NASA still investigating Orion heat shield issues from Artemis 1 moon mission

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I don't understand why they have to come in so hot (40,000 kph). Why must they enter the atmosphere at this extreme velocity?
Could they not slow down the vehicle before re-entry and thus avoid the extreme heat effects? In fact, if they could sufficiently reduce velocity, they could use parachutes to re-enter the atmosphere, yes?
 
I don't understand why they have to come in so hot (40,000 kph). Why must they enter the atmosphere at this extreme velocity?
Could they not slow down the vehicle before re-entry and thus avoid the extreme heat effects? In fact, if they could sufficiently reduce velocity, they could use parachutes to re-enter the atmosphere, yes?
I don't know anything about the specifics of this case but I can make a guess. If they are coming in real fast, it is because they didn't launch enough fuel to be able to slow down at the right time. Adding fuel makes everything bigger and bigger and vastly more expensive. The faster we can learn to come in, the more we can do with what we have.
 
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I don't know anything about the specifics of this case but I can make a guess. If they are coming in real fast, it is because they didn't launch enough fuel to be able to slow down at the right time. Adding fuel makes everything bigger and bigger and vastly more expensive. The faster we can learn to come in, the more we can do with what we have.
yes, this is an answer from an a.i.:
"The reason spacecraft like Apollo or Orion enter Earth's atmosphere at such high speeds is due to the laws of orbital mechanics.
When a spacecraft is orbiting the Earth, it is essentially in a constant state of freefall, balanced between its velocity and the pull of Earth's gravity. To return to Earth, the spacecraft must slow down enough to drop out of orbit and begin descending towards the surface.
However, slowing down in space is not easy. In the vacuum of space, there is no air resistance to naturally slow the spacecraft. The only way to decelerate is to use thrusters or engines, which requires carrying a lot of heavy fuel. It's much more fuel-efficient to let Earth's atmosphere do most of the work of slowing the spacecraft down.
As the spacecraft enters the denser layers of the atmosphere at high speed (around 17,000 mph for Apollo), the air resistance creates intense friction and heat, up to 5000°F or more. This is why the spacecraft needs a heat shield to protect it and its occupants during re-entry.
Theoretically, a spacecraft could use its engines to slow down before re-entry to reduce the heat. But this would require carrying much more fuel, greatly increasing the spacecraft's size and weight, and making the entire mission more expensive and complex. Using Earth's atmosphere to decelerate is a more practical approach.
So in summary, the high re-entry speeds are a consequence of the spacecraft's orbital velocity, and using the atmosphere to slow down, while very hot, is more efficient than carrying extra fuel to decelerate before re-entry. The heat shields and re-entry profiles are designed to manage these extreme but unavoidable conditions of returning from orbit to Earth's surface."
 
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This I believe has influenced SpaceX so much, The need to slow down before re-entry.
To that end the Starship V3 is growing in size. Thus allowing a Starship variant to refuel in orbit , that is then able to get to the Moon and back and have enough fuel to slow down to enter LEO where it can either dock with a lander or if the additional weight of a full heatshield is possible, Then it can the use the fact the Steel body that can withstand the re-entry temperatures.
 
It is possible Apollo heat shield Avcoat behaved the same, there is no mentioning of an effect comparison.

yes, this is an answer from an a.i.:
"The reason spacecraft like Apollo or Orion enter Earth's atmosphere at such high speeds is due to the laws of orbital mechanics.
While it is quite all right to use large language models to get some insight into a new problem - and importantly, to declare it as here - they are known to hallucinate and they won't give proper references. Think of it as a "popular science" take, but check its contribution.

Yes, it is a problem of fuel consumption. Earth is so massive that it is barely possible to launch chemically propelled rockets to orbit and even less to escape velocities, since the rocket equation means there is an exponential increase in fuel consumption with achieved speed. An orbital launcher has a few percent net cargo mass, typically in the 3-5 percent range. (This is why it is so expensive to send a kg of cargo to orbit, compared to an Earth destination.) Hence missions tend to use minimal propellant energy free ("Hohmann") orbits.

So the higher the orbit, the higher the return energy:
The atmospheric entry interface velocity upon return from the Moon is approximately 36,500 ft/s (11.1 km/s; 40,100 km/h; 24,900 mph)[4] whereas the more common spacecraft return velocity from low Earth orbit (LEO) is approximately 7.8 km/s (28,000 km/h; 17,000 mph).
https://en.wikipedia.org/wiki/Free-return_trajectory

The Hohmann maneuver often uses the lowest possible amount of impulse (which consumes a proportional amount of delta-v, and hence propellant) to accomplish the transfer, but requires a relatively longer travel time than higher-impulse transfers.
https://en.wikipedia.org/wiki/Hohmann_transfer_orbit

A more detailed analysis: https://www.faa.gov/sites/faa.gov/f...ffices/avs/III.4.1.7_Returning_from_Space.pdf

No, it is not possible to reentry from a 8 km/s orbital speed with a parachute designed for low atmosphere falls at ~ 0.2 km/s for a human in terminal speed fall (fall with aerodynamic lift steady state). The best you can do is a kind of heatshield protected foam bubble.
MOOSE, originally an acronym for Man Out Of Space Easiest but later changed to the more professional-sounding Manned Orbital Operations Safety Equipment,[1] was a proposed emergency "bail-out" system capable of bringing a single astronaut safely down from Earth orbit to the planet's surface.[2][3] The design was proposed by General Electric in the early 1960s. The system was quite compact, weighing 200 lb (91 kg) and fitting inside a suitcase-sized container. It consisted of a small twin-nozzle rocket motor sufficient to deorbit the astronaut, a PET film bag 6 ft (1.8 m) long with a flexible 0.25 in (6.4 mm) ablative heat shield on the back, two pressurized canisters to fill it with polyurethane foam, a parachute, radio equipment and a survival kit.
https://en.wikipedia.org/wiki/MOOSE

There are ways to reduce high orbital speeds towards low Earth entry speeds though. Besides using gravity assist braking you can try to double up with aerobraking in the case of Earth (or even Mars).

Gravity assistance can be used to accelerate a spacecraft, that is, to increase or decrease its speed or redirect its path.
https://en.wikipedia.org/wiki/Gravity_assist

Aerobraking is used when a spacecraft requires a low orbit after arriving at a body with an atmosphere, as it requires less fuel than using propulsion to slow down.
https://en.wikipedia.org/wiki/Aerobraking

They are somewhat risky methods, especially the latter as it relies on a complex and dynamic atmospheric pressure*, which is why e.g. SpaceX do not (yet, officially) rely on them for its projected Mars missions:
no spacecraft can yet aerobrake safely on its own

*) Think of how solar emissions heightened Earth stratospheric height and drag so a launch of Starlinks failed to reach orbit.
 
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To that end the Starship V3 is growing in size. Thus allowing a Starship variant to refuel in orbit , that is then able to get to the Moon and back and have enough fuel to slow down to enter LEO where it can either dock with a lander or if the additional weight of a full heatshield is possible,
Starship system design is primarily driven by its eventual interplanetary mission capabilities, not by NASA et al. Moon shenanigans. The Moon lander will be dedicated to travel to Moon to be used as a local non-heat shield ferry there, with Orion and its heatshield as the crew craft taking the astronauts to and from lunar orbit. Yes, it will need to be refueled in Earth orbit to reach lunar orbit, and later refueled in lunar orbit - and have a crew environmental control system - so it shares some deliverables with other manned Starship projects.

[Though the last week there has started to come out commentaries to the effect that NASA has realized that the Apollo project design is the safest and cheapest. E.g. they may slow down the project and decrease its per mission cost by testing out the Orion/Starship Moon lander docking interface and crew handling in a safer Earth orbit as the 3d mission, instead of trying it out in lunar orbit. If they would be adamant on "the less parts, the less risks (and costs)" they would eventually ditch Orion, skip the lunar orbital station, and rely on lunar modified but heat shielded Starships entirely.]

The 100 mt Moon cargo capability comes without a heatshield (but with Moon land and reorbit propellant). https://en.wikipedia.org/wiki/Starship_HLS No doubt they will have refrigerators and microwaves for "Moon Cordon bleu development" and astronaut R&R. Think of Earth food chains and restaurants that announce sodas and menus that "will take you to the moon"!
 
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I don't understand why they have to come in so hot (40,000 kph). Why must they enter the atmosphere at this extreme velocity?
Could they not slow down the vehicle before re-entry and thus avoid the extreme heat effects? In fact, if they could sufficiently reduce velocity, they could use parachutes to re-enter the atmosphere, yes?
Yes, in theory, they could enter the atmosphere at low speed. But it takes a lot of energy to slow down. In fact it takes as much energy to reduce speed as it took to gain all that speed. SO they would literally need a huge rocket like the one that sent them to the moon and of course all the fuel too.

As it turns out heat shields are MUCH lighter than big rockets and fuel tanks. The heart sheild converts all that kinetic energy to heat and has no moving parts to fail.

Look at how much harder it is to land on the Moon, they need a decent rocket. It is only possible to land on the moon becuase the gravity is so low. Landing one Earth using a lunar-landing-like rocket would be nearly impossible.

Mars is an intermediate problem. They can use a heat shield but they also need rockets and a parachute. Mars has a thin atmosphere and 38% of Earth's gravity.

Venus was the best, heat shields are very effective in the thick atmosphere. As we saw with Pinoneer Venus it is possible to crash land on Venus with nothing but a heat sheild. The air is so thick that the velocity of freefall is low-enough for a survivable crash.

Heat sheilds are the best and most cost effective why to slow a spacecraft for landing
 
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Hello readers!
I can understand the lethargy from the Orion team as can be seen from lack of innovation in many areas by their just copying items, shapes, technologies and materials largely from Apollo. I worked on Apollo 8-17 and also on planning for Space Shuttle and Space Station as early as 1968-1972.

We improved several new items that have proven much more resilience to extreme environment.

While Skylab except for solar panel damage used some prototypes, Shuttle and Station have more robust systems compared to Apollo.

Columbia and Challenger disasters were largely human errors and some of my suggestions were identical to those implemented for post Columbia post-disaster missions for a decade. In my opinion there was need to be more vigilant and we could have continued after 2013 till SpaceX could be ready.

Anyway that is history.

But just as Shuttle discovered materials for low level ablative losses and a lot is happening now in Space Force arena, we could expect SpaceX and Starship technologies to use better materials than Apollo vintage,

In my opinion not enough priority has been give to reentry materials and into preflight robotic missions to prove newer concepts.

Thanks.
Ravi
(Dr. Ravi Sharma, Ph.D. USA)
NASA Apollo Achievement Award
ISRO Distinguished Service Awards
Former MTS NASA HQ MSEB Apollo
Former Scientific Secretary ISRO HQ
Ontolog Board of Trustees
Particle and Space Physics
Senior Enterprise Architect
SAE Fuel Cell Tech Committee voting member for 20 years.
http://www.linkedin.com/in/drravisharma
.
 
Hello readers!

I can understand the lethargy from the Orion team as can be seen from lack of innovation in many areas by their just copying items, shapes, technologies and materials largely from Apollo. I worked on Apollo 8-17 and also on planning for Space Shuttle and Space Station as early as 1968-1972.

We improved several new items that have proven much more resilience to extreme environment.

While Skylab except for solar panel damage used some prototypes, Shuttle and Station have more robust systems compared to Apollo.

Columbia and Challenger disasters were largely human errors and some of my suggestions were identical to those implemented for post Columbia post-disaster missions for a decade. In my opinion there was need to be more vigilant and we could have continued after 2013 till SpaceX could be ready.

Anyway that is history.

But just as Shuttle discovered materials for low level ablative losses and a lot is happening now in Space Force arena, we could expect SpaceX and Starship technologies to use better materials than Apollo vintage,

In my opinion not enough priority has been give to reentry materials and into preflight robotic missions to prove newer concepts.
Ravi
(Dr. Ravi Sharma, Ph.D. USA)
NASA Apollo Achievement Award
ISRO Distinguished Service Awards
Former MTS NASA HQ MSEB Apollo
Former Scientific Secretary ISRO HQ
Ontolog Board of Trustees
Particle and Space Physics
Senior Enterprise Architect
SAE Fuel Cell Tech Committee voting member for 20 years.
http://www.linkedin.com/in/drravisharma
 
Scott

Interesting brain teaser. Thanks for being Knowledgeable on human capable spacecraft and Launch Vehicles! Enjoyable!

As you mentioned, I also worked on space tug although less duration compared to Shuttle and Station design Phase B studies in parallel to working on Skylab and Apollo missions 8-17.
I also studied thermal loads on spacecraft such as the Shuttle from Earth and Sun and the effects on such for different inclined orbits as a Principal Investigator for NASA HQ funded study.
Space tug was not actually built but today if studied again in the light of Artemis, it would be relevant for cislunar needs and for gateway build up at the Lunar orbit.
Your booster reuse for fueling is still a good potential option with more studies needed.
Best wishes.
Ravi
(Dr. Ravi Sharma, Ph.D. USA)
NASA Apollo Achievement Award
ISRO Distinguished Service Awards
Former MTS NASA HQ MSEB Apollo
Former Scientific Secretary ISRO HQ
Ontolog Board of Trustees
Particle and Space Physics
Senior Enterprise Architect
SAE Fuel Cell Tech Committee voting member for 20 years.
http://www.linkedin.com/in/drravisharma
 
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NASA is still studying the performance of the Orion capsule's heat shield during its reentry to Earth's atmosphere at the end of the Artemis 1 moon mission in late 2022.

NASA still investigating Orion heat shield issues from Artemis 1 moon mission : Read more

A KINETIC ENERGY CONVERTER (KEC)
When material science improves, Starships returning to Earth will have a different heat dissipation. One that is efficient and simple.
My suggestion is an additional outer skin on each section. A ten mm clearance is required to house thousands of nine mm long radial spikes pointing inwards, thus magnifying heat dissipation.
Superheated steam can be ejected through external venturis to add thrust against the atmosphere. Heat tiles can be minimized because the best part is no part.

PS This KEC could be used on the Orion. Would someone do the mathematics? KE[1/2mv(2)] = Latent Heat of Super Heated Steam. This leads to kgs of water required to replace the heat tiles.
 
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Orion crew capsule weighs 8500 kg.
Reentry speed is 40,000 km/h or 11,100 m/s.
Kinetic energy is thus 1/2 * 8500 * 11,100^2 or 5.2e11 J.
The heat of fusion of water is 333,000 J/kg.
The amount of water needed is thus 1.6e6 kg.
This would increase the capsule mass by almost 200 times.
This is probably not going to work.
 
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Bill, I agree that it probably won't work.

But you did not take into account Wisetoe's proposed use of the steam created to produce retro thrust to slow down the reentering craft. And, the total kinetic energy does not have to be counteracted, just enough to get to the Q value that can be accommodated by specially designed parachutes before the reentry craft's surface skin is damaged by heat of friction during reentry.

So, the performance calculations are much more complex than a simple calculation of how much water is needed to absorb all of the kinetic energy as latent heat of vaporization by water.
 
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It is possible Apollo heat shield Avcoat behaved the same, there is no mentioning of an effect comparison.
thank you for your classic style reserach! Torbjorn, if you can, could you tell me something about the story dating back to the dawn of russian and american space enterprises, when theorists of trajectory calculations were unable to obtain launches that corresponded to the laws of propulsion, ballistics and orbital mechanics, and had to introduce into equations some ad hoc terms - still used today and never explained - in order to correctly design the mission targets? thx very much
 
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Bill, As I posted already, my thinking is that you are correct that this would not be a useful design.

But, that is just the gut feeling of us two who have calculated such things for other ideas, not this one. Wisetoe asked for the simple calculation that you performed for him. My comment is simply that the calculation asked and given is not really what is needed to address the proposal's effectiveness, because it does not include the beneficial aspects I delineated. But I do NOT think those aspects can overcome the 200:1 weight difference. Here is why:

Based on https://www.nas.nasa.gov/pubs/stories/2022/feature_LAVA_Parachutes.html , I'll guess that a parachute could be deployed at about 1,500 mph = 670 m/sec. So, the total velocity change still would need to be about 10,000 m/sec. To accomplish that ONLY with steam emitted from the cooling jacked, and limiting the total amount of steam to the weight of the current ablative heat shield, about 4000 lbs = 1,800 kg (see https://www.space.com/22036-nasa-orion-spacecraft-heat-shield-photos.html), that steam would need to be emitted out forward facing nozzles at something like (10,000 m/sec x 8,500 kg / 1,800 kg =) 47,000 m/sec. But, the exhaust velocity of the Hydrogen/Lox powered J2 rocket engines (which emit superheated steam) is only about 1/10 of that. See https://en.wikipedia.org/wiki/Rocketdyne_J-2 .

Worse, that J2 performance is when the gas is going aft, not forward, and the back pressure on the nozzle is nearly the vacuum of space. The chamber pressure of the J2 engine is only 763 psi. The effective back pressure on a nozzle facing forward is dependent on its velocity into the air and the density of the air. I did a quick 'n' dirty scaling of some numbers I won't try to explain here, and it looks like the "max Q" for the Orion capsule reentry could get close to the J2 motor chamber pressure. So, the thrust would be very small when the heating was very large, unless the temperatures and pressures inside the cooling jacket far exceed what we can handle in a H2/Lox rocket motor.

Further, the surface temperature of the reentry craft would reach about 5,000 degrees F if it is not cooled, but the metals that could be used would need to be kept to more like 2,000 degrees or less. So, if it gets to the point that there is not much retro thrust, the cooling is rate is going to be very small until the pressure and temperature inside the cooling jacket get very high. So, now you are talking about making a rocket motor with a diameter the size of Orion capsule diameter out of super alloys, and make it far stronger than the J2 rocket motor. That in itself would have to be very heavy.

Exactly how all of this would play out in terms of design requirements would require a lot of internal and external thermodynamic and aerodynamic calculations integrated over the entire descent path. Obviously, I am not going to try to do that myself. But, I think the issues I raise here should be sufficient to convince folks that this is not likely to be a feasible substitute for a heat shield.
 
I have read and heard that high voltage and current potentials are generated from trailing wires in orbit. Some in a surprising violent way. If we could tame it and store it we might have a great orbit electric power source. With a retractable collector. And if we could polarize the returning craft, we might step accelerate it down, like locks on a canal, with magnetic rings, placed at different levels.

Faraday Reduction.
 
Yes, a craft in orbit can extend a wire such that it cuts the Earth's magnetic field and generates power. The energy comes from the orbital energy of the craft. This could be used to deorbit a craft but is not a source of power.
 
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Orion crew capsule weighs 8500 kg.
Reentry speed is 40,000 km/h or 11,100 m/s.
Kinetic energy is thus 1/2 * 8500 * 11,100^2 or 5.2e11 J.
The heat of fusion of water is 333,000 J/kg.
The amount of water needed is thus 1.6e6 kg.
This would increase the capsule mass by almost 200 times.
This is probably not going to work.
Really appreciate your interest to do the mathematics for me.
 
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Bill, As I posted already, my thinking is that you are correct that this would not be a useful design.

But, that is just the gut feeling of us two who have calculated such things for other ideas, not this one. Wisetoe asked for the simple calculation that you performed for him. My comment is simply that the calculation asked and given is not really what is needed to address the proposal's effectiveness, because it does not include the beneficial aspects I delineated. But I do NOT think those aspects can overcome the 200:1 weight difference. Here is why:

Based on https://www.nas.nasa.gov/pubs/stories/2022/feature_LAVA_Parachutes.html , I'll guess that a parachute could be deployed at about 1,500 mph = 670 m/sec. So, the total velocity change still would need to be about 10,000 m/sec. To accomplish that ONLY with steam emitted from the cooling jacked, and limiting the total amount of steam to the weight of the current ablative heat shield, about 4000 lbs = 1,800 kg (see https://www.space.com/22036-nasa-orion-spacecraft-heat-shield-photos.html), that steam would need to be emitted out forward facing nozzles at something like (10,000 m/sec x 8,500 kg / 1,800 kg =) 47,000 m/sec. But, the exhaust velocity of the Hydrogen/Lox powered J2 rocket engines (which emit superheated steam) is only about 1/10 of that. See https://en.wikipedia.org/wiki/Rocketdyne_J-2 .

Worse, that J2 performance is when the gas is going aft, not forward, and the back pressure on the nozzle is nearly the vacuum of space. The chamber pressure of the J2 engine is only 763 psi. The effective back pressure on a nozzle facing forward is dependent on its velocity into the air and the density of the air. I did a quick 'n' dirty scaling of some numbers I won't try to explain here, and it looks like the "max Q" for the Orion capsule reentry could get close to the J2 motor chamber pressure. So, the thrust would be very small when the heating was very large, unless the temperatures and pressures inside the cooling jacket far exceed what we can handle in a H2/Lox rocket motor.

Further, the surface temperature of the reentry craft would reach about 5,000 degrees F if it is not cooled, but the metals that could be used would need to be kept to more like 2,000 degrees or less. So, if it gets to the point that there is not much retro thrust, the cooling is rate is going to be very small until the pressure and temperature inside the cooling jacket get very high. So, now you are talking about making a rocket motor with a diameter the size of Orion capsule diameter out of super alloys, and make it far stronger than the J2 rocket motor. That in itself would have to be very heavy.

Exactly how all of this would play out in terms of design requirements would require a lot of internal and external thermodynamic and aerodynamic calculations integrated over the entire descent path. Obviously, I am not going to try to do that myself. But, I think the issues I raise here should be sufficient to convince folks that this is not likely to be a feasible substitute for a heat shield.
Really appreciate your interest to do the mathematics for me. Wisetoe
 
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A KINETIC ENERGY CONVERTER (KEC)
When material science improves, Starships returning to Earth will have a different heat dissipation. One that is efficient and simple.
My suggestion is an additional outer skin on each section. A ten mm clearance is required to house thousands of nine mm long radial spikes pointing inwards, thus magnifying heat dissipation.
Superheated steam can be ejected through external venturis to add thrust against the atmosphere. Heat tiles can be minimized because the best part is no part.

PS This KEC could be used on the Orion. Would someone do the mathematics? KE[1/2mv(2)] = Latent Heat of Super Heated Steam. This leads to kgs of water required to replace the heat tiles.
Really appreciate your interest to do the mathematics for me.
 
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