop it anyway.

Also, since NASA, and the ESA are not completely decided on Orion's service module, I only use the Orion capsule itself's dry mass then estimate some values for the mass of the service module based on the Ariane 5's hypergolic stage.

Bob Clark