Apollo 8, redux

RobertDyck · 2017-02-25 15:34:57

GW Johnson wrote:

I am ASSUMING that HIAD is the extendible heat shield, not the inflatable (I think, at least I hope, that ADEPT is the inflatable).

Other way around. ADEPT is the extendable one with coated carbon fibre fabric that opens like an umbrella. HIAD is the one that inflates.

GW Johnson · 2017-02-25 17:00:21

Well, I told everyone there were too many acronyms floating around, didn't I?

GW

SpaceNut · 2017-02-25 21:07:15

There is a third combination of the two in what some are calling the hypercone... inflateable tires lashed together in a shape of a cone with the carbon fiber fabric on the outside to act as a heashield that is expanded prior to atmospheric entry.

Whether nasa will become less risk adverse is the real question and put in the additional testing to make the first mission on time without pilot/prototype runs.....

kbd512 · 2017-02-27 11:33:44

GW,

IMLEO - Initial Mass, Low Earth Orbit - a measure of how much mass is required in LEO for a particular part of a mission or mission profile
HIAD - Hypersonic Inflatable Aerodynamic Decelerator - the inflatable donuts heat shield that uses mass shifting for directional control
ADEPT - Adaptable, Deployable Entry Placement Technology - the umbrella heat shield that uses mass shifting for directional control
DGB - Disc Gap Band - a type of supersonic ring sail parachute that NASA has used on Mars at Mars local Mach 2.5 - part of how Curiosity was soft landed

NASA's mass budget estimates say 20t (20 metric tons IMLEO) is required for an all-propulsive landing on Mars with a spacecraft containing humans (something with an Earth mass of 20t-40t when it de-orbits at Mars). NASA's mass budget estimates say 5t to 6t using HIAD, parachutes, and retro-propulsion. I think it's a reasonable assumption that someone at NASA did their homework regarding the mass budget figures, although that is government work. I was in the government, so I know that's an assumption based on a set of contrived conditions that may not accurately reflect a practical mission profile and I could be wrong as a result. I believe that number was derived using storable chemical propellants for retro-propulsion and a RL-10 powered upper stage delivered by SLS.

I say make the parachute a bit larger in diameter, forcibly deploy it using an inflatable donut in place of the "Band" (the parachute fabric just outside the air gap used to let high pressure atmospheric gases escape so the parachute isn't ripped off whatever it's attached to) in the DGB parachute using a CO2 inflated HIAD donut, and slow yourself a bit faster from a bit further up in the atmosphere. If necessary, use SEP (Solar Electric Propulsion) to put Cygnus into a stable 3.4km/s LMO (Low Mars Orbit) vs the 7km/s reentry from interplanetary space. An extra 30 to 60 days in Mars orbit will not make or break a Mars surface mission.

At a 3.3km/s reentry velocity, you're well within HIAD's peak heating limit and you don't have to kill 7km/s in a thin atmosphere. Screaming reentries from interplanetary only assure that inordinately expensive and complicated technologies are required to soft land. The mass budget figures I use for a Cygnus and HIAD combination correlate to a HIAD designed to withstand 8km/s reentries into Earth's atmosphere. There is no written rule stating that we can't make use of every technology at our disposal to increase survivability and decrease insane performance requirements for mission critical hardware.

Strangely, the standard Cygnus module increased pressurized volume compared to BEAM and reduced the mass for structures. The HIAD for Cygnus will not subject Cygnus to substantial heating. The next HIAD test should be a standard Cygnus flown aboard a Falcon 9 with a used booster.

Cygnus is not that minimalist. I land multiple pairs of crew members in Cygnus modules with all consumables required for the trip to Mars and the surface stay. My food masses were calculated using ISS maximum permissible daily allowance per astronaut. The donut water storage tank that surrounds the astronauts sleeping quarters contains substantially more water than the minimal amount required for radiation shielding. The placement of the consumables in the module makes the module balanced and base heavy for artificial gravity and HIAD stability. The total mass is between 11t and 12t. The only time you're without a substantial consumables reserve is the short period of time when you leave Mars for Earth. Failing to achieve orbit only has one outcome and we all know what that outcome is.

We can obviously afford to throw Cygnus modules away after missions, but your ride for the entire trip is a Cygnus module. I broke the mission hardware into chunks that Falcon Heavy can throw to Mars. There's still plenty of work for rocket scientists, GW. However, there's no need to make this mission cost more than it already will.

Falcon Heavy 1 - Mars Ascent Stage
Falcon Heavy 2 - Earth Return Stage
Falcon Heavy 3 - Cygnus and Earth Departure Stage
Falcon Heavy 4 - Optional pressurized RAMA rover (a well thought-out European rover design concept)

That's $300M to $400M in launch costs for every 2 explorers and probably close to half a billion dollars total investment. A $2B investment, spread over 4 years, purchases 4 simultaneous missions to the same or different locations. There's no need to make this mission entirely unaffordable by putting all of our astronauts into a single massively over-engineered set of vehicles that are also just plain massive. A requirement for 6 or 7 $500M rockets will guarantee that this won't happen in your lifetime. Do we really want to go to Mars or talk about going to Mars?

GW Johnson · 2017-02-27 14:02:57

Hi kbd5212:

Thanks for the acronym definitions. I cannot keep track of all that stuff without some actual names now and then. Too old, I guess.

If I understand correctly, you want to use a Cygnus module as the airframe of a Mars entry vehicle, with the inflatable heat shield protecting it. Unless I am mistaken, Cygnus is made of aluminum. You'll have to protect that thing from the wake gases during entry. They'll only be high subsonic scrubbing velocities, but the effective temperatures will be around 3500 K just after the hypersonics start. You'll at the very least need a coating of Avcoat on that Cygnus. Mercury and Gemini came back bare metal like that, but they weren't aluminum. They were inconel. Although you could do it with black ceramic oven coat on SS 316. It needs to be at 1500+ F, maybe 2000 F to re-radiate effectively.

Not sure what entry trajectory you had in mind, but I did a surface-grazer from a 200 km orbit, last time I did this, some years ago. It gave me about 50 m/s delta-vee to deorbit, and a very shallow angle (1.6 degrees off local horizontal). By the time I hit entry interface at 140 km, I had an entry velocity near 3.6 km/s. From there I just did a simple 1950's-style 2-D estimate of the entry dynamics, a la H. Julian Allen. This came out at low altitude (Mach 3 = 0.7 km/s near 5 km) because I was not using an extended heat shield, just a "normal" high ballistic coefficient for a few dozen tons.

Some stuff I did recently suggests that the entry speed coming off a true Hohmann min-energy transfer is closer to 5 km/s, instead of the 6-7 km/s that is more usual now with the faster trajectories. But either way, an Earth-qualified heat shield should work fine. And could later be thinned from experience for a one-shot item.

I believe you'll need some sort of small rockets to finally touch down, almost no matter what. Even at surface densities, terminal speed is going to be rather high with any practical chute: 200-300 mph vs a survivable 20 mph. These rockets need not be very large or heavy. The Russians have landed air-dropped tanks like that long before they used the same technique to land Soyuz capsules.

For 20 mph touchdown speeds, at .007 Earth normal density, the chute loading W / CD A needs to be 0.0075 psf. CD is only a bit bigger than 1, so W/A needs to be pretty near 0.0075 psf. If W = 6 metric tons = 13,200 lb, then A must be near 1.76 million sq. ft = 163 thousand sq.m. As a single chute, that's over 450 m diameter! Over 225 m dia if a cluster of 4 chutes. It seems much more practical to me to use 3 50-meter dia chutes, and come down high subsonic at 300 mph, then fire the touchdown rockets last second.

Glad to see both you and I are looking at sending men to Mars orbit to rendezvous with pre-sent stuff there. That's the right way to do it, especially if you decide to use a reusable manned transport for that orbit-to-orbit job. Once you decide to do that, you can afford to make the living space bigger and nicer, because you can amortize the cost of the thing over multiple missions to multiple destinations, not just Mars.

GW

Last edited by GW Johnson (2017-02-27 14:06:40)

Terraformer · 2017-02-27 16:16:14

We're going back!

It's only going to be a flyby, sure, but perhaps the next one will be orbital - and then, just maybe...

SpaceNut · 2017-02-27 17:05:16

While some of this can be used, we have drifted a bit to far with discussion of parts for mars use, So when I get a moment I will either copy those posts to a new topic or add them to an existing one that matches discussion more appropriately.

GW Johnson · 2017-02-27 17:14:27

Sorry Spacenut. That was just me and kbd512 trying to clarify.

Previously in this thread, I worked out ballpark what cargo Dragon, crewed Dragon, and Red Dragon might be able to do propulsively. That was given in posts 34 and 35 above. Those data show that only a lunar flyby is possible without some sort of significantly-propulsive service module. Whether the ESA SM being done for NASA/Orion could serve a lunar orbit crewed Dragon remains to be determined.

As it now stands, Spacex cannot re-do Apollo-8, but it can do the flyby mission that really is a reprise of the lunar flyby depicted in the 1957 Disney flick. One that we never actually flew. Thread title = "Apollo-8 Redux". Crewed Dragon needs a service module of about 3.4 km/s delta-vee capability to do that mission, which leaves the capsule's 0.7-0.9 km/s capability available to do a propulsive landing on land.

If you give up the land landing capability and land with chutes at sea, then, your service module could have only 2 km/s delta-vee capability. At worst. But, you STILL need a propulsive service module to recreate Apollo-8. Just maybe not quite so capable.

Service module delta-vee capability is NOT inherent, it depends very critically upon the capsule mass attached to it. That's just rocket equation physics. Crewed Dragon is quite a bit lighter than Orion, so the corresponding delta-vee from the ESA service module would be higher. If anybody has a weight statement for the ESA service module, and an in-space engine Isp for it, I'd take a look at this issue.

Meanwhile, kudos yet again to Spacex, for taking this on.

GW

Last edited by GW Johnson (2017-02-27 17:28:28)

SpaceNut · 2017-02-27 18:09:54

No problem, I gave my scaled down version of the sizing for mars of a falcon derived dragon lander for using the same first stage retropulsion and reuse for becoming a single stage Mav that is reuseable from the start.
http://newmars.com/forums/viewtopic.php … 91#p135191
The one area that I am not totally understanding yet is the Earth departure stage sizing from the TMI for mars yet.

kbd512 · 2017-02-27 20:36:47

GW,

Perhaps I misunderstood something about the upcoming plan to test HIAD using Cygnus (standard variant, same thing I want to use now that I know it weighs less than BEAM and has more pressurized volume), but I don't believe they're planning on using a back shell. Maybe I'm wrong about that and some sort of fabric is still required to shield the top of Cygnus since the heat shield does not completely envelop the module.

Edit:
HEART - High-Energy Atmospheric Reentry Test

HEART Flight Test Overview

I scaled up my mass figures linearly for a larger unit, using the geometry information provided, to keep mass loading and peak heating within the realm of what has already been tested or projected.

I just don't want a massive retro-propulsion system using liquid fuels and all the complexities that creates. A Soyuz style retro-rocket or something like that is fine. A multi-stage abort-to-orbit Apollo 11 style lander is not. If you de-orbit, then you're committed. It's no different than landing in a Soyuz. Some things are not survivable and the astronauts on that mission need to accept that.

SpaceNut, this relates to Apollo 8 and 11 redux like so:

I want to use Cygnus for those missions, too. We'd use smaller parachutes and retrorockets to land Cygnus on Earth instead of Mars. The Mars Ascent Stage would become the Lunar Descent / Ascent Stage. The Earth Return Stage is still the same Earth Return Stage with less propellant. It's more complicated than necessary for going to the moon, but it's also a dress rehearsal for going to Mars. It's the same basic hardware, with slight modifications, to account for the unique mission requirements.

Edit:
The critical events (launch, descent, ascent, reentry, and docking) are things we wanted tested multiple times in a slightly less hazardous environment. This mission architecture caters to testing requirements by using reusable boosters, maintaining a steady launch cadence, and enabling selection of multiple exploration targets - moon, Mars, Phobos, Deimos, and asteroids.

Although it's pretty clear that everyone here wants to go back to the moon or onto Mars, maybe some other target catches our attention and we want to go there. It'd sure be nice to have a universal hardware set of comparatively inexpensive individual pieces. It's not engineering perfection or anything close to it. It's just intended to get the job done for acceptable cost and risk to the crews.

If we have another successful lunar landing using the hardware set, then the next mission should be off to Phobos and Deimos, places that are similar to the lunar environment that add long duration deep space transits. If the Phobos or Deimos mission is successful, then we've adequately demonstrated our flight chops and can proceed to land on a planet with a substantial atmosphere.

Last edited by kbd512 (2017-02-27 21:15:14)

SpaceNut · 2017-02-27 21:55:04

This is Cygnus

Which is made from the manufacturer of the ISS modules so they have the ability to custom the designs of any number of features for use of them with as many docking ports that we might need.

Such as this proposed use of it for a habitat, ignore the Orion and picture it with a Dragon V2

kbd512 · 2017-02-28 00:40:31

Two key technologies required to make space exploration a sustainable proposition, seldom addressed here except in a derisive way, underpin our future successes in space exploration. For all practical purposes, we also need practical applications of these technologies here on Earth.

1. Propulsion without reaction mass

Physicists need to stop treating general relativity, thermodynamics, electricity, and magnetism like religion. In my estimation, these domains of physics have only served to produce more questions than have been adequately answered. It's time to stop pretending that we know more than we can experimentally prove and start being more receptive to new ideas. I don't mean we should accept the results of poorly executed experiments, but reproducible data should be king in science. If data repeatedly disagrees with a theory or assertion, then eventually it's time to accept that the theory or assertion is incorrect. No actual laws that govern our universe will change to align with the personal proclivities of experimenters.

Popularity with peers is no indication of how valid a theory is, period. Obviously every experiment performed should be replicated multiple times and rigorously evaluated to attempt to eliminate experimental errors. However, data should not be thrown out based purely on disagreement with theory or assertion. The actual measurements of the speed of light and its recorded variances are a good example. Sooner rather than later, an honest attempt to explain the results is warranted.

Generating linear movement using a gyroscope is supposed to be impossible, but multiple experiments have proven that supposition false.

Faster-than-light travel is supposed to be impossible, but experiments were devised with particle accelerators whereby particles were accelerated faster than light. This is a relatively new experiment and future experiments and analysis may indicate experimental error. Unless experimental error is proven, further experiments intended to confirm or refute previously generated data is required.

Pumping radio frequency energy into an enclosed metal cavity should not generate thrust, yet every experiment conducted to date indicates that it does. Any scientists worth their salt should be trying to determine how and why this happened by devising their own experiments, not simply claiming that the experimenters with the data produced erroneous results when they have no evidence at all to support claims of experimental error, one way or the other.

There's more heat energy trapped in the Earth's atmosphere than all the liquid hydrocarbons, coal, or nuclear power could ever reasonably produce. It's there for the taking, every day of the year, even in Antarctica. Astonishingly, we're not using simple and reliable heat pumps to produce electrical power from the most ridiculously powerful source available that virtually any human on the planet can access. There was never any requirement to use liquid hydrocarbons, coal, or nuclear power for electrical power production. A fusion reactor look like a curiosity for children, if total potential output from both systems is taken into consideration.

2. Advanced electrical power generation systems

All spacecraft built and flown to date have minimal means to produce electrical power, but future operations requiring humans to live and work in space, in appreciable numbers or for appreciable duration, will mandate highly compact, light weight, and power dense sources to sustain operations.

In decades past, solar panels, fuel cells, and fission reactors were the only options available. In the next decade, so-called low-energy or lattice-enabled nuclear reactions will rapidly replace all of those technologies as the go-to power source for human space exploration. It's economically infeasible to transport sufficient quantities of solar panels or the associated batteries required for electrical power storage to other planets, the best fuel cells available lack sufficient energy density for sustainment of operations, and fission reactors have come to symbolize nuclear weapons to a citizenry who are even more ignorant of how nuclear reactors work than they are about how their motor vehicles work.

The two technologies described were and are the most implacable impediments we have in our quest to take our place amongst the stars. When these two problems are adequately resolved, there are few practical limits to what we can achieve within our own solar system.

Terraformer · 2017-02-28 05:44:39

kdb512,

Though it's pretty obvious that the currently known "laws" of physics - thermodynamics, the light speed barrier, conservation laws - don't always hold true (the existence of the universe tells us that!), I'm fairly sure we don't yet have the ability to violate them...

elderflower · 2017-02-28 07:06:31

I don't know about the rest of you, but I'm pretty sure that I don't have time to wait for somebody to develop things that are not already demonstrated or just a short time off that.

Oldfart1939 · 2017-02-28 08:52:45

When I was just a kid back in 1953, it was when the first popularized space articles had been appearing in Collier's magazine, it was only 50 years since the Wright brothers started riding on, essentially, powered kites; at that time jets were coming into service that were supersonic. Only 10 years after that, men were riding rockets into LEO, and another 10 years later, men had walked on the Moon. In my college years, I was positive we'd be on Mars by 2000, and the asteroids by 2010. What went wrong? Answer: we lacked not the intelligence and science to do so--only the POLITICAL WILL to do so. No President since Kennedy has come forward with SCIENCE and SPACE as high national priorities. Maybe--just maybe--this new one will have a different perspective. I'm not "counting on it," but...hoping!

GW Johnson · 2017-02-28 12:06:14

This is for kbd512 concerning the Cygnus experiment with the inflatable heat shield:

I took a look around on the internet to try to pin down what Cygnus's external shell is made of. I couldn't find anything, but it surely does look like aluminum in the photos. It does ride up inside a payload shroud to protect it from M2-3 aeroheating below ~125,000 feet on the way up. Bare aluminum cannot survive that, either. Most aluminum alloys are crap at about 300-350 F material soakout temperatures, including the latest-generation aluminum-lithium alloys.

The paper you linked has a good analysis of the inflatable heat shield flight test experiment, but the guys planning that were obviously not looking at wake gas heating to the Cygnus itself. A lot of the time in papers like that, the other issues are "somebody else's problem". If such a project is really funded, someone somewhere must address this issue.

I'd hazard the guess that you could just wrap the thing in a ceramic blanket and bluff your way through this. That's a fairly lightweight solution, used on the later shuttles in lieu of some of the backside tiles. The only "trick" is TYPING {edit: not "typing", "tying"; I really hate autocorrect; my slide rule NEVER EVER misbehaved like this!!!!} tying it in place.

GW

Last edited by GW Johnson (2017-02-28 18:33:33)

SpaceNut · 2017-02-28 21:33:37

Modify the payload shroud flairing to stay as part of the unit and problem is taken care of for heatshielding with it being coated.

.

kbd512 · 2017-03-01 02:24:03

SpaceNut,

That would probably work but instead of modifying the payload shroud, and the tens of millions that could potentially cost for a single type of payload, why not just wrap the module in a protective fabric like the Russians do?

Cygnus is definitely not what comes to mind when I think of an interplanetary spacecraft and lander, but it's a lot closer to the Chicken of the Sea model than Dr. Zubrin's tuna can. It's the lightest pressurized module in the inventory with enough pressurized volume to deliver two people and consumables to the moon or Mars.

There are any number of purpose-built solutions that should work better than Cygnus, but at what cost and how long would it take to develop it? If Cygnus can survive a launch and long duration missions in space, then it can probably survive a landing, too. I'm trying to work with what we have right now that fulfills requirements.

Dragon can land on the moon or Mars, but it's not coming back since it's far too heavy. Orion can't land anywhere but an ocean here on Earth. Soyuz can't land anywhere but Earth. Dream Chaser can't land anywhere but a runway on Earth. That narrows the selection criteria a bit.

GW,

Incidentally, you are correct about the requirement for a back shell to protect Cygnus. I hadn't thought about that because I thought the HIAD engineers wanted Cygnus to remain usable after reentry. Obviously they weren't too concerned about that. How much will this fabric cost in terms of added mass? My mass margins for TMI aren't that great. I have less than 2t to play with.

MMOD 100% Al
MLI 15.3% Beta Cloth, 58.8% MLI, 7% Kevlar, 18.8% Nextel
Primary Structure 93.9% Al, 6.08% Steel
Secondary Structure 73.5% Al, 14.1% Zylon, 11.1% Zotek, 1.5% Steel
Internal Components 80% Al, 10% Cu, 10% Kapton

Spacecraft SPE Shielding SPE environment models and materials performance in a Cygnus PCM based EAM

Oldfart1939 · 2017-03-01 08:49:02

Before I make my engineering and science comments; SpaceNut: shouldn't this discussion be moved to a Mars thread, as inflatables for Mars entry need to be on a Mars thread? Feel free to delete this portion of this post, need arising.

What's really obvious underlying these discussions is just how poorly engineers have done w/r the Orion capsule. It is too limited in scope to be useful for anything other than just a joy ride in space. Terribly overweight and over budget. It reminds me of the old joke here in the west: a camel is a horse designed by a committee; modified form pertaining to Orion: An elephant is a racehorse designed by NASA engineers.

kbd512 · 2017-03-01 10:38:47

Oldfart1939,

We can use a Mars mission architecture for lunar, Mars, and asteroid missions if the architecture is not optimized for any of the various exploration targets selected.

I use 3 Falcon Heavy rockets for lunar and Mars missions. The Earth Return Stage performs TEI / EOI (TEI and velocity reduction only for Mars missions) and the combination Descent / Ascent Stage (only an Ascent Stage on Mars missions) performs both descent and ascent for Cygnus on the moon.

FH1 - Earth Return Stage is delivered to the exploration target ahead of the crew
FH2 - Descent / Asent Stage for lunar and Ascent Stage only for Mars missions is delivered to the exploration target ahead of the crew
FH3 - Cygnus module is a miniature habitat module, used from mission start to completion, for both lunar and Mars missions

Dr. Zubrin proposed a similar architecture for Mars Direct, but his mission components required further modifications for use on the moon because they were really designed for Mars. I'm going for plug-n-play here. The kick stages all use pump-fed NTO/MMH engines. The Earth Return Stage for Mars missions also carries a second HIAD for Earth Return of Cygnus. For lunar missions, Cygnus disconnects from HIAD and connects to the Descent / Ascent Stage.

Is any of this optimal for Apollo 8 / 11 redux? Absolutely not. Can it be made to work for the varied exploration targets without designing entirely new vehicles? Yes. It's within the mass constraints of what Falcon Heavy can deliver and if there is ever any reuse of boosters for the delivery of kick stages, then it becomes far less expensive than using SLS to accomplish the same task, even if the mission is more complicated as a result of the orbital rendezvous requirements.

Break the mission hardware set into roughly equal chunks that a comparatively inexpensive heavy lift rocket like Falcon Heavy or Vulcan Heavy can throw, make it truly modular, focus on rocket reusability rather than spacecraft reusability, and exploration of all realistic exploration targets we can reach with current technology are well within the realm of economic and technical feasibility with current budgets.

We can use this approach (doesn't have to be specifically what I proposed, it just needs to use existing hardware that can get the job done) or we can wait ten years for development of a purpose-built deep space habitat, ten years for a purpose-built lunar lander, and another ten years for a purpose-built Mars lander. It'll be 2030 before we go back to the moon and 2040 before we go to Mars using the purpose-buiilt approach to hardware development. The approach I proposed could be flight rated hardware in 6 years or less. Want vs need. Requirement vs nice to have.

Oldfart1939 · 2017-03-01 11:07:45

kbd512-

I'm not disagreeing with your overall mission concept--only stating that it seems misplaced here and lost in the shuffle.

I have proposed an architecture that also requires multiple Falcon Heavy launches. Overlooked in the shuffle sometimes is the simple fact that RP-1 fuelled vehicles cannot be used on long missions; RP-1 has terrible thermal properties, in that it gels up like Jello below minus 35 degrees C. We need to utilize a new 2nd stage--one using the methylox couple. That is subsequently completely compatible with a Mars landing and the Zubrin proposed Sabatier reaction for fuel production. See my model for a mission architecture on the Mars Direct, Mars Semidirect, etc. thread.

Last edited by Oldfart1939 (2017-03-01 11:39:18)

kbd512 · 2017-03-01 16:36:42

Oldfart1939,

I agree that RP-1 is not a good propellant choice for long duration missions. What does that have to do with going to the moon? The RP-1 in the Falcon Heavy upper stage will be expended in a matter of minutes to hours. To go to the moon (or Mars), you use storable propellants like NTO/MMH to land and come back, just like we did during the Apollo program.

This LOX/LCH4 fixation was Dr. Zubrin's answer to the requirement to get back to Earth from the surface of Mars using a two stage vehicle propulsively landed on Mars. There's no need for that for any lunar or Mars mission. We have too much experience with orbital rendezvous and assembly to mess around with new rocket engines and propellants. There are plenty of other ways to spend money that will be of greater benefit to human space exploration.

RobertDyck · 2017-03-01 17:20:06

N2O4 is toxic. NASA found out just how toxic at the end of the Apollo-Soyuz mission. The TV show "Secret Space Escapes" recently featured that incident. During atmospheric entry, something electronic failed in the Command module resulting in a loud sequel in the headphones of all astronauts. When the mission commander ordered the RCS thrusters disabled, the astronaut who was supposed to flick the switch didn't hear him, so didn't. When the drogue shoots deployed, the Command module started to sway. The RCS thrusters attempted to stop movement, so they fired. But as the capsule opened a value to allow air in, to equalize pressure from the low pressure that Apollo operates in space to ambient air pressure. The intake brought in N2O4 from RCS exhaust. The Command module filled with toxic vapour. Astronauts coughed and suffered toxic effects of N2O4. The mission commander flicked the switch to deploy main chutes at 10,000 feet. When the Command module hit the water, it hit a wave flat, causing high-G acceleration. It slapped the wave and bounced over, ending up floating up-side-down. Astronauts hung in their seat harness. Astronauts continued to suffer toxic effects, the commander released his harness so he could get to O2 bottles. He fell onto the control panel. He gave all astronauts their O2. One astronaut's oxygen mask had shifted to the side, he wasn't breathing O2. He passed-out, unconscious. The commander moved his mask back onto his nose/mouth, but as soon as that astronaut became conscious he thrashed with his arms and punched the commander in the face. And he shifted the O2 mask off his face again. Passed-out again. The mission commander gave that astronaut a bear hug, while the third astronaut shifted his oxygen mask back onto his face. The unconscious astronaut came to again, and again tried to thrash about with his arms, but the commander did not let his bear-hug go. When the astronaut breathed enough O2, he was able to think clearly enough that he stopped trying to push the O2 mask off his own face, and stopped trying to punch his commander. At that point the commander could let go. At this point the flotation bags inflated, turning the Command module upright. Then navy frogmen (as they were called at that time) arrived to put the flotation skirt around the command module. They gave a brief speech when the arrived on the carrier, but were taken directly to sickbay. The navy surgeon said they were all suffering from toxic induced pneumonia.

The Space Shuttle had to be secured, all N2O4 and MMH secured before they could open the hatch to the mid-deck so astronauts could walk out. Now you know why.

Oldfart1939 · 2017-03-01 17:32:30

N2O4 is toxic and corrosive. Monomethylhydrazine (MMH) is "only" poisonous, and exposure usually results in severe liver damage. The methlox couple is favored now because it has an excellent Isp, and in the Raptor engine is hypothecated to deliver 383 sec. This is second only toH2/LOX. Fewer problems with Methylox than H2/LOX.

kbd512 · 2017-03-01 21:46:47

Oldfart1939,

I agree that N2O4 and MMH are toxic, but they're also storable and available right now, whereas cryogens are not without insulation, cryocoolers, and continuous power to the cryocoolers. I'm not opposed to using HAN (HydroxylAmmonium Nitrate - AF-M315E formulation), instead of N2H4 (Hydrazine), for RCS. AF-M315E has 50% greater density impulse than hydrazine, higher specific impulse compared than hydrazine, it's non-toxic, and can be re-heated to liquefy it if it freezes so no tank insulation is required. If H2O2 (hydrogen peroxide) and HAN can deliver similar Isp to N2O4 / MMH, then we have a storable bi-propellant with fantastic density impulse.

Cryogen storage means more weight, more complexity, and more cost for no discernible benefit, apart from somewhat less hazardous handling qualities which is dubious since LOX is involved. Most cryogen tanks also leak and since Rob has done a bit of reading about the STS program, he must be aware that this is not conjecture on my part.

Rob,

There won't be any fuel aboard Cygnus after it reenters so I fail to see how orbiter processing procedures or horror stories from the Apollo era relate to opening the hatch on a Cygnus after the astronauts return to Earth. ASTP and STS are poor comparisons since Apollo capsules and STS orbiters have their RCS system integrated into the reentry vehicle, whereas Cygnus has a separate service module.

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#51 2017-02-25 15:34:57

Re: Apollo 8, redux

#52 2017-02-25 17:00:21

Re: Apollo 8, redux

#53 2017-02-25 21:07:15

Re: Apollo 8, redux

#54 2017-02-27 11:33:44

Re: Apollo 8, redux

#55 2017-02-27 14:02:57

Re: Apollo 8, redux

#56 2017-02-27 16:16:14

Re: Apollo 8, redux

#57 2017-02-27 17:05:16

Re: Apollo 8, redux

#58 2017-02-27 17:14:27

Re: Apollo 8, redux

#59 2017-02-27 18:09:54

Re: Apollo 8, redux

#60 2017-02-27 20:36:47

Re: Apollo 8, redux

#61 2017-02-27 21:55:04

Re: Apollo 8, redux

#62 2017-02-28 00:40:31

Re: Apollo 8, redux

#63 2017-02-28 05:44:39

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#65 2017-02-28 08:52:45

Re: Apollo 8, redux

#66 2017-02-28 12:06:14

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#67 2017-02-28 21:33:37

Re: Apollo 8, redux

#68 2017-03-01 02:24:03

Re: Apollo 8, redux

#69 2017-03-01 08:49:02

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#71 2017-03-01 11:07:45

Re: Apollo 8, redux

#72 2017-03-01 16:36:42

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#73 2017-03-01 17:20:06

Re: Apollo 8, redux

#74 2017-03-01 17:32:30

Re: Apollo 8, redux

#75 2017-03-01 21:46:47

Re: Apollo 8, redux

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