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?
Quote from: Michael Bloxham on 02/01/2011 06:20 amOkay, 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...
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.
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?
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.
Also, what happens if the engines are snuffed mid-flight?
Quote from: rklaehn on 02/01/2011 08:47 amThe 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.
But they still assume the same drag profile as the 10m reference case, don't they? I need to read it again...
"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 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?
Quote from: Michael Bloxham on 02/01/2011 05:07 amLH2/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.
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.
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?".
Quote from: Michael Bloxham on 01/29/2011 10:56 pmIt 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.
Quote from: Nathan on 02/01/2011 08:38 amQuote from: Michael Bloxham on 02/01/2011 05:07 amLH2/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.pdf1.6 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..."
Quote from: guru on 02/04/2011 09:46 pmQuote from: Michael Bloxham on 01/29/2011 10:56 pmIt 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?