While most "gold" thermal blankets don't have any gold (the orangish ones are aluminum and the color comes from the Kapton/polyimide plastic of the blanket substrate), there are indeed some gold products with very low emissivity: see

http://www.sheldahl.com/Product/TMThermal....ermal%20Control

and this may well be what NH is using.

Certainly. In fact, I might say that this is the only way to reliably determine the PSF, though it's complicated by the fact that you are limited by the sampling frequency of the system's detector and you would usually not oversample the PSF as much as you'd like to if you were just trying to measure the PSF, not take useful images.

However, I haven't seen anyone on this forum going through the PSF determination process, and without doing that, you can't do deconvolution, you can only do filtering.

Just a pedantic nitpick, perhaps, but I think it's important to be precise about these things.

Despite what's been said in other discussions in this forum, you need knowledge of a system's point-spread function to do deconvolution. Anything you do without knowledge of the PSF is just regular high-pass filtering, and you can do that with Photoshop perfectly well.

I think Bruce explained it pretty well. The mass of the motors and antenna is only a small fraction of the whole DTE system (if the total mass of that is 50 kg, they really aren't trying very hard, but that's another story) and trying to cantilever more mass up at the top of the RSM, which has to be stable enough to support Mastcam imaging, has to move rapidly to take mosaics, is a meter or more off the ground when the antenna has no need to be that high -- well, it just wouldn't end up saving mass.

Not the mention the fact that they have to have a fat waveguide between the power amps and the transmitter (presumably in the main rover warm box) and the antenna.

But we're just designing the cameras. JPL designs the mast and its pointing and deployment mechanisms. All I can say is that the RSM design is hard enough to make work as it is.

First, we call it the Mastcam. Or maybe the MastCam, I'm not sure. The whole mast is generically called the Remote Sensing Mast (RSM), since the navcams and the LIBS (aka Chemcam, or is it ChemCam?) is on it as well.

Second, there's just no way the extra mass and volume of an antenna could be supported by the RSM.

And the paper journals don't have that concern? I'm a bit confused about what web-based peer-reviewed venues you're talking about. I'd say that Dave Paige's new online journal MARS is as well-peer-reviewed as SCIENCE.

That said, you can't beat the prestige of SCIENCE. I guess NATURE is the same way, though I don't like it 'cause they spell all their words funny

Seems to me that the 2006 NASA budget I cited is clear proof that they have selected the MMRTG. And I'll just leave it at that.

Even if any of these things were true (they're not, as far as I know -- there's nothing terribly enabling about going to twice the RTG power. You still need batteries, you still don't have enough power for active thermal control of all elements so you have thermal cycling wearout, you still have mechanical wearout problems, and there is no way that MSL is ever going to drive 1 km/day under any circumstances no matter how much power it has) MSL is using the MMRTG anyway. So your drool is wasted.

The full document makes it pretty clear that yes, they mean two entire separate SRGs if you need the power from one.

If you were an advocate of the SRG you might say that was an unfair imposition, but them's the rules at the moment.

You make a fair point about the radiators, though even MPF and MER had radiator systems.

I can't assess the mass estimates of the RTG systems themselves without detailed technical descriptions.

However, from http://solarsystem.nasa.gov/scitech/display.cfm?ST_ID=705

"Initially, the MMRTG could have an advantage from a mass perspective, as current NASA/DOE guidelines recommend that early missions using SRGs carry at least one redundant SRG unit until its reliability has been verified [11]. This means that early missions using SRGs would need to carry a minimum of two SRG units. Thus, for early missions (where a redundant SRG would be required), the MMRTG (at <45 kg [10]) would be the lighter option for spacecraft requiring one or two RPS units."

And according to the NASA FY06 budget request ( http://www.nasa.gov/pdf/107489main_FY06_1_sae.pdf ), page 2-22 "MSL - Department of Energy for Multi Mission Radioisotope Thermoelectric Generators" and page 2-14 "Radioisotope Power System (RPS): Deleted Small RPS or second generation Sterling (SRG), and RPS Power Conversion Technology (RPCT)." I think it's clear from that that MSL is planning to use the MMRTG.

Not quite. The mission was Mariner Venus-Mercury 1973 (aka Mariner 10) but the realization was that if the bit error rate was increased the bit rate could also be increased, and with imaging it didn't matter much if the bit error rate was higher, since isolated bad pixels could be removed with simple filtering techniques (such as median filters.) MVM73's bit rate was 117 Kbps at a BER of 5e-3 for this reason.

See http://history.nasa.gov/SP-424/app-b.htm for more info.

Actually, doing a little more research, the SRTG is even less obviously a win. It saves Pu mass, but the additional mass of the more complex conversion hardware partially kills that advantage. Depending on exactly what mass the MMRTG finally comes in at, it can have a higher energy density than the SRTG. Getting rid of the additional waste heat, while an issue, is a fairly easy problem to solve.

Quoted from "Advances in Planetary Aerobots" by Erik Laan et al:

"The MMRTG will be designed to generate
110 Watts of electric energy over a minimum lifetime
of 14 years (3 years on the Martian Surface)
weighing 24-34 kg (218.8 - 309.1 kg/kW) including 4
kg Plutonium-238 (Boeing)
Another development is aimed at increasing the
efficiency of the conversion process. Current RTGs
are capable of converting heat into electricity with an
efficiency of ~8%. A number of new conversion
processes were proposed recently. The Stirling
Radioisotope Power Source is one of them. This
SRPS is using a Stirling heat engine, which produces
an acoustic pressure wave to drive a piston in a linear
alternator producing the electricity. The SRG
(Stirling RTG) delivers 93-114 Watt of electric
energy, weighing 27 kg (236.8 - 290.3 kg/kW) and
contains 1 kg of Plutonium-238 (LMM)."

The report I referenced is probably just a baseline; I don't know if the final RTG configuration has been picked. However, I'm not sure that the Stirling RTG is a clear-cut winner. It's more efficient and lighter for a given power output, true, but it's far more mechanically complex and has its own set of problems (vibration, electrical noise, etc.)

According to
http://trs-new.jpl.nasa.gov/dspace/bitstre...6/1/05-0708.pdf
MSL will use an MMRTG. I would think that the use of batteries is inevitable, as there's no way the avionics can run directly from RTG output, much less the payload.

I'm sure they've done a detailed trade study of alternative RTGs, but I haven't seen those.

Well, obviously a large portion of the proposal evaluation and selection process is assessing the cost realism of the proposal, so it's not like you can just claim anything and be believed. Of course, this cuts both ways -- people can end up claiming that a mission can't possibly be done for a given cost cap when it certainly could, just perhaps not by the organization doing the evaulating. That's one of the factors which IMHO has led to spiraling cost increases in aerospace in general.

Gee, it must be nice to be infallible.

Calling it the "english vs metric unit thing" is something of an oversimplification. You might want to read
ftp://ftp.hq.nasa.gov/pub/pao/reports/1999/MCO_report.pdf before being so casually dismissive of those of us who spent several years on the MS98 project (not that I had anything to do with the nav error.)

I don't think either would work; conditions in the upper atmosphere can only be assessed to the needed accuracy by entry probes at the time of entry.

*Maybe* atmospheric sounding, such as will be provided by MCS on MRO, will give some insight, but I don't think to the desired level.

Another reason that building systems that can handle the variations in real time would be better than trying to characterize and predict conditions in advance.

I don't recall seeing specific analyses of the trajectory based on final tracking data, though it certainly exists in some form. In the pre-launch mission plan, closest approach was over about 30N, 170W and the spacecraft was moving southward; the orbit was near-polar (inclination 92.93 deg) but I don't know if the closest approach coordinates varied with actual day of launch.

Maybe instead we should use entry systems with more margin against these sorts of problems. I don't know how much of what Bruce is saying about Spirit is reality and how much engineering conservatism, but it's always been understood that MER had fairly poor robustness in the face of some environmental factors. There are two engineering responses to this: ask "the scientists" for better environmental data, or design proper margins into the system based on the uncertainties (probably losing capability in the bargain.) One is looking for the proper mix, but it's likely not on one end or the other of this spectrum.

The most likely outcome stated by the Coffey board in 1993 was that MO was spun up to a high rate by helium expelled from a blown propellant line. The delta v this would impart could end up averaging to near-zero. I tried to get people to seach for MO in 2000 when it would have approached to about 0.2 AU from the Earth, but it was very dim, and I couldn't interest anyone in spending telescope time on such a long shot.

A paper written in 2001 by Carl Guernsey of JPL, though ( http://www.klabs.org/richcontent/Reports/F...y_a01-34322.pdf )
suggests it was more likely that MO suffered contamination in the He pressure regulator and that the oxidizer tank just exploded, which would have blown the spacecraft into pieces.

I think it's probably sensitively dependent on the state of the martian upper atmosphere, which we don't know, and on the breakup dynamics of the spacecraft, which would be very hard to predict. As was pointed out, it's not even known if MCO debris fell on Mars or ended up in solar orbit.

If you are using a first-difference-based lossless compressor, you can get this for free depending on what encoding table you use and what your black-space noise level is. For example, on MGS/MOC, we use a table that compresses black space 8:1, and we occasionally use this for star calibration images.

Lossless compression algorithms can certainly be fast enough to be applied in real time, if that's desirable. For example, the MGS/MOC lossless compressor can compress the raw pixel rate of 5 megapixels/sec. I'm not certain, but I suspect that the NH camera pixel rate is less than that; it probably doesn't need to be more than a megapixel/sec or maybe even less. It would also surprise me if NH was using lossy compression; we certainly didn't propose that for our unselected PKB instruments.

The pressure isn't as much a problem as the temperature. There are probably a lot of industrial uses for high-temp electronics -- jet-engine controllers, for example. When we were writing our Venus probe proposal (which I like to call "Tom Swift and His Nuclear Refrigerator" -- see http://www.msss.com/venus/vgnp/vgnp.txt.html ) there was a lot of work being done on silicon carbide (SiC) electronics for engine control. Haven't looked lately; it wasn't really viable in the mid-90s.

I don't know for sure, but I think the view from Jetty Park of LC41 would be pretty bad. LC41 is way north, very close to LC39. I would think any good site for a Shuttle launch would be better for LC41. It's not like a Delta launch, where the view from Jetty Park is probably as good as from the VIP site.

We watched the MRO launch from the causeway site and it was OK.