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I was reading about the problems of retro rocket landings on Mars. Apparently the rocket exhaust acts like an unstable nose cone and because Mars unlike the Moon has winds , it is a bit of a liability in trying to land smoothly and successfully.
Does anyone have any comments/thoughts on this?
One thought I've had is whether a "tripod" of rocket exhaust angled at three equidistant points on the side of the lander would help stabilise it.
Another thought: Harrier Jump Jets have mastered the art of landing with jet exhaust i.e. a retroactive system. Does that afford any insights?
Are NASA being unduly cautious about the scope for retrorockets?
Let's Go to Mars...Google on: Fast Track to Mars blogspot.com
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The Skycrane landing for MSL is just that system but I am sure that it has a mass limit just like the bouncing air bags did for the past mars rovers.
Part of the problem as you noted is the existance of the thin atmosphere which means the heat shield needs to stay intack for a very long time as it descends from orbit. Once the parachutes can open only then can the engines be exposed to the atmosphere as the thermal temperatures have final dropped such that it would not damage the engine.
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The problem is that Retro rockets may not work well at supersonic speeds as they have to fly backwards through their exhaust plume. Entry either from orbit or direct from Earth is at very high speed, the lander has to slow down sufficiently to use retro rockets and given the thin atmosphere it's difficult to achieve with heavy payloads without hitting the ground. Research continues.
Viking, Phoenix and MSL use retro rockets. MSL will be the heaviest, landing about 800 kg on the surface.
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CI -
This is probably a stupid question - but why can't you slow down the rocket before the descent. Why can't you say slow it down as you approach Mars orbit? Is it just the fuel use which is the issue?
I'm wondering why we can't slow to subsonic speeds and then descend gracefully to the planet's surface.
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Yep, slowing down before entering the Martian atmosphere needs fuel and that needs bigger fuel tanks. All that extra mass reduces the payload. As the payload gets heavier and bigger, the vital heat shield and parachute system also needs to be heavier and bigger. With current technology this type of system is mass limited and can't reduce the lander speed enough in the Martian atmosphere to land safely either with rockets, airbags or any other means.
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Ciclops -
I'm not disputing what you say, but I do want to understand it better.
What theoretically is the position in terms of fuel use if we:
Slowed the craft down from 000s of MPH to slow subsonic before we are anywhere near the Mars atmosphere and then for the descent we slowed down to speeds where the craft would NOT heat up as it passed through the atmosphere - a "straight down" drop.
Any ideas?
I'm looking to compare (i) the fuel use involved in a conventional landing using a mix of heat shield, parachutes and retro firing with (ii) pure retro as described above.
If the answer is 200% extra fuel compared with conventional landing I think it might still be worth pursuing the idea. If the answer is 10,000% extra, obviously it can be forgotten.
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Every drop of fuel has to be accelerated to over 11 kms/second to escape Earth's gravity. Once near Mars the pull of its gravity will be continuous. To give a feel for the magnitude of the problem, a spacecraft on a standard trajectory from Earth will enter the top of the Martian atmosphere at about Mach 27.
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It is always hard to understand values for comparison when they are mixed...so to save confusion give both with one in brackets or something.
We treat mach numbers as constants but in actuallity they are not as it is dependent on temperature and atmospheric composition.
Most consider the speed of sound to be at Sea level, 340.3 m/s (1,225 km/h, 761.2 mph, or 661.7 kn) in the Earth's atmosphere. Mach 1 however at 11,000 m (36,000 ft), would be traveling at 295 m/s (654.6 mph, 1,062 km/h, 86% of its speed at sea level).
Much like the mach number gravity as well is not a constant as distance form the object it will drop.
A range from 9.789 m·s−2 at the equator to 9.832 m·s−2 at the poles results from centrifugal force and equatorial bulge.
Body Multiple of Earth gravity m/s²
Sun 27.90 274.1
Mercury 0.3770 3.703
Venus 0.9032 8.872
Earth 1 (by definition) 9.8226[5]
Moon 0.1655 1.625
Mars 0.3895 3.728
Jupiter 2.640 25.93
Saturn 1.139 11.19
Uranus 0.917 9.01
Neptune 1.148 11.28
Pluto 0.0621 0.610
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Yes, it's better to quote velocity directly, for Mars EDL it's about 5 kms/sec at the entry interface around 130 kms above the surface. That's a lot of velocity to burn off with the propulsion system.
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I accept it's a lot - just be interested to know how much in terms of fuel use as a percentage of a "conventional" landing using parachute, heat shield and some retro rockets.
Has anyone out there got a calculator and the requisite knowledge?
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The amount of fuel needed to reduce the spacecraft velocity can be calculated using the infamous rocket equation:
Initial vehicle mass/Final mass = e^(dV/Ve)
where dV is velocity change and Ve is exhaust velocity
dV will be about 5000 m/sec and Ve about 4500 m/sec for the best LOx/LH2 engine (Isp 450 secs)
That gives a mass ratio of about 3:1 - so to change the velocity of a spacecraft with a dry mass of 1 MT by 5 kms/sec requires about 2 MT of fuel.
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Well that is not unreasonable.
Look at it the other way, the amount require to kill all its speed and land softly will be pretty close to what is required to send it back to Earth.
How much speed would you have to bleed to justify the risk of having an insertion burn, vs the risk of having a heavy highspeed aerocapture.
You will probably have an engine anyway for course correction, and a large engine for landing, so you are really looking at increasing your fuel load quite a bit. MD already calls for tanking 6 tons of H2 anyways its more of a cpacity thing than adding equipment.
Come on to the Future
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It may appear reasonable (assuming that calculation is correct) however that 1 MT has to include the engine, fuel tanks, parachutes and structure, leaving almost nothing for the payload. Current propulsion systems are too inefficient to be used for braking to parachute speeds, that's why all landers have used heat shields.
Yes the amount of energy is about the same as is required to escape Mars gravity. Low Mars orbit is about 4 kms/sec, and it's just possible to burn off the extra 1 kms/sec with current technology. For heavier payloads aerobraking has to be used, as was done with MRO. It's still not clear how a MD size payload (about 10 MT) can be safely landed on the surface.
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That's quite encouraging - if I've understood it right.
I'm a Mars Minimalist I think we can get there and establish permanent human settlement on a very small payload - but we have to change our way of thinking about the problem.
I'm also a retro rocket enthusiast.
So I;m suggesting a much smaller payload than some others do.
I think for Mission 1 we could go with two robot pre-flights delivering 10 tonnes each followed by two companion manned flights and landings of 10 tonnes each. So 40 tonnes total of which maybe 24 tonnes would be deployed following landing.
So do you have any thoughts on how big a single stage craft launched from earth using only retro rocket landing would have to be for a 10 tonne payload. Would something like 2000 tonnes be about right? Or is that too low?
Another way of addressing the problem once we had a lunar base would of course be to have the craft launch from earth with fuel tanks half or third empty, to refuel on the moon from lunar fuel. That seems like a very sensible way of proceeding to me.
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One reason why a large diameter heatshield is desired as it slows the craft as it glides through the atmosphere more than a small one would.
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Also do not forget, that 1MT must also include the fuel tanks for the fuel too. Overall, powered breaking isn't out of the question, but if you are going to do that and bring the fuel you need for the trip back to Earth, forget about chemical rockets. Nuclear rockets on the other hand could pull it off perhaps.
Oh and because of the thin Martian atmosphere, a big heat shield is probably preferred, unless your vehicle has a relatively low density.
And last, Lunar refueling is a bad idea, because it takes about as much fuel to get to the Moon then just to go to Mars directly.
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GCN -
You say:
"Also do not forget, that 1MT must also include the fuel tanks for the fuel too. Overall, powered breaking isn't out of the question, but if you are going to do that and bring the fuel you need for the trip back to Earth, forget about chemical rockets. Nuclear rockets on the other hand could pull it off perhaps."
That suggests you have an idea about how much fuel or overall rocket mass would be required to get 10 MT to Mars and back. Are you able to give the figure. I was thinking perhaps we were talking about 2,000 MT but are you saying it is much higher. If you're not why are you saying we should "forget it"?
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Food for thought would be to turn back to the old Apollo LM which was a 2 stage unit.
As you can see it had a huge by comparison descent stage to land versus the much smaller ascent. Keep in mind though it had to dock with the departure stage that was parked in orbit for the means to return home.
You will see these same sizings of stages even in the current Altair - Lunar Lander (LSAM) - status thread in this folder.
Mars will be no different in those respects.
I did quite a bit of document gathering for the landing issues for mars and posted then under the MarsDrive (Human) Mission Design in response to Michael Bloxham needs and noted that other threads existed for a mission discussion on other forums lower down on that page for alternative mission designing.
As you noted it is not a desire to land with the return fuel so you would leave that in orbit and the ascent is a question of size of the insitu processing mass that needs to be landed in order to refuel as a source of hydrogen is needed for it to happen.
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Wow searching for topics and came accross the title which is the Red Dragon retro propulsion landing rocket system long before space x was starting to make the dragon....
We sure were a forward thinking group back then before the crash of the web site....
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Spacex landing Falcon first stages disproves most of the feared risks of supersonic retropropulsion.
My ancient Hoerner "drag bible" has NASA wind tunnel data from 1961 with a Mercury capsule shape somewhere around Mach 3. It shows the effects they saw on drag due to the retropropulsion plume: a slight reduction in effective drag coefficient. Not a word about plume instability or the side forces that could induce.
The plume coming out of the rocket nozzle is moving at Mach 9 plus, depending upon the expansion ratio of the nozzle. The oncoming stream for an Earth entry (not at Mars!) can be faster than that, but the rocket plume exit plane pressure is far higher than local atmospheric pressures, so there is far more momentum in the exhaust plume than in the oncoming stream.
The rocket plume must shock down to subsonic (relative to the oncoming stream, before it can turn and go with that stream. That is where any plume instability can occur. Once shocked down, momentum is much smaller, which is why the plume instability effect is not large.
The EDL (entry, descent, and landing) dilemma is all about what happens if you keep using the same heat shields, ringsail chutes, and terminal retrorockets) as payload mass grows. It's not the chute, it's the hypersonic entry trajectory.
At higher ballistic coefficient (which grows with mass on a square cube effect), you reach Mach 2.5 (the max opening speed of a ringsail chute) at lower altitudes. That CANNOT be avoided.
At lower opening altitude, there is less time available for the chute to decelerate you from Mach 2.5 to high subsonic (the best it can do, by the way, on Mars).
As the mass grows, you very quickly reach a point (about a ton or so) at which you start hitting the ground still supersonic. Carried to a ridiculous extreme, you don't even have time to deploy and open the chute.
All that conundrum really says is "do something different".
Forget the chute and make the retrorockets much bigger, to shoulder the deceleration load about the time you come out of hypersonics about Mach 3-ish. Now you can land several tons.
The other piece of this puzzle is chute scaling. Bigger masses require bigger chutes. There is a limit to chute loading (mass/opened planform area) that is set by the ability to actually deploy and successfully open the thing. True here and on Mars, just different numbers. Chutes are infeasible under any circumstances (here and Mars) if the payload mass gets big enough. Again, it's a square cube scaling thing.
Now you know why Spacex did what it did with Falcon stages, proposed what it did for Crew and Red Dragon, and is developing for both stages of its BFR. Bigness drives you to retropropulsion as the most feasible landing means.
GW
Last edited by GW Johnson (2019-03-23 09:53:55)
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Multiple atmospheric passes will bleed off a lot of velocity, even at Mars with its nearly nothing atmosphere. I expect that the use of retro rockets would give better control of this process and minimise the risk of skipping off. Just thinking how we might improve our mass ratios...
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Mars arrival direct from the interplanetary trajectory is pretty near 7-7.5 km/s at entry interface. Mars escape is 5 km/s. You'd have to lose 2-2.5 km/s speed aerobraking, in one single pass, in order not to inherently skip off into space. That's a tall order, given how thin Mars's air is.
GW
GW Johnson
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"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Perculating a topic back to the fore front of discusion as GW has put together baseline numbers for landing starship on mars. The roseta stone of getting large mass to mars surface. Star ship is still going through design changes while in a build of prototype process that is being used to test theory.
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The real difference is that if the crane had long legs it would not need to winch the rover to the gound as it had enough fuel to finish the landing. The s curve entry path is used to aerobrake the crafts speed over a longer distance while its still in the upper atmosphere.
So all that we really think that a dragon would not be used is the parachutes but these would be needed for earth return.
When we look at the systems of both fuel and engines seem to be the answer for retro propulsion.
https://en.wikipedia.org/wiki/Mars_Scie … ding_(EDL)
The mass of this EDM system, including parachute, sky crane, fuel and aeroshell, is 2,401 kg (5,293 lb).
The descent stage is a platform above the rover with eight variable thrust monopropellant hydrazine rocket thrusters on arms extending around this platform to slow the descent. Each rocket thruster, called a Mars Lander Engine (MLE), produces 400 to 3,100 N (90 to 697 lbf) of thrust and were derived from those used on the Viking landers.
I think you see where I am going with the scability towards what the dragon and larger as a means to prove we are on the correct path.
More launch details
https://www.jpl.nasa.gov/news/press_kits/MSLLaunch.pdf
Mars Science Laboratory: Entry, Descent, and Landing System Performance
https://ntrs.nasa.gov/archive/nasa/casi … 007730.pdf
Just hoping to fill in details for scaling for any shape that we would land on mars
Same landing was used by MSL
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The insight and Pheonix landers are nearly the same for retro propulsion...
The craft will hit the top of the atmosphere at a speed of 5.7 kilometers per second (12,750 miles per hour). Within the next six and a half minutes, it will use heat-generating atmospheric friction, then a parachute, then firings of descent thrusters, to bring that velocity down to about 2.4 meters per second (5.4 miles per hour) just before touchdown.
There is no guarantee of a successful landing, despite extensive analysis, testing and review of the entry, descent and landing system.
About 75 seconds after the parachute opens and 140 seconds before landing, the spacecraft will start using its radar. The radar will provide information to the onboard computer about distance to the ground, speed of descent and horizontal velocity. It will take readings at a pace of 10 times per second until touchdown.
Descent speed will have slowed to about 56 meters per second (125 miles per hour) by the time the lander separates from the back shell and parachute, about a kilometer (six-tenths of a mile) above the ground. The spacecraft will be in free fall, but not for long. Thrusters will begin firing half a second later and will increase their thrust three seconds after Phoenix sets itself free from the parachute. Touchdown will still be about 40 seconds away. The onboard computer will use information from the radar to adjust the pulsed firings of the 12 descent thrusters.
By the time the lander gets to about 30 meters (98 feet) above the surface, it will have slowed to about 2.4 meters per second (5.4 miles per hour) in vertical velocity. Continuous adjustments to the thruster firings based on radar sensing will also have minimized horizontal velocity and rocking. Touchdown will be about 12 seconds away.
ouch if the thrusters to slow the ship do not work as you will find mars in less than 20 sec....
Twelve thrusters mounted around the bottom edge of the lander will slow the descent during the last half-minute before the legs touch the surface. These can each pulse on and off for fine-tuning the velocity and for maintaining the lander’s stability -- controlling its pitch, yaw and roll -- as it approaches touchdown. They each provide about 293 newtons (65.9 pounds) of thrust.
https://en.wikipedia.org/wiki/Phoenix_(spacecraft)
Launch mass 670 kg 1,477 lb)
landing Mass 350 kg (770 lb)
eight 1.0 lbf (4.4 N) and 5.0 lbf (22 N) monopropellant hydrazine engines built by Aerojet-Redmond
https://mars.nasa.gov/insight/timeline/ … t-landing/
Compared with Phoenix, though, InSight's landing presented four added challenges:
InSight entered the atmosphere at a lower velocity -- 12,300 miles per hour (5.5 kilometers per second) vs. 12,500 miles per hour (5.6 kilometers per second).
InSight had more mass entering the atmosphere -- about 1,340 pounds (608 kilograms) vs. 1,263 pounds (573 kilograms).
InSight landed at an elevation of about 4,900 feet (1.5 kilometers) higher than Phoenix did, so it had less atmosphere to use for deceleration.
https://www.jpl.nasa.gov/news/press_kits/insight/
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