Smallest Human Ascent or Descent Lander for Mars Or Earth

GW Johnson · 2015-03-27 12:52:39

Louis:

The reason an Apollo-style lander is inappropriate for Mars is the presence of that near-vacuum-of-an-atmosphere. There are hypersonic drag and friction-heating phenomena (as in the photos posted just above), and there is a real need for both proper aerodynamic shape and friction heat protection (just not as much heat shield as here).

Low Mars orbit (LMO) velocity is near 3.5 km/s. The typical deorbit burn is only a very few dozen m/s, so you hit "air" at pretty near orbital speed. For direct shots to Mars, your entry speed is even faster, maybe 6-ish km/s. The trick is to hit at a very shallow angle so you don't strike dirt before the hypersonic drag can slow you down. The "air" is really near-vacuum-thin, so that's a lot trickier to do than it is here at home.

You'll come out of hypersonics at around local Mach 3, which on Mars is about 0.7 km/s velocities. Depending upon your angles, your shape's hypersonic lift capability, and your ballistic coefficient, the altitude when this happens could be quite dangerously low. Big, heavy stuff might be well under 5 km altitude, maybe less than 2 km. If so, chutes are useless: no time to deploy, much less decelerate.

You will need for all practical purposes roughly about 2 x the delta-vee-to-kill for maneuvering rocket braking to the surface, including some hover time, with no chutes. That's about 1.5 km/s landing delta-vee required as a maximum. If you can use chutes, this is reduced, but only somewhat. Even terminal speeds with chutes are quite high in that thin "air". Significant rocket braking to touchdown is just absolutely required on Mars.

So from LMO, you need about 1+ to maybe 1.5 km/s terminal rocket brake capability, with the thin "air" killing all but 0.7 km/s of your entry speed of 3.5 km/s. So, air drag kills about 2.8 km/s of your orbital entry speed for you (even more for direct entry). That's totally unlike the moon situation. Descent propellant requirements are rather modest, even with no chutes.

That and heat protection is why an Apollo-style lander is the wrong thing to attempt at Mars. A one-way Mars lander is closer to a traditional space capsule than anything else, just one that is powered with a lot of propellant. Like manned Dragon v.2.

Takeoff from Mars is much more like the moon. You never hit high speeds until you are already outside the thin "air". The vehicle hardly needs any streamlining, although that does help a little. You'll need a minimum delta-vee of LMO velocity plus a bit of maneuvering margin, say 4 km/s total. The thin "air" doesn't help you, the way it does on descent. Thus any ascent craft will need a lot more propellant than a descent-only craft, even if you make it by ISRU on the surface.

Yet if (1) you take full advantage of descent drag, and (2) you only need the bare minimum to reach min LMO again, this descent/ascent job can be done in a potentially-reusable single stage with ordinary chemical propulsion, because the total delta-vee for the round trip is only near 5.5 km/s or maybe just a tad more. The payload fraction is quite low (crudely 3% or so) at realistic inert mass fractions (15-25%?), but the reusabilty aspect potentially offsets this, as long as propellant can be made available to re-fly it.

GW

SpaceNut · 2015-03-27 19:18:34

The Entry, Descent and Landing is the issue with the current estimated (40T landed) mass that is suggested to be necessary for mars manned landings which is currently beyond our capability which sort of fostered this topic. The fact that we need to send any landing mission as a demostrator of that size means we waste a huge amount of cash and still are possible going down the wrong path with those futuristic designs that we might come up with to make that 40T possible. Its less of an issue to get the mass to mars orbit so why not solve for current state just like we are when we say we need on the order of 500 T plus to LEO by smaller building blocks, so why not use the same approach going down for mars making the pieces more manageble by sending down smaller chunks. If it solves the problem of landing what is needed.

kbd512 · 2015-03-27 19:22:01

RobertDyck wrote:

You keep repeating the same error. Yes, I agree that defence (Canadian spelling) contractors want big spending. Yes, we need to keep cost down. Congress said "No!" to the 90-Day Report because it cost too much. The problem is you keep assuming that splitting a single, shared spacecraft into multiple single-person landers would actually reduce mass. I've explained why it won't.
Let me put it another way. Physics is real life. This isn't some Hollywood movie or TV show. The TV show "Stargate SG-1" showed aliens with technology that looked like ancient Egypt. It's based on the initial premise of the show. A couple episodes showed an escape pod that looked like an Egyptian sarcophagus. Some episodes showed a horizontal sarcophagus as some sort of medical device, other episodes showed a vertical sarcophagus in a Tel'Tak ship as an escape pod. Well, in real life a sarcophagus is a coffin. Just a box to hold a dead body. It isn't an escape pod. In real life an escape pod cannot be a form-fitting coffin-shape. In real life an entry vehicle must deal with high speed atmospheric entry.
Studies by NASA in the 1950s showed a blunt body works much better than a sharp, streamlined vehicle. A blunt body means a round heat shield, and an aeroshell behind it. Wind tunnel studies showed air flow behind the heat shield in a specific pattern. The reason a Mercury spacecraft has the shape it does, is that fits within the wake behind the heat shield. A cone of specific size, with specific angle to the side walls. At a specific distance behind the heat shield, the wake stops collapsing inward, forming a fairly straight cylinder of turbulent air. The nose bit of Mercury, fits within that. Mercury used that to store its parachute, flotation bags, and life raft. And above that was the antenna fairing. But you don't need to use that space. A spacecraft that docks to something, such as ISS, would have a rounded-off, domed top with a docking hatch rather than that nose piece.

Rob,

Get the whole MOOSE concept out of your head. NASA isn't doing space jumping stunts on Earth or Mars or anywhere else there is a potential alternative EDL method.

The only way you need a capsule the size of Mercury is if you insist on laying the astronaut flat on his back. The human doesn't have to be oriented that way in the capsule for a Mars landing because we're not executing a Mercury style landing on Mars.

Imagine for a second that we're using an inflatable heat shield, which is what I proposed, and that the capsule's seating arrangement is such that the astronaut is sitting in a fabric seat in a more vertical orientation, relative to the nose of the capsule, instead of laying on his or her back. If the capsule was designed with that seating arrangement, then it doesn't have to be as wide as mercury.

I never proposed using a Mercury capsule with a conventional heat shield because neither the capsule geometry nor its seating orientation are required for a Mars landing. As far as your physics lesson is concerned, in real life the IRVE payload container had a vertical orientation. In real life, an inflatable heat shield can be deployed from the base of the service module and would permit a much smaller diameter capsule and therefore lighter capsule. Most of the capsule's mass, even after expenditure of propellant, is in the base.

RobertDyck · 2015-03-27 22:20:02

Vertical. How many G's do you envision astronauts enduring? Perhaps GW can help us with G-load for a realistic Mars descent.

As for MOOSE, that is the only way you can reduce the mass for a single person re-entry capsule for Earth. Without MOOSE, you get Mercury. MOOSE is inspired by a book: "Starship Troopers" written by Robert A. Heinlein. Excellent book! You should read it. Some back-story: after the book "Stranger in a Strange Land", Mr. Heinlein found hippies making a pilgrimage to his house. He was an ex-navy man, so quite disconcerted by this. His rebellion to the hippie movement was "Starship Troopers". The idea of powered armour was stolen from Doc Smith's books, the Lensman series. He was quite open and honest about this. But "Starship Troopers" had a unique Robert A. Heinlein flavour. Great book! I thought this would be ideal for Hollywood. Lots of action. Lots of shooting monsters. But when they did make a movie, they ruined it! Don't judge by the movie, read the book.

This relates to MOOSE because that book included "drop capsules". Paratroopers drop from a ship in orbit. They had a minimum entry capsule, parachute, and the capsule would disintegrate as soon as the parachute opened, creating a cloud of chaff. That left the paratrooper dangling by a parachute for the landing. In the book, the armour included a rocket pack. This rocket pack automatically activated when close to ground, cutting parachute cords, landing the individual trooper with propulsive landing.

So your argument is Mars allows different physics. Let's see what GW has to say. Just to stir up GW, the ADEPT team claims using a parachute and sub-sonic propulsive landing has lower total mass. That's for 40 metric tonne landed payload. What does it take to do that? How many G's during thermal atmospheric entry? On Mars.

GW Johnson · 2015-03-28 09:17:06

Bear in mind that my spreadsheet "model" uses the 1953-vintage scale-height approximation originally developed for warhead entry, and is inherently 2-D. It's crude at best. It's been 2 years since I last ran it.

The last model I ran was 212 kg/sq.m ballistic coefficient at entry angle angle 1.63 deg below horizontal, and entry interface speed 3.47 km/s at 135 km. That object came out of hypersonics at an inherently-overpredicted altitude of 22 km. Peak gees was just over 0.7. With a rather blunt nose radius of about 8 m (appropriate to a conical capsule shape flying blunt side forward), I got a peak heat rate of only 2.34 W/sq.cm.

Both the gees and the heating are actually quite modest. Although a pulse of 0.7 gees to a person weakened by 8-9 months in zero-gee might be rather tough on someone sitting upright. For someone 1-gee "fit" from artificial gravity on the trip, a 0.7 gee pulse on entry is nothing. This would really require some modest lift to maintain that shallow angle relative to local horizontal.

I'd be more inclined to think the real end-of-hypersonics would be down closer to 5-10 km at a much steeper downward angle. But the other numbers are likely more-or-less representative.

As for the craft's "design", it was nothing but a conical capsule with retro-propulsion engines built in. I didn't use any fancy new technologies like flexible or foldable or inflatable heat shields. I did some high-school physics kinematics for my rocket-brake landings. It's all very crude, just trying to get into the right ballpark. I don't have any real trajectory code for this. (I guess I could try to write a simple one.)

GW

kbd512 · 2015-03-28 12:24:17

A micro capsule equipped with an inflatable heat shield with the same attitude control mechanism that IRVE uses would decrease rate of descent and the thin Martian atmosphere would decrease peak deceleration.

http://ntrs.nasa.gov/archive/nasa/casi. … 012170.pdf

The high deceleration forces imparted to the test articles has everything to do with deliberate use of plunging descents on suborbital flights to produce maximum heating and aerodynamic pressure on the shield for testing purposes.

http://solarsystem.nasa.gov/docs/p484.pdf

If I thought the micro capsule concept had no merit or posed significant technical challenges, I would not have proposed it.

It's a relatively simple solution to a complex problem. The primary disadvantage being that the micro capsule leaves nothing but a water ration and spare oxygen tank or two for the explorer to use to reach the habitat module (which is mobile and comes to him or her in my proposals).

To be clear, what I desire is a Mars EDL solution for humans that could be developed in 6 years time or less, at a cost of $3B or less. This is relatively inexpensive and therefore workable. All the realistic multi-person EDL solutions I've seen proposed aren't even in the same ballpark. Most involve a decade or more of development time and projected costs of around $10B or more. In other words, a human rated lander would be within NASA's budget using my proposal but a multi-person lander would take so long to develop or cost so much that it would most likely be cancelled.

If you disagree with my micro capsule proposal, then counter with a realistic multi-person lander with a development time of 6 years or less, costing $3B or less, that significantly improves the situation of our explorers if they happen to land off course. An all-propulsive EDL using Red Dragon would obviously function correctly, but it lands so little payload that the astronauts contend with the exact same surface survivability issues that astronauts landed with the micro capsule would have. Both solutions provide very little in the way of contingency supplies.

Apart from keeping the crew together, ensuring that they all die together if anything goes wrong (something that's obviously pretty high on the priority list of everyone but me), Red Dragon provides precious little advantage over my solution and would almost certainly cost more in terms of development and per unit purchase price. I like Red Dragon better than my proposal because I really want a cost effective multi-person lander, but it won't put the astronauts in any better a position if they land off course.

To complete the context of what I'm proposing, I would like two separate EDL solutions. I want to use a minimum mass and complexity solution like HIAD for humans and a somewhat more involved solution like ADEPT for heavy cargo.

If Red Dragon can be developed and properly tested for less money, then I would drop my proposal. My guess is that the weight and higher ballistic coefficient of the vehicle will make landings on Mars marginal at best.

Last edited by kbd512 (2015-03-28 12:26:14)

RobertDyck · 2015-03-29 13:48:12

GW Johnson wrote:

Peak gees was just over 0.7.
...
nothing but a conical capsule with retro-propulsion engines built in. I didn't use any fancy new technologies like flexible or foldable or inflatable heat shields.

That may explain why you keep raising concern about coming out of hypesonics too low. What happens if you use something like ADEPT? That's a carbon fibre foldable heat shield. The goal is to catch more air when entering Mars thin atmosphere. Would that exit hypersonic flight high enough for a parachute to be useful? What G-load would that put on astronauts?

SpaceNut · 2015-03-29 17:55:38

Something like this product Lava Shield

Made from crushed volcanic rock, our Lava Shield Mat has similar benefits as our HP Heatshield Mat with that trick carbon fiber look! This insulator helps to reflect and dissipate heat making it perfect to be an air box heat shield or a heat shield barrier to protect the paint on your carbon fiber/fiberglass hood. Due to basalt’s (mineral from volcanic rock) natural chemical and acid resistant properties, Lava Shield is excellent for use in harsh environments Our Lava Shield may be used on a fire-wall to shield radiant heat it is also perfect to shield heat in transmission tunnels. This heat shield barrier cloth, can be used with less than 1” of airspace, 3/8” minimum is recommended. It is also better to use Lava Shield when there are instances of little air flow, making it a superior insulator to aluminum and gold barriers fabrics in these occurrences. Bulk rolls of basalt heat shield cloth are available. Our Lava Mat withstands 1200°F direct continuous and 2000°F intermittent.

Now something for Adept: MECHANICALLY-DEPLOYED HYPERSONIC DECELERATOR AND CONFORMAL ABLATOR TECHNOLOGIES FOR MARS MISSIONS.

The pages have just what we are looking for in ballistic angle and G force....

kbd512 · 2015-03-29 19:50:49

SpaceNut wrote:

The pages have just what we are looking for in ballistic angle and G force....

SpaceNut,

The document in your last post in this thread shows the peak deceleration force that the technology would likely produce if deployed at Venus with the specified area and loading parameters. How does that help us understand what the peak deceleration would be on Mars?

Let's have a look at the following document for more information about peak deceleration forces from actual Mars missions:

http://solarsystem.nasa.gov/docs/4_17WELLS.pdf

Last edited by kbd512 (2015-03-29 19:52:38)

kbd512 · 2015-03-29 20:15:47

RobertDyck wrote:

That may explain why you keep raising concern about coming out of hypesonics too low. What happens if you use something like ADEPT? That's a carbon fibre foldable heat shield. The goal is to catch more air when entering Mars thin atmosphere. Would that exit hypersonic flight high enough for a parachute to be useful? What G-load would that put on astronauts?

According to JPL, ADEPT would permit you to deploy a supersonic parachute at a high enough altitude on Mars to produce useful deceleration. However, the mass of the payloads we're talking about landing is so high that I think supersonic retro-propulsion is a better method for supersonic and subsonic deceleration and landing, even if it takes a little more propellant to do it. I like approach #3 in their infographics.

Peak G's for all proposed ADEPT control strategies is less than 4. ADEPT and HIAD shift the payload mass to "catch more air", as you put it.

SpaceNut · 2015-03-29 22:35:12

The chart yes is for Venus kbd512 but the attack angle and G's are still there in the up levels of the chart as they are the same for Mars its only as we approach the 75,000 ft level that we are in an atmosphere as thick as earths....

http://www.lpi.usra.edu/opag/jan2014/pr … uchamp.pdf page 16 starts the woven materials

Ablative Thermal Protection Systems

kbd512 · 2015-03-30 00:35:17

SpaceNut,

Without an altitude at which peak G's are experienced, how can we determine that the reentry peak deceleration force on Venus would be the same as it is on Mars? Do you have an atmospheric density model for Venus? MSL reentry started around 125km at 5.9 km/s, experiencing a peak deceleration force of 12.9g, and it was ~127kg/m^2. One of the hypothetical sphere cone reentry vehicles in Fig 2 experienced 32g at an unspecified altitude and was 44kg/m^2. Between 200km and 75km in altitude, is the density of the atmosphere on Venus an analog for Mars? What am I missing?

http://www.spaceflight101.com/msl-landing-special.html

Last edited by kbd512 (2015-03-30 00:37:54)

RobertDyck · 2015-03-30 04:24:49

From the web page that kbd512 just linked. Note that MSL did not use a deployable or inflatable heat shield, just a normal aeroshell.

Atmospheric Entry
This portion of the Mission begins at the Point of Entry Interface and ends with Parachute Deployment. The Atmospheric Entry itself consists of four different phases.
Once reaching and detecting Entry Interface, the Guidance System is switched to Entry Mode and the Reaction Control System is Pressurized. As soon as Entry begins, Guidance Computers start to modulate the lift vector by rotating the vehicle to achieve the required angle of attack to reach the desired downrange and cross range target for Parachute Deployment. The Entry Controller of the vehicle generates roll, pitch and yaw torque commands based on real-time navigation data. These commands are translated to on/off commands for the Reaction Control System that consists of eight thrusters installed in pairs outside the aeroshell. At the Point of Entry Interface, MSL is at an Altitude of 125 Kilometers above the Martian Surface (131.1km above Gale Crater) traveling at a velocity of 5,900 meters per second.
Once Entry Guidance is active, the Guided Entry Portion of the Process begins with MSL holding its pre-bank attitude until sensing 0.5G (Earth Gs). This marks the start of the Range Control Phase during which the bank angle is modified in a way to minimize the predicted downrange error at parachute deployment. Bank reversals are conducted as required with cross-range error margins being maintained at a manageable level. During this Guidance Phase, MSL experiences Peak Heating at about EI+85 Seconds. At that point, the PICA Heatshield of the vehicle has to withstand a thermal load of more than 5,700J/cm² with a Turbulent Air Flow around the Entry Vehicle. Peak Heat Rate will be greater than 200W/cm². Eleven Seconds after Peak Heating, the Vehicle passes Maximum Deceleration experiencing a G Force of approximately 12.9. Once MSL’s Velocity drops below 900m/s, the Heading Alignment Phase of the EDL Process starts. Residual Cross-Range Error is minimized during this phase with bank angles being adjusted so that the vehicle achieves a direct-flyover of the Parachute Deploy Target. Approximately 15 Seconds prior to Parachute Deployment at about EI+230 Seconds, six 25-Kilogram weights mounted on the inner portion of the Aeroshell are jettisoned with two-second intervals to eliminate the Center of Gravity Offset. This maneuver is known as SUFR – Straighten Up and Fly Right. Also part of the maneuver is a 180-degree azimuth turn of the vehicle to align the Terminal Descent Sensor for proper Ground Acquisition. SUFR is velocity triggered. Transients occurring due to SUFR will be dampened by the Reaction Control System and the Angle of Attack will be reduced to nearly zero.

Note: Peak deceleration is 12.9 Gs. There is no way a human can remain conscious while standing or sitting in a vertical orientation at that G-load. This is one major reason why entry capsules are designed as they are. I saw a documentary about G-suits for fighter pilots. Most fighter pilots can only handle a maximum of 9 Gs. Many pilots wash out because that can't handle even that. One new G-suit compressed the pilots chest and upper abdomen, not just legs and lower abdomen. The test pilot was able to withstand 12 Gs. That's only one guy; and MSL experienced a peak of 12.9 Gs, which is above that. Don't expect astronauts to remain conscious at 12.9 Gs. This is with a normal aeroshell. What is the peak G-load with a deployable or inflatable? I doubt it will be below 4 Gs, considering MSL's peak. So entry vehicle design is back to lying down, like Mercury.

Last edited by RobertDyck (2015-03-30 08:10:47)

GW Johnson · 2015-03-30 08:06:15

It's the entry angle that is causing all the confusion. I used a very shallow entry angle from LMO to minimize retro burn requirements, essentially converting circular to a surface-grazing ellipse. By the time I reached 135 km "interface" from a 200 km orbit, my speed was up to 3.7 km/s, but the local angle below horizontal was only 1.63 degrees. To come in steeper at higher gee from orbit, you need a bigger deorbit burn (expensive!). Or, you need to be coming in from deep space, which frees you to enter at any angle you want.

Here on Earth, the air is thicker, and so you hit higher deceleration much higher up. Steeper angles work fine, except for the gees! Steeper is what the warheads use, by the way. Returning from the moon, they had to hit at about 2 degrees, give or take a fraction, to limit deceleration to 11 gees in Apollo. In Earth's thicker air, you come out of hypersonics and pop your drogue chute very high up: 30+ km, if memory serves.

You cannot come in steeper on Mars unless your ballistic coefficient is very, very low (the way the probes do for entry from deep space) because the air is so thin. Drag cannot build up and slow you down until you are too low. That is what the JPL guys are talking about when they say that 1 ton articles give them problems designing for landing. They've been using the same aeroshell shape for some years now. That's also what the extendible, inflatable, and flexible heat shield technologies are all about: means to reduce ballistic coefficient well under 50 kg/sq.m.

Venus has a thicker atmosphere. Gees will build up quicker than here, so entries must be shallower-angle to limit gees. You can get this effect higher up in the thinner air if your ballistic coefficient is low. If it's higher, you will penetrate deeper before drag has a chance to decelerate you, and you will see more peak gee. Same as here, just a bit worse, that's all.

GW

RobertDyck · 2015-03-30 08:38:22

I saw a presentation at a Mars Society convention. I think it was by Gary Fisher. He suggested parking a Mars orbiter in highly elliptical, high Mars orbit. I mean extremely elliptical. Periapsis just barely above the atmosphere, because orbit would be achieved by aerocapture. Apoapsis so high that the craft is barely in orbit about Mars. That means just a tiny nudge is necessary to depart orbit. When the lander falls from apoapsis, how much propellant is necessary to achieve the entry angle you're talking about? Wouldn't a tiny nudge at the highest point of elliptical orbit make a great change in atmospheric entry angle?

GW Johnson · 2015-03-30 15:41:03

A small nudge at apoapsis (high point) makes a change in periapsis (low point) altitude. The entry angle would change, yes, but only by a very little. With an elliptical orbit, all entry angles would still be steep, assuming a surface-grazing orbit. With a circular orbit at low altitude, the entry angles are inherently very much lower. Make a scale plot and get your protractor out. You'll see what I mean.

GW

kbd512 · 2015-03-30 22:17:23

Rob,

Moreover, it wouldn't matter if an astronaut was flat on his or her back at 12.9 G's, he or she would be unconscious either way. Perhaps I'm oversimplifying things, but it seems as if JPL has two problems to solve here.

P1. In order to decelerate fast enough for supersonic parachutes or retro-propulsion to function properly, our reentry vehicle has to bleed away enough of its initial velocity at an altitude high enough above the surface of the planet to permit effective use of parachutes and/or retrorockets. This generally means a shallow entry angle that will increase peak heating.

S1: Develop lightweight materials capable of withstanding higher heating and reentry vehicle shapes or structures that lower the ballistic coefficient.

P2. In order to reduce peak deceleration force to levels tolerable by humans, a lower rate of descent is required. This generally means producing lift to counter gravity, thereby decreasing the rate of descent.

S2: Use the reentry vehicle's shape and center of gravity manipulation to produce lift.

MSL's reentry angle was around 20 degrees to reduce peak heating. The combination of the weight of the vehicle, relatively small lifting body, and poor lift generation created a plunging descent, increasing peak deceleration force. JPL and NASA favor plunging descents for unmanned reentry vehicles subject to substantial peak heating, as unmanned vehicles can readily be designed to survive deceleration forces that would injure or kill humans.

The micro capsules I proposed would reenter at shallow angles, have far lower ballistic coefficients than MSL, and generate more lift than MSL was capable of generating. You keep trying to invalidate this micro capsule concept using faulty logic that does not correlate with what NASA has published with regards to reentry. Even if what you stated applies to the micro capsule concept was true, which it is not, then it applies to an even greater degree to larger capsules with higher ballistic coefficients and lower lift generation capability.

A single person capsule can be made so light as to be capable of landing by parachute alone, especially if the service module and inflatable heat shield are discarded before landing. The same cannot be said for the multi-person capsules and heavier rovers or surface habitats.

RobertDyck · 2015-03-30 23:25:04

MSL already used PICA. Apollo used Avcoat, and Orion uses Avcoat. That's enough for direct entry into Earth's atmosphere from the Moon. Inspiration Mars planned to use Orion, with a free return trajectory that would fly by Venus before returning to Earth. That implies Avcoat is good enough for that. Hopefully good enough. But in 1970, when Russia talked about going to Mars to trump Apollo 11, NASA determined Avcoat is not good enough for direct entry from a trajectory directly from Mars. So NASA developed PICA, specifically for that. Since PICA can protect a capsule with humans directly entering Earth's thick atmosphere when returning from Mars, a trajectory that falls down in the Sun's gravity the whole way, then it should be able to handle direct entry into Mars atmosphere. Dragon uses PICA-X.

To reduce speed quickly enough to deploy parachutes at an altitude to be useful, Mars Direct proposed a deployable carbon fibre heat shield. ADEPT is that deployable carbon fibre heat shield.

We need lower the ballistic coefficient to decelerate at higher altitude, so spread deceleration over longer duration in order to reduce peak deceleration. That's accomplished by the deployable heat shield, aka ADEPT.

We also need to enter Mars orbit first, in order to reduce speed before final plunge into Mars atmosphere. Aerocapture using ADEPT.

Sounds like it's handled. We will need to test ADEPT on Mars. And test/demonstrate aerocapture with an orbiter.

Your argument about microcapsules does not hold. I already gave you a basic grade school description of surface area vs volume. This means more hull weight per person, not less. The issue of higher ballistic coefficient due to lower aeroshell surface area is resolved by a deployable heat shield. In fact, you have lower mass per person with a single capsule, not greater. You talk of using an inflatable heat shield with the microcapsule, but I do not see how that is any different than a deployable heat shield with a single capsule for the entire crew. An inflatable has size limit, beyond a certain size the inflatable will bend/kink, losing shape. That's one reason to use ADEPT for the full-size capsule.

As for landing by parachute alone, you realize your microcapsule would be as large as MSL. It required skycrane. MER rovers (Spirit & Opportunity) were smaller, but used air bags to cushion landing. Their peak acceleration on impact with the ground was 35 Gs. And that's with airbags; parachute alone would experience even higher acceleration. Don't expect an astronaut to survive that. Sorry, but Mars atmosphere is just too thin. You need rockets and landing legs with shock absorbers like Viking, Phoenix, or the Apollo LM.

kbd512 · 2015-03-30 23:53:13

Rob,

Dragon's problem with reentry on Mars doesn't have anything to do with its heat shield material, it has everything to do with its weight and ballistic coefficient. It doesn't land enough mass to be anything more than a more expensive flying coffin than a micro capsule would be, if either land substantially off course.

Your second argument was that this thing will experience a higher peak deceleration force than a sitting astronaut can tolerate. That's been pretty thoroughly debunked by GW and the documentation by NASA regarding peak reentry deceleration forces encountered on Mars from actual missions. MSL encountered a greater peak declaration force for the reasons I already noted.

Your basic grade school education about area and volume ignores the fact that Red Dragon is already substantially more massive than 4 micro capsules would be because Red Dragon lands propulsively.

Regarding how inflatable heat shields are "different" from using deployable articulating heat shields for the entire crew, the only multi-person capsule that could potentially land on Mars would have to be substantially redesigned to use this novel new ADEPT technology. I don't care if NASA does that, and it would be my preference, but so far there's no talk of doing that.

Regarding MER, it's not a valid comparison. The lander weighed 533kg and it was dropped from a height of 10M onto the surface of Mars. The capsule plus astronaut won't be anywhere near 533kg and we're not dropping it from 10M with the astronaut in it.

RobertDyck · 2015-03-31 00:24:22

I'm sorry you are having difficulty reading what I wrote. I won't repeat myself. What you day about Dragon shows you didn't read what I wrote. Read it.

And I said Dragon is appropriate to return crew to Earth. The heat shield material is appropriate for Mars. A custom lander is required for Mars. Dragon can be used either by Mars Direct as the ERV capsule, or as an emergency escape pod for a reusable ITV during aerocapture at Earth. Dragon is the not the Mars lander. I never said it is. You are obsessing about Red Dragon from the Mars One mission plan. I don't know anyone here who believes that's a good plan. MIT did an analysis, they believe settlers would survive 2 weeks after arriving on Mars. Yes, a multi-person capsule does have to be substantially redesigned to use ADEPT. One option is the Mars Direct habitat (hab). I came up with a mission plan to use Russia's Energia, so suggested a minimum size landing capsule and all-inflatable surface habitats. That plan was developed in 1999-2002 when Russia was behaving itself. The idea of using Energia was from Robert Zubrin's book "The Case For Mars". As Dr. Zubrin said in his book, using Energia requires splitting the mission into 3 parts instead of 2, because Energia has 2/3 the lift capacity of Saturn V / Ares / SLS. Doing it today with SLS might permit a hard wall habitat. But that's all custom landers. At no time did I ever claim that Dragon was the Mars lander. What I do say is Dragon can be used for return, and Dragon is overall a much better choice than Orion. That's because Dragon has a dry mass of 4.3 metric tonnes, with full fuel tanks (wet mass) is 8.0 metric tonnes. Orion is much heavier, with fairing and LES and full tanks it's 28 metric tonnes. I've said this many times. That doesn't mean I claim Dragon is the crew lander.

As for MER, yes it is a fair comparison. Don't you think they would have landed with parachutes if that would have worked?

kbd512 · 2015-03-31 00:33:07

Rob,

I'm tired of arguing the point with you. You keep using arguments that apply equally to any EDL solution or are demonstrably false. When that doesn't work, you propose something that I did not in order to construct something with weight/dimensions/reentry profile/etc that won't work on Mars without lots of expensive, complicated, and failure prone technology.

The micro capsule is an unpressurized aluminum trash can with a fabric seat bungee corded to the walls and a service module attached to its rear end with an inflatable or deployable heat shield (I don't care which) and reaction control system for de-orbit. Anyone with the desire and enough time on their hands can engineer it into a Rube Goldberg contraption that won't work reliably on Mars, but you have to work overtime to do it.

I wanted to keep EDL for humans on Mars spectacularly simple, which would be a first for anything designed by NASA, but some people apparently have complexity cravings that rival NASA's. Multi-person landers are an admirable goal, something I would love for us to build, but there's simply no funding to do that because Congress forced us to squander so much of our available funding on Orion and SLS rather than a spacecraft that could actually land on something other than Earth or a rocket that we could afford to operate. Unfortunately for us, funding hasn't been coming out the wazoo lately.

kbd512 · 2015-03-31 01:01:16

RobertDyck wrote:

And I said Dragon is appropriate to return crew to Earth. The heat shield material is appropriate for Mars. A custom lander is required for Mars. Dragon can be used either by Mars Direct as the ERV capsule, or as an emergency escape pod for a reusable ITV during aerocapture at Earth. Dragon is the not the Mars lander. I never said it is. You are obsessing about Red Dragon from the Mars One mission plan. I don't know anyone here who believes that's a good plan. MIT did an analysis, they believe settlers would survive 2 weeks after arriving on Mars. Yes, a multi-person capsule does have to be substantially redesigned to use ADEPT. One option is the Mars Direct habitat (hab). I came up with a mission plan to use Russia's Energia, so suggested a minimum size landing capsule and all-inflatable surface habitats. That plan was developed in 1999-2002 when Russia was behaving itself. The idea of using Energia was from Robert Zubrin's book "The Case For Mars". As Dr. Zubrin said in his book, using Energia requires splitting the mission into 3 parts instead of 2, because Energia has 2/3 the lift capacity of Saturn V / Ares / SLS. Doing it today with SLS might permit a hard wall habitat. But that's all custom landers. At no time did I ever claim that Dragon was the Mars lander. What I do say is Dragon can be used for return, and Dragon is overall a much better choice than Orion. That's because Dragon has a dry mass of 4.3 metric tonnes, with full fuel tanks (wet mass) is 8.0 metric tonnes. Orion is much heavier, with fairing and LES and full tanks it's 28 metric tonnes. I've said this many times. That doesn't mean I claim Dragon is the crew lander.

What is your proposal for this custom multi-person lander that we don't have money to develop, Rob? Please share your proposal. When do you think we'd have money for that? Around 2030, maybe?

Dragon's going to be an ERV? Got it. Now we just need to land a rocket big enough to launch it off the surface of Mars. What would be easier to design and test on Mars, rockets capable of lifting 1t to LMO or 8t to LMO?

I'm obsessing over Mars One? Find a post of mine, except this one, where I mention it. I've never seen any of their proposals, so I'll have to take your word for it.

RobertDyck wrote:

As for MER, yes it is a fair comparison. Don't you think they would have landed with parachutes if that would have worked?

The retro rockets on MER brought the lander to a full stop 10M above the surface and then dropped the lander onto the surface. The peak deceleration that you referenced was the result of the lander slamming into the ground after it was dropped. MER weighed 533kg. Micro capsules will be significantly lighter.

RobertDyck · 2015-03-31 12:23:30

The following discussion thread was started by someone else, but it became about my mission plan starting with page 2. My architecture uses a light-weight Mars Ascent vehicle, a reusable ITV that goes from ISS to Mars orbit and back, and has Dragon attached to the ITV as an escape pod.
Click here: Yet another Mars architecture

I discussed updating Mars Direct. That means sticking with Robert Zubrin's mission architecture, not mine, but building it with current equipment. Robert Zubrin's ERV includes a capsule, and a 2-stage rocket to throw directly from Mars surface to trans-Earth trajectory. That 2-stage rocket would use LCH4/LOX engines, and would be a new rocket. One point I made is rather than develop a new capsule, use Dragon. However, to survive 6 months you would need a small, light-weight module attached to the nose with recycling life support. I suggested the same life support as ISS, and just barely enough room for one person to float in the centre to do repairs/maintenance. Base the hull of that module on Cygnus, but smaller. That module would be discarded before entering Earth's atmosphere.
Click here: Light weight nuclear reactor, updating Mars Direct

And I argue that NASA has enough money to do all this right now. They don't need more money. It would require cancelling Orion and ARM, complete SLS, give up on the Moon and Constellation, focus exclusively on Mars.

kbd512 · 2015-04-01 05:30:35

RobertDyck wrote:

I discussed updating Mars Direct. That means sticking with Robert Zubrin's mission architecture, not mine, but building it with current equipment. Robert Zubrin's ERV includes a capsule, and a 2-stage rocket to throw directly from Mars surface to trans-Earth trajectory. That 2-stage rocket would use LCH4/LOX engines, and would be a new rocket. One point I made is rather than develop a new capsule, use Dragon. However, to survive 6 months you would need a small, light-weight module attached to the nose with recycling life support. I suggested the same life support as ISS, and just barely enough room for one person to float in the centre to do repairs/maintenance. Base the hull of that module on Cygnus, but smaller. That module would be discarded before entering Earth's atmosphere.

Using a minimum mass ERV sounds expensive, difficult, and potentially far more hazardous than an off-course landing on Mars.

The capsule I proposed is for LMO to Mars and Mars to LMO only, rather than any type of interplanetary transfer.

RobertDyck wrote:

And I argue that NASA has enough money to do all this right now. They don't need more money. It would require cancelling Orion and ARM, complete SLS, give up on the Moon and Constellation, focus exclusively on Mars.

I like the fact that the plan involves going to Mars instead of playing with space rocks, but I also think that what you're proposing would cost every bit as much as Orion's development and then some. Dragon, like Orion, is incapable of sustaining the astronauts for the period of time they're in deep space between Mars and Earth and docking a sustainment module to the front of the capsule doesn't make it a deep space habitat. If you're sending your explorers to Mars in a ITV/MTV anyway, why not bring them back in that vehicle? We now have reliable propulsion capable of doing that without blowing mass budgets and therefore breaking the bank.

Even if we complete SLS, what does SLS do for us except suck down a few billion in development and operations each year? It's entirely wrong for any type of sustainable Mars exploration campaign.

With NASA's current budget, it can't spend a few billion on ISS, a few billion on SLS, and a few billion on Orion every year and have funding remaining for closed loop ECLSS, deep space habitats, landers, and ISRU. Two of the three would need to be cancelled if we ever want to go to Mars. ISS has the potential to have real utility to assist with that goal, whereas an unaffordable rocket and an insanely heavy capsule don't.

RobertDyck · 2015-04-01 07:41:34

I'm sorry, you're still having difficulty understanding what I said. When I first gave a presentation about my mission plan, it was the 2002 Mars Society convention. Most of the members that said they wanted Mars Direct. Only Mars Direct. Nothing but Mars Direct. They didn't want to hear anything else. So what I've posted on this forum is two separate mission plans. They are not integrated, they don't work together, you have to pick one or the other. But these are two options.

One is Robert Zubrin's Mars Direct. It's just updated. When he and his partner David Baker came up with Mars Direct in late 1989 and early 1990, the military was still working on the SP-100 nuclear reactor for space. In fact, development for SP-100 was completed in 1992. So they used what was current at that time. Since then the SAFE-400 reactor has been developed, completed in 2007, same purpose and same amount of electricity, but lower mass. So I propose using the newer one. And Mars Direct used a capsule to return astronauts to Earth. Rather than "re-invent the wheel", I proposed using Dragon for that purpose. Dragon does not have life support for the 6-month journey, so will need a life support module added. But this is a 3-module design, something that has proven to be efficient. When Apollo began development, contractors bid to design and build the spacecraft. GE (General Electric) was one of the bidders. Their proposal for the Apollo CSM in 1960 was Apollo D2. It was a 3-module design: entry capsule, service module, plus mission module with life support and equipment for the mission. So what I'm saying is the Mars Direct ERV would also have multiple modules. In this case, 2 stage launch vehicle to lift off from Mars and inject the capsule into a trans-Earth trajectory, plus capsule for Earth EDL, plus life support module for the journey. Technically that's 4-modules, but one of the propulsion stages would be discarded during ascent from Mars. And Mars Direct called for developing recycling life support, but I suggested using the same life support system as ISS. That has already been demonstrated in space. Actually, life support would require a few improvements, but should be tested/demonstrated on ISS first.

The other mission architecture is mine. I called my mission "Mars Orbit Rendezvous". However, years ago one member who posted on this forum asked Dr. Zubrin about it, and he reported that Dr. Zubrin called my plan "Hybrid Direct". Ok, so Dr. Zubrin himself gave my plan a name. I'm flattered. This plan has a reusable Interplanetary Transit Vehicle (ITV). I borrowed that name from NASA's DRM. My plan has a light-weight Mars Ascent Vehicle to return astronauts to Mars orbit, rendezvous and dock with the ITV. The ITV uses aerocapture at both planets. Because it uses aerocapture to enter Earth orbit, that manoeuvre has risk. To mitigate risk, the ITV has an escape pod. If aerocapture in Earth's atmosphere goes wrong, astronauts can "punch out" in an escape pod, and land safely on Earth. But it's just an escape pod. Rather than re-invent the wheel, in 2002 I said use a pair of Soyuz spacecraft without their orbital modules. However, now we have Dragon, so use Dragon as the escape pod. But that's only for emergency. If everything goes according to plan, Dragon will not be used. The ITV will enter Earth orbit, the use aerobraking to drop orbit, rendezvous with ISS, and dock. Then any spacecraft that can return crew from ISS can return the Mars astronauts to Earth. If using a reusable craft to return crew, in 2002 that would have been Shuttle, then Dragon would stay attached for the next mission to Mars. If using an expendable craft, then astronauts may as well use Dragon, then the second mission would start with a new Dragon. That is crew would ride dragon to ISS, then that Dragon would be attached to the ITV before departing for Mars.

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