Author Topic: Mars EDL technologies  (Read 175801 times)

Offline Warren Platts

Re: Mars EDL technologies
« Reply #180 on: 02/01/2011 06:50 am »
Okay, I'm confused. Are you envisioning an ACES-71, presumably with payload on top, thrusting backwards all the way through Mars' atmosphere from around ~3100m/s orbital velocity until touchdown?

The DTAL (Dual Thrust Axis Lander) land horizontally. So the payload's "in front". But yeah, I guess it would have to go backwards, unless they mounted the RL-10's in the middle somehow...

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Offline Downix

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Re: Mars EDL technologies
« Reply #181 on: 02/01/2011 06:52 am »
Okay, I'm confused. Are you envisioning an ACES-71, presumably with payload on top, thrusting backwards all the way through Mars' atmosphere from around ~3100m/s orbital velocity until touchdown?

The DTAL (Dual Thrust Axis Lander) land horizontally. So the payload's "in front". But yeah, I guess it would have to go backwards, unless they mounted the RL-10's in the middle somehow...


Yes, it would be going backwards until near the end, when the horiz thrusters would kick in.
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Offline Michael Bloxham

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Re: Mars EDL technologies
« Reply #182 on: 02/01/2011 07:57 am »
Right. But this study assumes a 10m diameter capsule shape - where full aerodynamic drag is preserved even when thrusting - which I assumed meant a truncated cone shape of limited volume. Presumably the engines would also have to be arranged around the periphery of the "heatshield" - or some other exotic configuration - to allow some stability during the retropropulsive maneuvers? (Which makes me wonder whether you will have to protect the craft against potential exhaust impingement?) And of course, you assume that supersonic (and hypersonic?) retropropulsion is viable in the first place?

Also, what happens if the engines are snuffed mid-flight?

Offline Nathan

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Re: Mars EDL technologies
« Reply #183 on: 02/01/2011 08:38 am »
You're kidding, right? LH2/LO2 propulsion will have a higher dry mass fraction for the same amount of thrust. You won't achieve a 10% decrease in ProMF if the final dry mass of the system is higher. You can't eat into your margin to achieve better performance. And just because 4.6% is "missing" doesn't mean it is left over and you can magically add it to your PaMF.


If the Hydrogen is gelled with 4-5% Methane then the tank size reduces considerably with a hit to ISP of ~30s.

I'm quoting another paper I've read that I cannot find.
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Offline rklaehn

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Re: Mars EDL technologies
« Reply #184 on: 02/01/2011 08:47 am »
Right. But this study assumes a 10m diameter capsule shape - where full aerodynamic drag is preserved even when thrusting - which I assumed meant a truncated cone shape of limited volume. Presumably the engines would also have to be arranged around the periphery of the "heatshield" - or some other exotic configuration - to allow some stability during the retropropulsive maneuvers? (Which makes me wonder whether you will have to protect the craft against potential exhaust impingement?) And of course, you assume that supersonic (and hypersonic?) retropropulsion is viable in the first place?

Also, what happens if the engines are snuffed mid-flight?

I think you are mixing two things here. The 10m diameter vehicle with propulsion integrated into the heat shield is for the "reference case" of partially propulsive descent, which is ignored by everybody even though IMHO it makes the most sense.

The idea of partially propulsive descent is to let the atmosphere help you as much as possible without having to add parachutes or other aerodynamic control surfaces other than a simple, vehicle-mounted heat shield.

The whole point of fully propulsive descent is to fully eliminate the heat shield.

See these two statements at the bottom of page 11:

"As offered in this study, a fully-propulsive descent may be used to avoid harsh heating environments usually encountered in atmospheric transit. In this light, the difficulty of developing and employing an enhanced propulsion system is being traded for that of TPS.."

"Throttling authority allows for added control of the vehicle and enables the heat rate limited trajectories which allow elimination of the TPS."
« Last Edit: 02/01/2011 08:55 am by rklaehn »

Offline mmeijeri

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Re: Mars EDL technologies
« Reply #185 on: 02/01/2011 08:56 am »
The 10m diameter vehicle with propulsion integrated into the heat shield is for the "reference case" of partially propulsive descent, which is ignored by everybody even though IMHO it makes the most sense.

I have said at least twice that it is my preferred option - apart from the 10m diameter. Restrict it to what will fit inside an EELV fairing (at least 7m), and I'll be happy with it. Replacing as much of the mass as possible with water for transpiration cooling would be even better, as water is just as good for RLVs as propellant.
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Offline Michael Bloxham

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Re: Mars EDL technologies
« Reply #186 on: 02/01/2011 08:59 am »
But they still assume the same drag profile as the 10m reference case, don't they? I need to read it again...

Offline rklaehn

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Re: Mars EDL technologies
« Reply #187 on: 02/01/2011 09:05 am »
Also, what happens if the engines are snuffed mid-flight?

It is simply not possible for a typical rocket engine to be snuffed out by the incoming flow. The chamber pressure of a rocket engine designed for operation in the atmosphere is much higher than the stagnation pressure of the supersonic flow.

And the choked flow through the throat of the engine effectively isolates what happens inside the combustion chamber to what happens outside.

We may get some data about firing a rocket engine into a hypersonic flow this year when armadillo and masten start doing boosted hops of their VTVL vehicles to higher altitudes.

Offline rklaehn

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Re: Mars EDL technologies
« Reply #188 on: 02/01/2011 09:52 am »
The 10m diameter vehicle with propulsion integrated into the heat shield is for the "reference case" of partially propulsive descent, which is ignored by everybody even though IMHO it makes the most sense.

I have said at least twice that it is my preferred option - apart from the 10m diameter.

I was not talking about you, but about the discussion in general. Usually it's some rube goldberg device that uses 5 stages of aerodynamic deceleration to avoid any propulsive braking, versus fully propulsive braking with its mass ratio disadvantages.

Quote
Restrict it to what will fit inside an EELV fairing (at least 7m), and I'll be happy with it. Replacing as much of the mass as possible with water for transpiration cooling would be even better, as water is just as good for RLVs as propellant.

I think the 10m diameter design was just a reference design. What matters is the ballistic coefficient, so a 7m vehicle would have to be just 30t instead of the 60t of the reference design. Still plenty for a manned mission with surface rendezvous.

Offline rklaehn

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Re: Mars EDL technologies
« Reply #189 on: 02/01/2011 10:07 am »
But they still assume the same drag profile as the 10m reference case, don't they? I need to read it again...

The way I read it, they just keep the 10m reference case for ease of comparison, and because they have to use some shape. But at the low heat rates they use as limits, that 10m shape might as well look like your typical upper stage, as long as the ballistic coefficient is similar.

By the way, a typical uper stage does have some heat shielding as well to cope with radiative heating by the exhaust and by recirculating exhaust.

Offline Warren Platts

Re: Mars EDL technologies
« Reply #190 on: 02/01/2011 02:40 pm »
Quote from: Marsh and Braun
"As offered in this study, a fully-propulsive descent may be used to avoid harsh heating environments usually encountered in atmospheric transit. In this light, the difficulty of developing and employing an enhanced propulsion system is being traded for that of TPS.."

The other thing is you'll save hugely on development costs: on the Moon, aerodynamic deceleration isn't an option, so there is no trade between enhanced propulsion system versus a TPS. Since presumably, we will be going back to the Moon before we launch any manned Mars missions, then the Lunar lander will have to be constructed first. So, the choice is to basically build a single lander that is itself an evolutionary descendant from the Centaur upper stage, or to build two clean sheet designs. The beauty of the fully propulsive lander is that it can land on practically any rocky body in the Solar system, with the exceptions of Earth and Venus.

The other thing is that the fully propulsive lander is easily reused. Isn't it the case that a partially propulsive landing would have to discard the TPS prior to the actual landing?
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Offline rklaehn

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Re: Mars EDL technologies
« Reply #191 on: 02/01/2011 03:00 pm »
The other thing is that the fully propulsive lander is easily reused. Isn't it the case that a partially propulsive landing would have to discard the TPS prior to the actual landing?

For a partially propulsive lander the engines will have to fire during descent when the heat shield is doing its job. Either through the heat shield, or on the side like the propulsive dragon. You can not let the heat shield cover the engines. So why would you discard the heat shield for landing?

The new vertical landing dragon concept comes pretty close to what a partially propulsive lander would look like, except that it would probably use more efficient pump-fed engines with a decent expansion ratio. The engines would also have to fire straight down to avoid cosine loss.

An ablative PICA heat shield would be able to survive several mars descents, so it would be a waste to throw it away.

Offline Warren Platts

Re: Mars EDL technologies
« Reply #192 on: 02/01/2011 04:20 pm »
LH2/LO2 propulsion will have a higher dry mass fraction for the same amount of thrust. You won't achieve a 10% decrease in ProMF if the final dry mass of the system is higher.
If the Hydrogen is gelled with 4-5% Methane then the tank size reduces considerably with a hit to ISP of ~30s.

I'm quoting another paper I've read that I cannot find.

If you really want a dense, high-thrust propellant, then mix in a little ISRU aluminum into the LO2/LH2 mix.

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19860018652_1986018652.pdf

1.6 MB pdf

"The specific impulse of the H2/O2 system is only slightly increased by the addition of aluminum. ... However, aluminum is much denser (2700.7 kg/m3 (168.6 lb/ft3)) than beryllium or lithium; therefore, because of the increase in propellant density, significant payload benefits can result from adding aluminum to bipropellant systems. ...Therefore, since the addition of aluminum to H2/O2 results in an increase in propellant density while performance remains essentially constant, significant payload benefits can result..."
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Offline 93143

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Re: Mars EDL technologies
« Reply #193 on: 02/01/2011 08:18 pm »
On the other hand, an engine that can handle solid aluminium particles is probably unavoidably much less efficient than, say, an RL-60.  Combined with the other practical issues inherent in metal-loaded triprop systems, I suspect it would turn out to be better to just use biprop...

Of course, the idea should be studied impartially to make sure...  we wouldn't want to dismiss a superior option based on gut instinct alone...

Offline Warren Platts

Re: Mars EDL technologies
« Reply #194 on: 02/02/2011 12:18 am »
There's a long thread here on Al/LO2 propellant. It turns out that probably the best method would be Wickman's who did some experiments mixing Al powder in a gelled liquid oxygen--so it was actually more of a monopropellant. So possibly, you could load this mixture into the oxygen tank of an RL-10 and get it to work....

Actually, the Lunar cold traps are loaded with mercury--and it wouldn't be too hard to extract it; it could literally be distilled. That would give you a dense propellant!
« Last Edit: 02/02/2011 12:20 am by Warren Platts »
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Offline guru

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Re: Mars EDL technologies
« Reply #196 on: 02/04/2011 09:46 pm »

It is interesting to me that, as far as I can tell, not a single study has been done to address the question of "how far can heritage EDL tech be pushed?".


I don't have a link to it, and I have not read it, but a reference, or at least a reference to a reference, for such a study is this:

J. Cruz, A. Cianciolo, R. Powell, L. Simonsen, R. Tolson. Entry, descent, and landing technology concept trade study for increasing payload mass to the surface of Mars. Fourth International Symposium on Atmospheric Reentry Vehicles and Systems, Arcachon, France, 2005.

The sentence that references the above study is found in this paper (which I have read):

Ashley M. Korzuna,Gregory F. Dubosa,  Curtis K. Iwataa, Benjamin A. Stahlb, and John J. Quicksallc. A concept for the entry, descent, and landing of high-mass payloads at Mars. Acta Astronautica, Vol. 66, issues 7-8, April-May 2010, pp. 1146-1159.

As the immediately preceding material is copywritten and not openly available, I will forego quoting it directly. But the introduction basically says that, yes, NASA studies have shown that MSL is the limit of what you can do with Viking heritage technology.

(You could probably find this publication at a university library if you don't live too far away from one.  You can also purchase it online. The paper describes a specific method for landing 20 tonnes on Mars, which requires aerocapture, a large diameter aeroshell, an inflatable aerodynamic decelerator, and some retro-propulsive supersonic thrusters (not extensively tested, though some experiments were done by NASA back in the late 60s and early 70s.))

edit: changed hypersonic to supersonic.
« Last Edit: 02/04/2011 09:51 pm by guru »

Offline Robotbeat

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Re: Mars EDL technologies
« Reply #197 on: 02/04/2011 10:06 pm »

It is interesting to me that, as far as I can tell, not a single study has been done to address the question of "how far can heritage EDL tech be pushed?".


I don't have a link to it, and I have not read it, but a reference, or at least a reference to a reference, for such a study is this:

J. Cruz, A. Cianciolo, R. Powell, L. Simonsen, R. Tolson. Entry, descent, and landing technology concept trade study for increasing payload mass to the surface of Mars. Fourth International Symposium on Atmospheric Reentry Vehicles and Systems, Arcachon, France, 2005.

The sentence that references the above study is found in this paper (which I have read):

Ashley M. Korzuna,Gregory F. Dubosa,  Curtis K. Iwataa, Benjamin A. Stahlb, and John J. Quicksallc. A concept for the entry, descent, and landing of high-mass payloads at Mars. Acta Astronautica, Vol. 66, issues 7-8, April-May 2010, pp. 1146-1159.

As the immediately preceding material is copywritten and not openly available, I will forego quoting it directly. But the introduction basically says that, yes, NASA studies have shown that MSL is the limit of what you can do with Viking heritage technology.

(You could probably find this publication at a university library if you don't live too far away from one.  You can also purchase it online. The paper describes a specific method for landing 20 tonnes on Mars, which requires aerocapture, a large diameter aeroshell, an inflatable aerodynamic decelerator, and some retro-propulsive supersonic thrusters (not extensively tested, though some experiments were done by NASA back in the late 60s and early 70s.))

edit: changed hypersonic to supersonic.
Did that paper (which concludes MSL is the limit) actually look at a bipropellant descent stage (with other improvements), a 7 meter diameter heatshield (possible with EELV guppy fairings), and lower altitude landing sites?
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Offline tnphysics

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Re: Mars EDL technologies
« Reply #198 on: 02/04/2011 10:26 pm »
LH2/LO2 propulsion will have a higher dry mass fraction for the same amount of thrust. You won't achieve a 10% decrease in ProMF if the final dry mass of the system is higher.
If the Hydrogen is gelled with 4-5% Methane then the tank size reduces considerably with a hit to ISP of ~30s.

I'm quoting another paper I've read that I cannot find.

If you really want a dense, high-thrust propellant, then mix in a little ISRU aluminum into the LO2/LH2 mix.

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19860018652_1986018652.pdf

1.6 MB pdf

"The specific impulse of the H2/O2 system is only slightly increased by the addition of aluminum. ... However, aluminum is much denser (2700.7 kg/m3 (168.6 lb/ft3)) than beryllium or lithium; therefore, because of the increase in propellant density, significant payload benefits can result from adding aluminum to bipropellant systems. ...Therefore, since the addition of aluminum to H2/O2 results in an increase in propellant density while performance remains essentially constant, significant payload benefits can result..."


Al in LH2-that might allow SSTO from Earth!

Offline guru

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Re: Mars EDL technologies
« Reply #199 on: 02/04/2011 10:51 pm »

It is interesting to me that, as far as I can tell, not a single study has been done to address the question of "how far can heritage EDL tech be pushed?".


I don't have a link to it, and I have not read it, but a reference, or at least a reference to a reference, for such a study is this:

J. Cruz, A. Cianciolo, R. Powell, L. Simonsen, R. Tolson. Entry, descent, and landing technology concept trade study for increasing payload mass to the surface of Mars. Fourth International Symposium on Atmospheric Reentry Vehicles and Systems, Arcachon, France, 2005.

The sentence that references the above study is found in this paper (which I have read):

Ashley M. Korzuna,Gregory F. Dubosa,  Curtis K. Iwataa, Benjamin A. Stahlb, and John J. Quicksallc. A concept for the entry, descent, and landing of high-mass payloads at Mars. Acta Astronautica, Vol. 66, issues 7-8, April-May 2010, pp. 1146-1159.

As the immediately preceding material is copywritten and not openly available, I will forego quoting it directly. But the introduction basically says that, yes, NASA studies have shown that MSL is the limit of what you can do with Viking heritage technology.

(You could probably find this publication at a university library if you don't live too far away from one.  You can also purchase it online. The paper describes a specific method for landing 20 tonnes on Mars, which requires aerocapture, a large diameter aeroshell, an inflatable aerodynamic decelerator, and some retro-propulsive supersonic thrusters (not extensively tested, though some experiments were done by NASA back in the late 60s and early 70s.))

edit: changed hypersonic to supersonic.
Did that paper (which concludes MSL is the limit) actually look at a bipropellant descent stage (with other improvements), a 7 meter diameter heatshield (possible with EELV guppy fairings), and lower altitude landing sites?

I am not certain what all the original study looked at, as like I said, I haven't read it.  However, the papers I have read (most especially the second one just referenced) have all stated that we can't do any more with what we have without better heat shield materials, better (as yet undeveloped) parachutes / aerodynamic decelerators, or other technologies.  All of the incremental improvements that can be made with Viking qualified technologies, the studies state, have pretty much been done, and anything bigger requires new qualification programs - the Viking one was very costly to start with, which is why they haven't done any since.

Changing to a bi-propellant landing stage will offer only modest improvements in your landing payload.  It is only used for the last portion of descent after the parachutes have already slowed the craft to subsonic velocities anyway.  For example, MSL has about 300 kg of propellant; dry, the rover + skycrane weigh about 1400 kg.  Even if you replaced the mono-propellant with a bi-propellant you would still only reduce your propellant requirement by about 100 kg for the same amount of deceleration, so you could potentially get 1500 kg instead of 1400 kg.  Adding more propellant than that actually reduces your payload because the Martian atmosphere still has to decelerate that mass.

Larger heatshields mean, necessarily (barring breakthroughs in structural or material sciences) higher ballistic coefficients.  They don't have to scale (correction): times 3/2, but they do still scale somewhat linearly with diameter, as opposed to not at all.  Dropping your ballistic coefficient can actually be somewhat detrimental as entry forces on the MSL are already 12 g's at an entry velocity of 5.9 km/s; reducing the ballistic coefficient would increase the rate of deceleration.  This is one reason why I think it is better to start re-entry from low Mars orbital velocity rather than doing a direct entry - the lower velocity helps to limit the re-entry forces to under 8 g's.

Heating also goes up with the diameter because the Reynold's number increases (unless the velocity decreases), causing the ship to transition to turbulence prior to peak heating.  The increased heating is just a fact of turbulent mechanics.  Current heat shield materials can't handle higher heating loads, and the second paper cited does mention that specifically.

They did plan the 20 tonnes to land at a mean altitude of 0 km, so that was considered.  You could improve on the MSLs landing payload by dropping down to 0 km, as opposed to the 1 km it is slated for, but I don't know how much exactly - I too would be interested to know that.

« Last Edit: 02/04/2011 10:54 pm by guru »

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