Apollo 11 REDUX

Oldfart1939 · 2017-02-17 17:26:59

I wasn't referring to the Habitat module as a throwaway, simply the ultralight lunar landers. I'm willing to bet the cost of development from Northrup-Grumman would exceed $250 Million! Using existing and future useful hardware could keep the entire program under a half Billion.

The other problem: we don't HAVE a Mars habitat module, either.

SpaceNut · 2017-02-17 18:25:39

Lets see if we can refocus on just getting to the moon and back if we can.

From the other topic the EDS needs to be able to produce 4.1 Km/s from LEO to get us to the moons orbit.

From the moons orbit to return home we will need just 0.7 km/s for the return journey.

Since our limitations for the moons first mission on the sunny side would be a max of 14 days stay with the use of solar panels.

Lets see if we can use just 2 Falcon heavy rockets to launch the capsule, service module, EDS stage to LEO and on the other launch the New LEM and lunar return stage.

The New Lem is based on the Dragon Truck less parachute, and heat shield with a new first and second stage to place under where the truck would be. The new first stage would be only slightly longer than the extended truck less solar panels to make the fuel requirements for ascent to orbit.The engine still needs to spec'd out for this new stage. The descent to surface or second stage would be would be 2 if not 3 full extended trucks with the set of solar panels at the top of the stage. It would carry the fuel to land and legs that would be left on the surface just like in the Apollo mission.

SpaceNut · 2017-02-17 18:26:51

Substitute for the original LEM

https://en.wikipedia.org/wiki/Draco_(ro … ne_family)
Propellant NTO / MMH[1]
Performance Thrust (vac.) 400 N (90 lbf)

Super DrACO
Thrust (SL) 73 kN (16,400 lbf)
Isp (SL) 235s
Burn time 25 seconds
Propellant capacity 1,388 kg (3,060 lbs)
The engine has been tested “in both a launch escape profile and a landing burn profile” and can be throttled between 20 percent and 100 percent thrust levels for precision landing.

http://www.astronautix.com/r/rs-18.html
Thrust: 15.56 kN (3,499 lbf). Specific impulse: 310 s.

lunar ascent engine redux with methane:
http://www.spaceref.com/news/viewpr.html?pid=26327

The cryogenic propellant-fueled RS-18 engine, Engineers from NASA and Pratt & Whitney Rocketdyne successfully completed a series of hot-fire altitude tests using liquid methane

https://en.wikipedia.org/wiki/Raptor_(r … ne_family)
Raptor is a family of cryogenic, methane-fueled rocket engines under development by SpaceX.
Thrust (vac.) ~3,285 kN (738,000 lbf)
Thrust (SL) 3,050 kN (690,000 lbf)
Chamber pressure 30 MPa (4,400 psi)
Isp (vac.) 361 s
Isp (SL) 334 s

https://en.wikipedia.org/wiki/TR-201

https://en.wikipedia.org/wiki/LR-87

SELECTED ROCKETS AND THEIR PROPELLANTS
PROPERTIES OF ROCKET PROPELLANTS and ROCKET PROPELLANT PERFORMANCE
http://braeunig.us/space/propel.htm

Altair Lunar Lander
https://www.nasa.gov/pdf/278869main_092 … 4final.pdf

Aerojet OMS or OME
http://www.astronautix.com/o/ome.html

Aerojet AJ10-118K
http://www.astronautix.com/a/aj10-118k.html

https://en.wikipedia.org/wiki/DragonFly_(rocket)

"DragonFly RLV" in that PDF:

- up to four steel landing legs
- weighs 14000 lbs unfueled
- maximum proplellant load is 400 gallons
height of 17 ft and a base width of 13 ft.

http://1.bp.blogspot.com/-_aFERa82jaI/U … t1_1_1.jpg

SpaceNut · 2017-02-17 18:28:44

The original LEM Dimensions
Height 22 ft 11 in. (with legs extended)
Diameter 31 ft (diagonally across landing gear)
Weight 32,500 lb (approx.) (with propellant and crew)
Weight (dry) 9,000 lb (approx.)

Pressurized volume 235 cu ft
Habitable volume 160 cu ft

Ascent stage
Height 12ft 4 in.
Diameter 14ft 1 in.
Weight (dry) 4,850 lb (approx.)
Batteries: two 28–32 volt, 296 ampere-hour silver-zinc batteries; 125 lb (57 kg) each
Power: 28 V DC, 115 V 400 Hz AC
Atmosphere: 100% oxygen at 4.8 psi (33 kPa)
Water: two 42.5 lb (19.3 kg) storage tanks
Coolant: 25 pounds (11 kg) of ethylene glycol / water solution

Descent stage
Height 10ft 7 in.
Diameter 14ft 1 in.
Weight (dry) 4,300 lb (approx.)
Water: one 151 kg (333 lb) storage tank
Batteries: four (Apollo 9-14) or five (Apollo 15-17) 28–32 V, 415 A·h silver-zinc batteries; 135 lb (61 kg) each

Propellant
Ascent stage 5,170 lb tanked
APS thrust: 3,500 lbf (16,000 N)
APS propellants: Aerozine 50 fuel / nitrogen tetroxide oxidizer
APS pressurant: two 6.4 lb (2.9 kg) helium tanks at 3,000 pounds per square inch (21 MPa)
APS specific impulse: 311 s (3,050 N·s/kg)
APS delta-V: 7,280 ft/s (2,220 m/s)

Descent stage 17,880 lb tanked
DPS engine: TRW LM Descent Engine (LMDE)[16]
DPS thrust: 10,125 lbf (45,040 N), throttleable between 10% and 60% of full thrust
DPS propellants: Aerozine 50 fuel / nitrogen tetroxide oxidizer
DPS pressurant: one 49-pound (22 kg) supercritical helium tank at 1,555 psi (10.72 MPa)
DPS specific impulse: 311 s (3,050 N·s/kg)
DPS delta-V: 8,100 ft/s (2,500 m/s)

RCS 605 lb tanked
RCS thrusters: sixteen x 100 lbf (440 N) in four quads
RCS propellants: Aerozine 50 fuel / nitrogen tetroxide (N2O4) oxidizer
RCS specific impulse: 290 s (2,840 N·s/kg)

http://www.nextbigfuture.com/2016/05/sp … anned.html

SpaceNut · 2017-02-18 13:31:45

Much like the moon flyby we are stuck with what we do not have and with programs that got cancelled that had they been allow to finish we would have engines and fuels for this mission.
Many of the engines are no longer in production and while they could be revived its not likely to happen.

AIAA-2013-5479 American Institute of Aeronautics and Astronautics 1 Lunar Lander Designs for Crewed Surface Sortie Missions in a Cost-Constrained Environment

Which leaves us to modifiy design based on what do we have rather than what we would want to use.
Since we need the EDS and all we have for engines are the J2 version, RL-10 for LH2/Lox and the Kerosene/Lox engines we will be making a new stage to go to the moon no matter what we launch it on.

SpaceNut · 2017-02-23 21:56:06

Computational Fluid Dynamics Validation of a Single Central Nozzle Supersonic Retropropulsion Configuration

A VALUE PROPOSITION FOR LUNAR ARCHITECTURES UTILIZING PROPELLANT RESUPPLY CAPABILITIES

The Gryphon: A Flexible Lunar Lander Design to Support a Semi-Permanent Lunar Outpost

Robust Lunar Exploration Using an Efficient Lunar Lander Derived from Existing Upper Stages

Oldfart1939 · 2017-02-24 10:15:49

SpaceNut-

After reading the article listed in your previous comment, Robust Lunar Exploration...Existing Upper Stages, it struck me that use of LH2/LOX is very seductive, but brings many problems to the table. And what has really happened, the dramatically reduced cost per launch to LEO brought about by reusability changes the entire paradigm. LH2 has a great exhaust velocity and hi Isp, but is offset by a very low density requiring larger and well-insulated tankage. My argument to all these other proposals is--live with the non cryogenic MMH and NTO, use more physically robust hardware, and simply offset the weight penalties by using cheaper heavy lift to LEO. Maybe my approach is too "brute force" for the geniuses at ULA and NASA.

elderflower · 2017-02-24 11:50:20

Even using liquid methane would save a lot due to increased density and smaller temperature difference from ambient, compared to liquid hydrogen. The reduction in Isp is not so big that it wont be offset by the reduction in vehicle weight due to smaller tanks and less insulation.

Oldfart1939 · 2017-02-24 12:29:59

Using LCH4 is fine, but the density is still fairly low, requiring larger tankage than--say--MMH. The combination of vacuum engine performance as stated by Musk for the new raptor engines is very good, though. The big imponderables are the freezing properties of various fuels and a stay on the lunar surface. LH2 is definitely NOT going to freeze into a lump in the tanks. I haven't checked the various melting points of the fuels, other than to say RP-1 is a total no-go.

SpaceNut · 2017-03-05 21:02:25

Oldfart1939 wrote:

GW-
An interesting concept would require use of another component in the system: the Russian-built Proton M 3rd stage. Here are the details:
Type: Storable Propellant Stage
Inert Mass: 4,185 kg
Diameter: 4.14 meters
Length: 6.5 meters
Propellant: UDMH
Oxidizer: NTO
Fuel & Ox: 46,562 kg
Total thrust:613.8kN
Burn time: 238 seconds
Isp(vac): 325 seconds
My suggested model incorporates the Falcon 2 + Trunk, + dry Proton M 3rd stage to ~20,000kg, including upgraded onboard fuel for retro propulsive landing on Earth. I'm figuring on a delta V around 3.5 km/sec for the rocket equation. This gives a mass ratio of ~ 3.0.
If we configure the Dragon 2 for a crew of 4, with onboard supplies for food, an atmosphere purifier, no water recycling at 11,000 kg, adding tankage and additional fuel over and above the "normal" 1900 kg at 3,800 kg (total fuel = 5,700 kg), the empty mass of the Proton M 3rd stage at 4,185 kg, and a trunk with a motor at 1500 kg, we get a payload mass of 20,485 kg. Fully fueled, this results in a total "wet" rocket mass of 67,047 kg with a mass ratio of 3.27. This implies there will be fuel remaining after departure for insertion into Lunar orbit. The increased fuel for the Dragon capsule also implies sufficient fuel for de-orbit burn and adequate fuel for dry land propulsive landing.
This is not the calculation of a sophisticated aerospace engineer, just an amateur. Please make comments! This requires orbital assembly and 2 loads to LEO. The mass of the Proton M 3rd stage is well within the capability of the Falcon Heavy.

SpaceNut · 2017-03-05 21:04:04

SpaceNut wrote:

Now you all see why we are having trouble keeping topics organize with a scope of discusion an ther is so much out of the box comments.
The trouble with a progress stage is the coupling of them does not match up as the core of the falcon is 3.7 m in diameter which now means interstage coupling collars are need to make an off the shelf work at all.
The again it is just a small thing as even an extended stage from Space x would also need a interstage coupling collar to mate the stages together for such a flight to occur thou these would be simpler as they already make use of them. I am sure with modifications to those that we would have a better match up of the sections to build the lunar rockets components up at the ISS.
Space x can only move so fast with all the changes that are needed in order to progress what they have to being able to do the flyby and then to proceed to being able to create something to think about landing with.
If I were the other rocket launch companies out there I would be looking to build up from what they have a rocket capable of the same feat. This is what speers on competition as its availability to provide at a cost of affordability that we need.

SpaceNut · 2017-03-06 19:28:28

GW Johnson wrote:

I think it was in another thread here in the "human missions" group where I addressed getting into lunar orbit with existing Spacex hardware (Falcon-Heavy flinging crew Dragon). It's very hard to tell, and it's a close thing with no margins, but I think the second stage would have enough propellant left to get into lunar orbit with a lightly-loaded crew Dragon.
[edit 3-4-17: it was in post 15 above in this very thread that I mentioned getting crew Dragon into lunar orbit at reduced weight.]
There's barely enough delta-vee in a lightly-loaded crew Dragon's Super Draco/Draco system to get back out of lunar orbit (~0.8 km/s). That forces parachute-only landings at sea. I rather doubt they really want to do that, precisely because it's risky without an alternate for landing. They really do need a service module for crew Dragon, or at least a trunk with extra propellants for the thrusters and a way to connect them.
Once that service module or extra propellant is in place, they would have the capability to reprise Apollo-8 style lunar orbit missions with crews, and without the light-load limitations they have now. I'm going to hazard the guess that this is something they are working on, but have not yet announced publicly. It just makes too much sense for them not to be looking at it
From there, you use a second rocket to shoot unmanned a lunar lander to lunar orbit, then shoot the men there to go rendezvous with it. At that point, they can reprise the Apollo 11 type missions, as long as the lander falls in the same 12 ton class of flung weight that crew Dragon does. That lander would need something like a total 3.4 to 3.5 km/s minimum delta vee capability to get from lunar orbit to the surface and back again.
GW

SpaceNut · 2017-03-06 19:30:14

GW Johnson wrote:

Referring to my post 34 just above, I ran my estimated weight statements and performance for crewed Dragon with extra propellant tanks installed in its trunk, staying within the 3000 kg published trunk cargo limit. I put 2800 kg extra propellants contained in 147 kg of tank inerts in that trunk, and just 2 suited crew plus 500 kg samples and supplies in the capsule. It weighs just a smidge over 12 metric tons as flung by its booster.
That extra tankage is a 95-5 split on propellant vs tank to contain it. Some sort of link between trunk propellant and what the Super Dracos can use is implied, but I suspect that's plumbing hardware around 5-10 kg or less, so I ignored that, as being well within my weight statement uncertainties.
It's capsule-plus-trunk delta-vee that is of interest here, since the electricity comes from solar cells mounted on the trunk skin, in crewed Dragon. You'll need that for the trip from the moon back to Earth, jettisoning the trunk right before Earth entry interface.
My results under those assumptions show a total capsule-plus-trunk delta-vee capability of 1.6 km/s. That's enough to cover 0.8 km/s to get from lunar orbit into a return trajectory, some significant midcourse and attitude control, plus (after trunk jettison and entry) something like 0.5+ km/s for a propulsive landing on land back on Earth. Maybe a big +.
The total mass for crewed Dragon rigged like this was just over 12 metric tons. Spacex says its Falcon-Heavy can fling 13.6 tons to Mars, although they do not specify what trajectory. For a min-energy Hohmann transfer ellipse, and 5% gravity and drag losses on the first 8 km/s, I show a trajectory to Mars as at least 12.1 km/s. I show (under the same 5% losses) a delivery into lunar orbit as no more than 12.4 km/s. Those correspond to the factored theoretical delta-vees, the factored ones being the ones that set your mass ratios for your designs.
12.2 tons vs 13.6 tons. 12.4 km/s vs 12.1 km/s. Looks to me like Falcon-Heavy can deliver a 12.2 ton crewed Dragon to lunar orbit, from which it can return, if rigged with 2800 kg extra propellant in the trunk. That's a crew of 2 in suits with 500 kg of samples and supplies. Add a third crewman with suit and supplies, and you start pushing toward 13 tons.
GW

SpaceNut · 2017-03-06 19:31:49

GW Johnson wrote:

Relative to my post 43 just above, I compiled and posted what I did reverse-engineering what the variants of Dragon could do. This is posted as an article dated 3-6-17 entitled "Reverse-Engineered "Dragon" Data". The site is my "exrocketman" blog site http://exrocketman.blogspot.com.
I agree with Oldfart1939: Mars is not premature at this time. We have known everything we need since the mid 1990's, and we have had far cheaper launch rockets for almost a decade now.
As for putting a Bigelow inflatable around the moon, well, why not? B330 lists as 20 tons, not within what Falcon-Heavy can deliver to lunar orbit, using the second stage to enter that orbit. But a smaller version at 12 tons could be.
SLS could do that B330 20 ton thing to lunar orbit. Atlas-5 cannot. Delta-4 cannot. Falcon-Heavy cannot quite. Nor can Falcon-9. How big a station you want? B330's dock together. Reduced-size modules could, too.
Refuel in Earth orbit, and Falcon-Heavy could do the job, even with a full-size B330.
GW

SpaceNut · 2017-03-06 19:54:12

As post #62 copy indicates we can go to the moon and now land a lander once designed comeback up rendevous with it change back to the dragon decouple and come home leaving the lander for a future use parked in orbit.

Post #63 indicates only minimal change to the trunk but its just a guess that there is nothing to alter on the newer man capable vehicle cargo unit.

#63 is really a waste of time early on for moon missions but as a hotel in orbit I would think that it needs lots more protection than the inflateable would have.

Oldfart1939 · 2017-03-06 23:03:57

For those wondering about toxic propellants, i.e. Hydrazine and NTO:

https://youtu.be/bDRKeM9kKxs

As a retired chemist, these compounds are not that big a deal if good handling practices are used.

elderflower · 2017-03-07 04:07:40

Would methylamines (probably mixed, as it comes out of the reaction) serve for fuel with NTO? Its relatively easy to make once you have made ammonia. Not so toxic as Hydrazine derivatives, and more stable. I don't know whether it would be hypergolic but I suspect that it could be.

Oldfart1939 · 2017-03-07 10:54:13

In order to deal with toxic propellants such as hydrazine, the Aerospace Engineers need to get in touch with some very good Chemical Process Engineers or experienced chemical plant managers. I worked with huge quantities of Hydrazine at times and am still kicking around. In chemical process industry, and using big plant scale reactors, a step is taken in addition of hydrazine to the reactors called "blowing down the transfer lines," which uses high pressure Nitrogen or Argon to literally blow out the lines from all residual reagents. Worried about the rocket motor having hydrazine residues dripping out? Blow down the lines and remove all vestiges from the motor. Yes, here on Earth that would contaminate the immediate area temporarily, but on the Moon, it would undoubtedly evaporate and vanish. Ditto Mars. Cleanup of Hydrazine contaminated equipment is easily accomplished with ordinary household bleach--Sodium Hypochlorite solution. Standard treatment for Hydrazine spills in the chemical industry. In my plant, we normally had at least two 55 gallon drums of hydrazine on hand almost all the time. We also had 6 drums of double strength Sodium Hypochlorite solution stored in the event of a Hydrazine spill. We also had full face mask chemical protection with cartridges specific to nitrogenous bases. My wife, in addition to being a Case Institute of Technology Chemistry graduate, was also a highly skilled plant process chemist and operator. In 24 years of operation, we suffered ZERO lost time incidents w/r Hydrazine and doing chemistry with it. And no EPA reportable incidents, either.

RGClark · 2017-03-08 08:50:59

Orbital Sciences has been investigating giving their Cygnus capsule life support capability for use as a deep space habitat. They've also investigated giving it a heat shield to make it reusable. A Cygnus given a heat shield and life support could work as a lightweight capsule for a lightweight lunar sortie design, launchable on a single Delta IV Heavy or Ariane 5:

http://exoscientist.blogspot.com/2013/0 … ights.html

Bob Clark

SpaceNut · 2017-03-08 20:50:47

Was looking for another stage to make use of and its not a good fit but at least its still in production. https://en.wikipedia.org/wiki/Delta-K
The Delta second stage uses a single AJ10-118K rocket engine, fueled by Aerozine 50 and dinitrogen tetroxide, which are hypergolic.

RGClark · 2017-03-09 00:31:58

Another hypergolic stage still in production is the hypergolic upper stage on the Ariane 5:

http://www.astronautix.com/a/ariane5-2epsl10aestus.html

The Ariane 5 can use a hypergolic upper stage or hydrolox one depending on mission.

Going for all hydrolox stages for the in-space propulsion for a lunar mission would make for a smaller mission size. But hypergolics would offer surety of ignition at least for the stages needed for Earth return.

Bob Clark

RobertDyck · 2017-03-09 05:34:41

Depends what you're trying to do. Orion currently uses a European service module, from ATV. That is capable of returning Orion from Lunar orbit. Since the CST-100 Starliner or Dragon capsule are both lower mass, you could use that service module for Earth return. It doesn't have enough propellant for both Lunar Orbit Insertion (LOI) and Trans-Earth Injection (TEI), but does for the latter. Current plans are that Exploration Mission 2 (EM-2) will use the Exploration Upper Stage (EUS) for both Trans-Lunar Injection (TLI) and LOI. That allows Orion to TEI using the ATV-based service module. One simple option is to use the ATV-based service module for Dragon or CST-100. However, there are issues with both.

Currently the launch escape system for Dragon v2 keeps the trunk attached until the launch abort system runs out of propellant. The Dragon v2 trunk has fins for aerodynamic stabilization. The control system would have to be modified to launch abort without the trunk. A service module of any sort would have substantially more mass than any trunk, so launch abort would never work that way. Orion includes a fairing around the ATV-based service module; you could design a fairing for Dragon that remains attached during launch abort and has fins, but a far far far better system would be to launch abort without any trunk.

CST-100 Starliner uses its service module for launch abort. That requires a brief high thrust burn. I don't believe the ATV-based service module can do that. So CST-100 Starliner would really require a new service module for a Lunar mission.

Whatever combination you choose for launch abort and TEI, you can use hydrolox (LH2/LOX) for Earth orbit departure aka Trans-Lunar Injection (TLI).

GW Johnson: to keep acronyms clear, Apollo used the term "Launch Escape System" (LES), while Orion calls it "Launch Abort System" (LAS).

Last edited by RobertDyck (2017-03-09 09:12:37)

Oldfart1939 · 2017-03-09 08:33:57

Just comment, not aimed at anyone in particular: too many acronyms can make an article virtually unreadable to most.

GW Johnson · 2017-03-09 10:07:23

My weight statement numbers for crew Dragon differ some with some of the other numbers being bandied about in these forums. I have those numbers, and their pedigree/derivation (for however much that is worth) posted over at "exrocketman".

I was able to identify from numbers posted on Spacex's website for Falcon-Heavy that it has enough "oomph" to put a crew Dragon into lunar orbit from a direct surface launch from Earth. I'm not sure, but I think this could be with a stop in LEO just like Apollo. This would not use any propellant from the Dragon, just a third and final burn from the rocket second stage. This ignores what you do with that spent stage now in orbit about the moon.

Crew Dragon with a trunk as it is so far, could possibly get back out of lunar orbit, if the crew/cargo were small enough, because the delta-vee is 0.8 km/s to do that. My weight statements say it barely has that capability if the load is light enough. That leaves nothing for landing, forcing an ocean splashdown with chutes. And I don't know about the budget for attitude control, critical during reentry.

Thus it makes very good sense to first test that system as a flyby mission, not an Apollo-8 style lunar orbit mission. Too marginal, too many ways to fail.

Adding propellant to the trunk changes that picture, if you can connect that supply to the thrusters in the capsule. You have to stay within the weight and volume limits for the trunk so as not to be too overweight to launch, or to overload the structures. I added that to my weight statements posted, and got 1.6 km/s delta-vee capability for a lightly-loaded capsule. That makes lunar orbit departure possible with a budget for attitude control and a budget for propulsive landing on land.

It does affect launch abort: when the capsule sheds the trunk, that trunk crashes, of course. Up to now, that's a dry structure. If you add propellant to the trunk, its crash site in a launch abort scenario will be contaminated with tons of MMH and NTO. It's nothing that can't be handled, but it is something that adds extra fuss and bother. Especially with onlookers adjacent to the launch site.

Once you do have lunar orbit capability with crew Dragon/Falcon-Heavy by adding propellant to the trunk, then all you need is a lunar lander to reprise the Apollo landing missions. If the lander fits within Falcon-Heavy's 13.6-ish ton throw weight to lunar orbit, then you just fling it there with a second Falcon-Heavy launch. You could reprise an Apollo-type landing that way.

If you need a bigger lander to start building some sort of base on the moon, then you either resort to assembly in Earth orbit, or you wait and pay the exorbitant price to fling it there with SLS.

Or maybe Bezos's New Glenn might do the job. Hard to say yet. Probably ~5 years off, with SLS ~1-2 years off. Falcon-Heavy should fly this year for the first time. So it's maybe ~1 year off, assuming it doesn't fail in test.

So, what you fly really depends more about what you choose to accomplish on the moon, more so than just getting there at all (as it was with Apollo). The real question before us: who is going to recreate lunar capability first? NASA or one or more of the private firms? Neither Musk nor Bezos is working on a lander, as near as I can tell. The various potential lander guys need to step forward and get into this.

The price is what determines who can stay in the game longer. From that standpoint, NASA/SLS is a loser.

GW

Last edited by GW Johnson (2017-03-09 10:22:40)

Tom Kalbfus · 2017-03-10 10:34:18

Oldfart1939 wrote:

Just comment, not aimed at anyone in particular: too many acronyms can make an article virtually unreadable to most.

But it makes you look intelligent if you use them! Anyone who doesn't know them will just nod their head to look thoughtful and polite.

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#51 2017-02-17 17:26:59

Re: Apollo 11 REDUX

#52 2017-02-17 18:25:39

Re: Apollo 11 REDUX

#53 2017-02-17 18:26:51

Re: Apollo 11 REDUX

#54 2017-02-17 18:28:44

Re: Apollo 11 REDUX

#55 2017-02-18 13:31:45

Re: Apollo 11 REDUX

#56 2017-02-23 21:56:06

Re: Apollo 11 REDUX

#57 2017-02-24 10:15:49

Re: Apollo 11 REDUX

#58 2017-02-24 11:50:20

Re: Apollo 11 REDUX

#59 2017-02-24 12:29:59

Re: Apollo 11 REDUX

#60 2017-03-05 21:02:25

Re: Apollo 11 REDUX

#61 2017-03-05 21:04:04

Re: Apollo 11 REDUX

#62 2017-03-06 19:28:28

Re: Apollo 11 REDUX

#63 2017-03-06 19:30:14

Re: Apollo 11 REDUX

#64 2017-03-06 19:31:49

Re: Apollo 11 REDUX

#65 2017-03-06 19:54:12

Re: Apollo 11 REDUX

#66 2017-03-06 23:03:57

Re: Apollo 11 REDUX

#67 2017-03-07 04:07:40

Re: Apollo 11 REDUX

#68 2017-03-07 10:54:13

Re: Apollo 11 REDUX

#69 2017-03-08 08:50:59

Re: Apollo 11 REDUX

#70 2017-03-08 20:50:47

Re: Apollo 11 REDUX

#71 2017-03-09 00:31:58

Re: Apollo 11 REDUX

#72 2017-03-09 05:34:41

Re: Apollo 11 REDUX

#73 2017-03-09 08:33:57

Re: Apollo 11 REDUX

#74 2017-03-09 10:07:23

Re: Apollo 11 REDUX

#75 2017-03-10 10:34:18

Re: Apollo 11 REDUX

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