Apollo 11 REDUX

SpaceNut · 2017-03-23 18:11:22

A contest for competition starting with the same specifications are not only bound to have the desired out come but also pricing is stated in the list of requirements of the specifications for the winner to forefill.
Pretty much the COTS contest to bring the cability for ISS resupply but for the other stuff that we would want with scaled tiered payoffs for bench marks being achieve upon reviews of progress.

SpaceNut · 2017-05-13 21:38:56

It is looking more an more like Nasa is not even going to be ready for even this small step for man and a giant fuzzle as the SLS has slipped again....

SpaceNut · 2017-05-19 20:05:19

What we know is that the Falcon 9 heavy will be able to send a crewed Dragon and Cyngus for room to stretch out in quite nicely to the moon and back but we still need a lander for the moon once we get into lunar orbit.
https://planetplots.blogspot.com/2017/0 … lunar.html
How much modification would Dragon 2 or Red Dragon require to serve as a lunar lander?
Use of Dragon Capsule for Lunar Lander
Dragon 2 or Starliner derived dedicated lunar lander?

RobertDyck · 2017-05-20 10:24:13

In 1961 NASA intended to land an Apollo CSM on the surface of the Moon, and return directly back to Earth. The problem is it was too heavy to launch with Saturn V. They tried various changes to make it work, but nothing did. Then one engineer showed them splitting into a Lunar Excursion Module (LEM) and Mothership would reduce mass. The "Mothership" would be the same Apollo capsule NASA was already working on, but a smaller Service Module. This total stack had reduced mass, so much that it would fit on Saturn V. Why believe anything has changed today? Physics is physics. Stick with the same idea.

Both Dragon and CST-100 Dreamliner were originally designed under the Crew Exploration Vehicle competition to carry astronauts to lunar orbit and back. They lost, Orion won. Most losing competitors quit, but SpaceX didn't, they just continued development of Dragon slowly. When Obama asked for commercial crew, SpaceX proposed the Dragon they were already working on, but with the service module replaced with a "trunk". They would just have to change the trunk to be a real service module again, with engines and propellant tanks to return from lunar orbit to Trans-Earth trajectory. Boeing also changed their CST-100 for Obama's commercial crew, they shrank their service module to something appropriate for ISS. They could expand their service module for the Moon.

We would then need a lunar module. Call it LEM or LM or whatever. I already posted my preferred design.

SpaceNut · 2017-05-20 17:13:40

I found this article interesting but now wonder what happened to it...as its dated 9/12/2009 Orbital Sciences to Develop Cygnus Crew Spacecraft

Launched on a human-rated version of Taurus 2, could be developed at a cost of $2 billion to $3 billion. The potential partners of Boeing were to be involved with adding a new liquid-hydrogen second stage for the Taurus 2.
Which would provide the thrust needed to carry around 8 metric tons to the space station and make it so that it could accommodate an Apollo-sized capsule based on the Cygnus cargo vehicle design along with a service module.

https://www.faa.gov/about/office_org/he … 202009.pdf

I found this while looking for lunar lander design work...
Lunar Lander Designs for Crewed Surface Sortie Missions in a Cost-Constrained Environment

What I found was the mass tables for the nasa designs are quite heavy and require SLS to get to the moon....

SpaceNut · 2017-06-23 19:11:16

Still wanting Space x to design a modified lunar lander followed by a mars capable MAV as well...

Can the Dragon V2 land on the moon

Descending and ascending the moon is much easier than on Mars. The Moon is just 17% of earth's gravity while Mars is 38%. The Apollo LM did not need any fancy inflatable re-entry vehicle to land on the moon. Rocket thrusters will suffice.
The Dragon-2 has enough power and fuel to land and ascend back to lunar orbit. The Dragon's SuperDraco has higher performance than the Apollo Lunar Module. Apollo LM descent engines has 10,000 lbf thrust, and 3,500 lbf thrust for ascent (the lander part is left behind). The Dragon's SuperDraco has 16,000 lbf thrust. The Dragon also uses computers to control throttle and land (much more fuel efficient). Hence, the Dragon-2 has enough power and fuel to land and ascend back to lunar orbit even if it carries all it's parts during ascent. However, it may need to cut crew from 7 to 3 to loose more weight for fuel.
On Earth terminal velocity for capsule of shape and size of DragonV2 or Orion at sea level is something like 96 m/s, so for powered descent you need you need decelerate about that speed in 6s to touch ground at near 0 m/s. (producing about 2,7 g-force) On Mars Terminal Velocity is much bigger ~1207 m/s at "see level" so you need to much more dV to gracefully land. Finally on Moon, there is no terminal velocity per se, since there is almost no atmosphere so 1870 m/s dV to land on Moon (from Low Lunar Orbit).

https://www.quora.com/Could-we-use-2-Sp … ar-landing

http://exoscientist.blogspot.com/2012/0 … -cost.html

Dragon V2 with truck

SpaceNut · 2017-06-23 19:22:12

Not to be forgotten GW's Reverse-Engineering What the Versions of “Dragon” Can Do

SpaceNut · 2017-06-24 18:07:09

As much as we would want a lunar landing mission to the Apollo landing site we would also want to plan a second mission to stake a claim to explore something new and to see if we can stay. Such as the South pole area of Shackleton Crater Rim https://www.nasa.gov/pdf/163896main_Exp … 120406.pdf slide 10 - 11

Slide 13 has the Altair lander to which I think we can modify the Cygnus to be able to make this design quickly a do able item. To which for additional space to live within a beam could be added to the second docking port.

The moon to mars planning.
http://images.spaceref.com/news/2007/AI … E.2007.pdf

To which the moon we will need to process the regolith for insitu oxygen.
Water may be processed from the shadows of the crater.
Radiational shielding would be via regolith piling.
Food would be grown underground possibly with natural light piped lighting.
Power will be a mix of nuclear and solar use.

RobertDyck · 2017-06-25 03:18:49

altair-constellation-returns-humans-to-the-moon-3-728.jpg?cb=1225347351
Apollo LM vs Soviet LK

Soviet LK:
Height: 5.20 m
Diameter: 2.25 m
Gross Liftoff Mass: 5,560 kg (5.56 Mt)

I said we could launch both crew Dragon with a service module instead of trunk, and a new Lunar Module based on the Soviet LK but designed to carry all 4 astronauts, no science instruments at all, and a modern docking hatch to mate with Dragon. This could be launched on a single SLS. In fact, we could use SLS block 1, however it would require an additional stage. The Interim Upper Stage (IUS) is a Delta IV upper stage, we could use that for TLI. The additional stage would be used for LOI (Lunar Orbit Insertion) and to de-orbit the new LM. The Dragon service module would only be used for TEI (Trans-Earth Injection).

This would be a crew taxi to deliver crew to the Lunar surface. A Mars Direct habitat would be pre-landed, used as a permanent Moon base, launched by SLS block 2B. The Mars Direct habitat would have long-duration life support, science instruments, and a rover.

Key statistic: crew variant of Altair = 45 metric tonnes, LK = 5.56 metric tonnes

RGClark · 2017-06-25 05:47:31

SpaceNut wrote:

Still wanting Space x to design a modified lunar lander followed by a mars capable MAV as well...
Can the Dragon V2 land on the moon
Descending and ascending the moon is much easier than on Mars. The Moon is just 17% of earth's gravity while Mars is 38%. The Apollo LM did not need any fancy inflatable re-entry vehicle to land on the moon. Rocket thrusters will suffice.
The Dragon-2 has enough power and fuel to land and ascend back to lunar orbit. The Dragon's SuperDraco has higher performance than the Apollo Lunar Module. Apollo LM descent engines has 10,000 lbf thrust, and 3,500 lbf thrust for ascent (the lander part is left behind). The Dragon's SuperDraco has 16,000 lbf thrust. The Dragon also uses computers to control throttle and land (much more fuel efficient). Hence, the Dragon-2 has enough power and fuel to land and ascend back to lunar orbit even if it carries all it's parts during ascent. However, it may need to cut crew from 7 to 3 to loose more weight for fuel.

The Dragon V2 has enough thrust to takeoff and land on the Moon but not nearly enough delta-v. I've seen the delta-v estimates it's capable of in the range of 400-600 m/s. But it takes ca. 1,850 m/s to takeoff or land on the Moon from lunar orbit.

Edit: proposals I've seen to make the Dragon, either cargo or manned version, a lunar lander involve giving it greater propellant load by adding tanks to the trunk section.

Bob Clark

Last edited by RGClark (2017-06-25 09:43:42)

Oldfart1939 · 2017-06-25 10:29:38

Would this Apollo 11 Redux just be another Flag and Footprints mission, or the start of a permanent base? That alone would influence the hardware designs. Reusability of vehicles? If this is to become a fixed base for ongoing habitation, then something needs to be done for long term solar flare radiation protection, as well as some means of attenuating the GCR loads. If the length of stay per astronaut will be < 30 days, probably OK with the radiation exposure outside of a solar flare event.

I for one, am against developing a completely new set of hardware. Make modifications to the SpaceX Dragon system to enable both landings and returns to Lunar orbit as well as Earth return. How many years will it take for development of a new Altair system? Time is not our ally in this type of development, since we've seen over the past 16 years what can happen through the 2 previous Presidencies. On again, off again, on --sort of--again. Then what? It's called use what we have in hardware and modify to make it work.

Try throwing a $5 Billion contract competition at both Orbital ATK and SpaceX and watch what happens!

SpaceNut · 2017-06-25 15:04:15

I am thinking that the first 1 or 2 mission of the "Apollo 11 Redux just be another Flag and Footprints mission" would be used to debunk to hoak sayers.... But then we would move onto starting a polar base for longer stays.

RGClark wrote:

The Dragon V2 has enough thrust to takeoff and land on the Moon but not nearly enough delta-v. I've seen the delta-v estimates it's capable of in the range of 400-600 m/s. But it takes ca. 1,850 m/s to takeoff or land on the Moon from lunar orbit.
Edit: proposals I've seen to make the Dragon, either cargo or manned version, a lunar lander involve giving it greater propellant load by adding tanks to the trunk section.

Could you point to a poor mans equation for solving to the stage enhancements....

Just found the topic you had created back in 2012 A Return to the Moon by the Apollo 11 50th Anniversary

GW Johnson · 2017-06-25 15:46:37

There is no one equation, Spacenut. Sorry. You have to use several, plus Spacex plays close-to-the-chest with some of the required data.

But I did evaluate delta-vee potential for all the Dragon versions, for propellant tanks added to the trunk regarding crew Dragon into and out of lunar orbit, and some roughouts of clean-sheet-of-paper lunar lander designs.

All of these are posted at http://exrocketman.blogspot.com.

GW

SpaceNut · 2017-06-25 15:56:38

I would think any body ascent equation as you add or remove atmospheric drag, gravity parameters would be the same and that retro propulsion would run simular to the first stage landing but thanks....

It appears that the entire Altair topic is gone from the crash.....

Will continue to add what I find still works from the other topics....
https://www.nasa.gov/pdf/203072main_LAT … 20rev2.pdf

https://www.nasa.gov/pdf/203084main_ISR … 7%20V3.pdf
Insitu production for longer stays would be a 1, 2 of water and oxygen

To which a propelant depot is not out of the realm
http://images.spaceref.com/news/2011/21.jul2011.vxs.pdf

SpaceNut · 2017-06-25 18:35:20

I do not think that there is any reason for the Altair to have a 45 mt mass.....
https://en.wikipedia.org/wiki/Altair_(spacecraft)

https://www.nasa.gov/pdf/289914main_fs_ … lander.pdf

We should be able to do better than that as the only thing for LOX and LH2 is fuel cell and water from boiloff. Sure the down mass can be greater for descent but is at the much large fuel tank and extra insulation to slow the rate of boil off...and this gets more severe with the ascent stage also using the same fuel...

Lunar Orbit Insertion Targeting and Associated Outbound Mission Design for Lunar Sortie Missions

From Apollo LM to Altair: Design, Environments, Infrastructure, Missions, and Operations

Oldfart1939 · 2017-06-25 19:12:05

I would prefer to see a non-cryogenic fuel system in place: NTO + UDMH. This eliminates all the oversize tankage at a penalty of a somewhat reduced Isp.

SpaceNut · 2017-06-25 19:22:17

This 2 pg in the above post https://www.nasa.gov/pdf/289914main_fs_ … lander.pdf was with those types of fuels for the ascent stage but it was still oversized with the descent stage still being cryogenic.....

Hold the presses I did find the altair topic or at least what still remains http://newmars.com/forums/viewtopic.php?id=4713

Lunar Lander Ascent Module Configuration and Propulsion Studies

RobertDyck · 2017-06-25 19:54:04

Oldfart1939 wrote:

I would prefer to see a non-cryogenic fuel system in place: NTO + UDMH. This eliminates all the oversize tankage at a penalty of a somewhat reduced Isp.

LOX and LCH4 are soft cryogens. They remain liquid at relatively warmer temperatures. The temperature of space is -270°C, but the Sun adds a lot of heat. Spacesuits for use in Earth orbit are designed to endure -150°C to +120°C. At pressure of typical propellant tanks at launch, these soft cryogens remain liquid in that temperature range. So all we need is good insulation.

Oldfart1939 · 2017-06-26 08:21:08

A lot depends on the duration of the Lunar stay. For another "Flag and Footprints," sure; use the soft cryogens. For longer duration surface missions, even the slow boil-off of LOX would eventually be problematic. I'm actually proposing a fairly conservative approach, and since the combination I suggest is hypergolic--no worries about an engine restart. Yes, CH4 has a great Isp, but the density is low and requires larger and insulated tankage. The Id of Aerozine-50 is excellent, and whether we use NTO or LOX---it will work well with either oxidizer.

RobertDyck · 2017-06-26 09:48:51

The Moon has 2 weeks of sunlight followed by 2 weeks of dark. I'm envisioning a minimal lunar module to act as a crew taxi. A Mars Direct habitat left permanently as a Moon base. The new stage I said would be used for LOI and to de-orbit the lunar module; it only has to endure 3 days transit to the Moon, then will be expended. The lunar module itself will park on the surface of the Moon for 2 weeks in direct sunlight, then leave. It will be expended. The service module for Dragon or CST-100 Starliner will be parked in lunar orbit for those same 2 weeks, then use that propellant for TEI. Wouldn't LOX/LCH4 work for that?

Yes, aerozine-50 has higher density. But remember the Soviet N1? It used LOX/RP1 for all 4 stages: blok A/B/V/G. Even though the N1 blok A stage was larger and more powerful than the S-1C stage of Saturn V, total throw mass to trans-lunar trajectory was lower. That's because Saturn V used LOX/LH2 for S-II and S-IVB stages. LH2 has the lowest density of all, but highest Isp.

Oldfart1939 · 2017-06-26 10:11:52

In times past, we used the highest performing fuels available for sake of simplicity. The entire Apollo project was in reality, a "quickie,' which demanded results FAST. We made some mistakes, as did the Soviets. Today, the Russian heaviest lifter is the Proton M, which uses NTO and UDMH as fuels. We now have orbital assembly abilities we lacked at the time of Apollo, and a demand for reuse of components. We also know that Hydrazine alone is more energetic than Aerozine-50 and UDMH, but is too unstable for use in regeneratively cooled motors; Aerozine-50 is stable enough for that application and still energetically robust. The subsequent tankage can be considerably smaller and lighter than tanks for either Hydrogen or Methane storage and containment, since Aerozine-50 doesn't require pressure containment or lots of insulation. Using multiple Falcon Heavy launches in place of the huge Saturn V, we can actually get what we need to Lunar orbit with greater mass and considerably cheaper.

RGClark · 2017-06-26 10:44:08

RobertDyck wrote:

https://image.slidesharecdn.com/vonbrau … 1225347351
Apollo LM vs Soviet LK
http://astronautix.com/nails/l/lemvslk.jpg
Soviet LK:
Height: 5.20 m
Diameter: 2.25 m
Gross Liftoff Mass: 5,560 kg (5.56 Mt)
I said we could launch both crew Dragon with a service module instead of trunk, and a new Lunar Module based on the Soviet LK but designed to carry all 4 astronauts, no science instruments at all, and a modern docking hatch to mate with Dragon. This could be launched on a single SLS. In fact, we could use SLS block 1, however it would require an additional stage. The Interim Upper Stage (IUS) is a Delta IV upper stage, we could use that for TLI. The additional stage would be used for LOI (Lunar Orbit Insertion) and to de-orbit the new LM. The Dragon service module would only be used for TEI (Trans-Earth Injection).
This would be a crew taxi to deliver crew to the Lunar surface. A Mars Direct habitat would be pre-landed, used as a permanent Moon base, launched by SLS block 2B. The Mars Direct habitat would have long-duration life support, science instruments, and a rover.
Key statistic: crew variant of Altair = 45 metric tonnes, LK = 5.56 metric tonnes

As described on the Astronautix.com page, the Soviet LK did not do the full landing on its own. It required the Soviet Block D stage to do most of the propulsion towards the lunar surface, bringing the LK to within 5 km and 100 m/s of the landing. The rest of the landing was done by the LK, as well as doing the take-off.

This type of landing used what's called a "crasher stage" since the Block D would crash into the lunar surface. I'm not a fan of this type of landing, though it's probably just a visceral reaction with this stage crashing and exploding on the lunar surface, even though kilometers away from the lander.

In any case the mass of the LK stage was comparable to the Apollo's ascent stage only, and the full lander was not at a saving in mass to the Apollo full lander.

About the Block D stage:

Block D space tug.
http://www.russianspaceweb.com/n1_d.html

For a replacement for this Block D stage, I like the storable propellant version of the Ariane 5 upper stage:

Ariane 5-2
N2O4/MMH propellant rocket stage. Storable propellant, restartable upper stage for use with Ariane 5. Chamber pressure 10 bar; expansion ratio 83.0; propellant mix ratio 2.05. Empty mass without VEB payload fairing support ring and avionics is 1200 kg.
AKA: L-9. Status: Active. Thrust: 27.40 kN (6,160 lbf). Gross mass: 12,500 kg (27,500 lb). Unfuelled mass: 2,700 kg (5,900 lb). Specific impulse: 324 s. Burn time: 1,100 s. Height: 3.36 m (11.02 ft). Diameter: 3.96 m (12.99 ft). Span: 5.46 m (17.91 ft).

Cost $ : 6.000 million.
http://www.astronautix.com/a/ariane5-2.html

Note that it's actual dry mass without the extra equipment used for its Ariane 5 function would only be 1,200 kg. Given also its high Isp, it could serve as a full stage descent stage, i.e., no separate crash landing for the stage.

Also, nice is its low cost.

Bob Clark

Last edited by RGClark (2017-06-26 10:49:57)

RobertDyck · 2017-06-26 11:37:30

Ok, let's compare.

Apollo

Soviet L3, consisting of Soyuz LOK, LK, Block D, a fairing around LK to support weight of Soyuz, and Block G.
- Soyuz LOK: 9,850 kg
- LK: 5,560 kg
- Block D: 13,360 kg
Total: 28,770 kg

L3 was said to have gross mass 95,000 kg, and Block G was 61,800 kg, so that means the fairing should mass 4,430kg. But there were 2 fairings: an inner fairing to support the weight of Soyuz during launch, and an outer aerodynamic fairing. So I'll leave it out.

Apollo: CM, SM, LM
- CSM: 30,329 kg
- LM (Apollo 9-14): 15,200 kg
- LM (Apollo 15-17): 16,400 kg
Total (Apollo 15-17): 46,729 kg

Soviet L3 would carry 2 cosmonauts, of which 1 would descend to the Moon, while Apollo carried 3 astronauts, of which 2 descended. But still, total mass is significant.

One thing that got me going on this was Orion. The Orion service module is only designed for TEI; it isn't capable of LOI. Even the original designed service module would only have been capable of TEI. Constellation would have used an over-size descent stage of Altair for LOI. Since announcement of SLS, architecture for Orion has been a stage for LOI just like Block D. Since they're using the Soviet architecture anyway, let's continue to the lunar surface. (But using a much lower mass capsule.)

Last edited by RobertDyck (2017-06-26 11:38:08)

Oldfart1939 · 2017-06-26 19:44:14

Seriously, given the rapid advances made by SpaceX, I doubt the Orion capsule will ever accomplish anything significant. It's simply too heavy and unwieldy for anything beyond a ballistic flight beyond the moon. The key to the future will be modularity and in orbit assembly. Have to give Armstrong, Aldrin, and the follow-on lunar landers the award for having big balls--just flying those flimsy and uber-light spacecraft to the surface and then taking off for a rendezvous with the Earth return vehicle was incredible.

Last edited by Oldfart1939 (2017-06-26 19:45:01)

RobertDyck · 2017-06-27 00:50:42

Oldfart1939 wrote:

I doubt the Orion capsule will ever accomplish anything significant. It's simply too heavy and unwieldy
...
flimsy and uber-light spacecraft

Sounds like you contradicted yourself in one paragraph. When you criticize "flimsy and uber-light" you give an excuse for the over-heavy Orion. I think we all know the real reason. Orion was designed to launch on Ares I, aka "The Stick". A single SRB produced too much vibration, it required payload mass to dampen the vibrations. So they designed Orion to be heavy; overly heavy. That's backwards. Any spacecraft has to be light.

Apollo Command Module (CM): 5,560 kg
Dragon v1, including trunk: 4,200 kg + 1,290 kg propellant = 5,490 kg
Orion capsule, not including service module, LES, blast shield, or fairing: 10,387 kg

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