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I'm still unsure what dry mass 4200 kg Dragon v1 and dry mass 6400 kg Dragon v2 really mean. That has to include the two trunks, and those cannot be the same. The difference is the Super Dracos, the landing legs, and the revised nose caps, plus an evident difference in propellant capacity that is listed as about 1200 kg for v1 and 1800 kg for v2, and likely a slightly-thicker variant of the heat shield. And there's nothing about the dry mass difference with the extended trunk option on v1.
It is clear Spacex takes pains to conceal such detailed information from the likes of enthusiasts such as us. I did see one reference to a v1 parachute test article that massed 5400 kg. That would be trunkless and without nose cap, and likely with most of the propellant gone. I've seen nothing for how much the nose caps weigh.
I also ran across a specific gravity of PICA-X as 0.30. A flat disk of this stuff 3.7 m dia and the listed 8 cm thick would mass just about 258 kg. The dia is actually slightly less and the shape is not really flat, so roughly 260 kg is probably a decent estimate of the v2 heat shield weight adequate for a from-Mars free-return.
That's the best I've got so far. Not good enough to generate reliable weight statements yet.
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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From the Dragon wiki
SpaceX completed a parachute drop test to validate the new parachute design. This involved carrying a 5,400 kilograms (12,000 lb) Dragon test article by helicopter to an altitude of 2,400 meters (8,000 ft) above the Pacific Ocean. 2,500 kg of return pressurized cargo, driven by parachute limitations.
That would be the capsule only with a simulated down mass payload capability.....
The Draco escape testing
During this test, the Dragon used its abort engines to launch away from a test stand at Launch Complex 40. It traveled to an altitude of 1,187 meters (3,894 ft), separated from its trunk, deployed its drogue parachutes and then the main parachutes. It splashed down into the ocean and was recovered. The vehicle was planned to reach an altitude of 1,500 meters (5,000 ft) but one of the engines underperformed due to an abnormal fuel mixture ratio.
Same for this test....
The manned dragon would be greater due to the 30 day limits for what a crew would consume for oxygen and water, which eats into the payload but also makes the capsule heavier.
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bump
Sure would be nice to be able to do this before nasa could....
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The modified dragon cargo for use as a service module docked to a drewed Dragon seems quite reasonable for a flyby mission. Other optional service module units are the European ATV currently being altered for Orion use, Cygnus from ATK/Orbital is another possible and finally the Japanese HTV is another which could be alterred to form this function. The 1 off mission is to redux just in time for the anniversary date which is fast approaching.
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So what do you think, should NASA sell lottery tickets for an Apollo 8 style mission. I don't really know what good professional astronauts would do by sitting in a capsule and orbiting the Moon. They could read the book of Genesis, or they could snap photos of the Moon through their windows.
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How about a $50 million seat to take the ride around the moon to any one that wants to pay..The mission would pay for its self then.
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Many of the same issues are cropping up as we make what we do not have for the mission to work, such as the EDS, Service module ect... which brings up fuel type / engines and power sources.....
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For the EDS, should use the most powerful combination not requiring extreme cryogenic cooling; that means CH4/LOX. For the lunar orbit/deorbit & ERS: MMH + NTO. Keep things as simple as possible. I'm a big believer in the KISS principles.
Last edited by Oldfart1939 (2017-02-17 20:22:34)
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I found enough data to estimate dry trunk vs dry capsule weight for cargo Dragon, based on the chute drop test. The uncertainty is down to the mass of propellant residuals still on board at ocean splashdown, which I assumed to be zero. Cargo Dragon flies as capsule-plus-trunk until after re-entry burn. Between Spacex’s site and the associated Wikipedia articles, I worked out a typical weight statement for cargo Dragon, and rough-estimated the total theoretical delta-vee its propellant load could muster, knowing that about half of this will go toward attitude control, not maneuver delta-vee:
Capsule-only dry 2900 kg
Max cargo to ISS 3310 kg
Capsule-only burnout 6210 kg
Propellants 1290 kg
Capsule-only ignition 7500 kg
Empty trunk 1250 kg
Trunk cargo 0 kg
Loaded trunk 1250 kg
Nose cap 50 kg
Total to launch 8800 kg
Capsule-plus-trunk ignition 8750 kg
Less propellant 1290 kg
Capsule-plus-trunk burnout 7460 kg
Capsule-plus-trunk MR 1.1729
Capsule-plus-trunk dV 0.523 km/s
Remember that only about half the delta-vee is really available for rendezvous maneuver, because of attitude control. That might be around .25 effective km/s loss for attitude control. But I do not know how realistic that assumption is.
The nose cap mass of 50 kg is just a guess. The chute drop test was 5400 kg and the max-rated down-cargo mass is 2500 kg, and had no nose cap. If you assume no propellant residuals at landing, that difference gets you a capsule dry mass of 2900 kg. My dV is based on an assumed Isp in vacuum of 335 sec. Note that my dry mass figures add to what was on Wikipedia for dry mass: capsule 2900 kg + nose cap 50 kg + trunk empty 1250 kg = 4200 kg dry mass as from Wikipedia.
Crewed Dragon is operated the same way as capsule-plus-trunk until after the reentry burn. It does not shed the nose cap. The trunk is different, but one might assume roughly the same mass. The Wikipedia dry mass is 6200 kg. Subtract trunk empty assumed as 1250 kg, and get 5150 kg capsule dry mass.
The propellant loadout is higher at 1890 kg, and the capsule should have a higher dry mass because of the crew fitments, life support, and the big Super Draco thruster pods, each of which has two Super Draco engines and four Draco thrusters, all feeding from the same propellant supplies. If you assume the same 1250 kg for the different empty trunk, you can estimate a similar weight statement. I did this for 7 suited crew plus some cargo in the capsule, and nothing in the trunk.
Capsule-only dry 5150 kg
Crew/cargo in capsule 2800 kg
Capsule-only burnout 7950 kg
Propellants 1890 kg
Capsule-only ignition 9840 kg
Empty trunk (assumed) 1250 kg
Cargo in trunk 0 kg
Loaded trunk 1250 kg
Nose cap (not shed) 0 kg
Total to launch 11090 kg
Capsule-plus-trunk ignition 11090 kg
Less propellant 1890 kg
Capsule-plus-trunk burnout 9200 kg
Capsule-plus-trunk MR 1.2054
Capsule-plus-trunk dV 0.613 km/s
Assuming the same .25 km/s loss of delta-vee for attitude control, the remaining maneuver delta vee is around 0.36 km/s. I don't know how realistic that is.
I’m still looking at Red Dragon as a variant of crewed Dragon. No life support, no crew fitments, no chutes, a jettisoned nose cap, a slightly-lighter heat shield, but the same landing legs. It does its thing at Mars without the trunk, which is why capsule-only dry mass is so important to determine.
GW
Last edited by GW Johnson (2017-02-18 11:05:33)
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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OK, here’s what I have reverse-engineered about Red Dragon, building upon what I already did with crewed Dragon and cargo Dragon. By the way, DragonLab is just cargo Dragon with an unpressurized instrument bay that can opened to space and closed for reentry.
I’ve seen very little about Red Dragon other than some images of it setting down on Mars with landing legs, and a very sparse article on Wikipedia claiming mass at entry is around 7500 kg. But from the landing legs, it is quite obviously a variant of crewed Dragon. You delete the chutes, the life support, the crew seats and fitments, lighten the heat shield a bit, and swap out the retained nose cap for the jettisonable nose cap from cargo Dragon.
There’s no information, but it has to launch with a trunk, because that is its adapter to the rocket. It’ll need electricity not to freeze solid going to Mars, so that trunk needs solar panels that work in the increasingly-dim sunlight during the voyage. The larger arrays are on the cargo Dragon trunk, so I bet they use a version of that one. It would be jettisoned prior to direct entry at Mars from its interplanetary trajectory.
Here is how I tried to guess a realistic capsule-only dry weight:
Crewed Dragon est. dry mass 5150 kg
-delta heat shield 50 kg kg “educated guess”
-life support equipment 100 kg “WAG”
-crew seats and interior fitments 100 kg “WAG”
-chutes 100 kg “WAG”
-nose cone 50 kg “educated guess”
Est. Red Dragon dry mass 4750 kg
It’s still heavier than cargo Dragon, but that reflects the Super Draco engine pods. I got the lighter heat shield delta by thinning it from 8 cm to 6 cm thick, for a 3.7 m dia flat disk, at specific gravity ~ 0.3. Then I just rounded to the nearest 10 kg, because this is so crude.
I used these guesses to generate a weight statement similar to those I generated for cargo and crewed Dragons in the other posting just above. I did this for three levels of cargo aboard, to look at delta-vee capability versus load delivered to Mars. Earlier comments from Spacex were talking about 1, maybe at most 2, metric tons of payload delivered to Mars. Later comments talk about 2 to at most 4 tons. Just over 3 tons is where my calculated delta-vee capability matches a rough criterion I have been using for propulsive landings after aerobraking at Mars. Coincidence, perhaps, but encouraging that I might be in the ballpark. I ran 1, 2, and 3.2 metric tons delivered payload.
We are interested in the capsule-only weights and mass ratios for this landing, which takes place after the trunk has been jettisoned. That is substantially different than the capsule-plus-trunk mass ratios of interest for cargo and crewed Dragons. Do not cross-compare delta-vees directly!
Loadout min medium heaviest units
Capsule-only dry mass 4750 4750 4750 kg
Cargo on board 1000 2000 3200 kg
Capsule-only burnout 5750 6750 7950 kg
Propellant 1890 1890 1890 kg
Capsule-only ignition 7640 8640 9840 kg
Trunk dry mass 1250 1250 1250 kg
Trunk cargo 0 0 0 kg
Total loaded trunk 8890 9890 11090 kg
Nose cap 50 50 50 kg
Total to launch 8940 9940 11140 kg
Capsule-only MR 1.328696 1.280000 1.237736 none
Capsule-only delta-vee 0.934 0.811 0.701 km/s
Note that my 1-ton cargo mass at ignition (which is pretty much mass at entry) is 7640 kg, not very far at all from the 7500 kg listed in the Wikipedia article. Coincidence, perhaps, but it does suggest that I am well within the ballpark. The Wikipedia article also listed 1 metric ton as payload.
My criterion of 0.7 km/s is the bare minimum required to successfully land. It assumes no maneuvering is required to avoid obstacles on rough ground. This also takes no account of some of the propellant being used for attitude control instead of delta-vee. So this is very rough. But it does match fairly well with “the talk” about what Red Dragon can do.
It looks to me like a 2-3 ton payload could be feasible with the 1890 kg crewed Dragon propellant supply. There is some talk about increasing that propellant load, but I’m not at all sure where they would put it, unless they use a new inner pressure hull which has less interior volume.
My landing criterion of 0.7 km/s is also pretty rough. It is based on entering the atmosphere at a very shallow angle, and maintaining a tad of lift by angling the capsule axis a few degrees. At these ballistic coefficients, you penetrate quite deep before the “hypersonics are over” at rule-of-thumb local Mach 3. For the average speed of sound on Mars at about 5 km altitude, that’s a velocity of around 0.7 km/s. Which is what you have to “kill” before impact, only several seconds away. The Mach 3 rule-of-thumb comes from the definition of “hypersonic” for blunt objects.
Even so, it looks to me like one could send Red Dragon into some pretty rough landing zones, if one could provide a maneuvering autopilot and some obstacle detection equipment (radar?), and one reduces payload down into the 1-2 ton range. That gives you a propellant budget for hover and maneuver, even with the original propellant quantity of 1890 kg.
I’m still looking at landing Red Dragon vehicles on other destinations, such as asteroids, the outer planet moons, and or moon. But at 0.7 to 0.9 km/s, it doesn’t look good for a lot of them without some sort of propulsion stage. Without such a stage, the 11.5 ton Red Dragon is near the limit for what Falcon-Heavy can fling to Mars, based on the 13.6 metric tons listed on Spacex's website. The other destinations (except our moon) would have similar throw-weight limits.
GW
Last edited by GW Johnson (2017-02-18 12:33:29)
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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Here's the escape velocity and surface gravity for several destinations for one-way probes. Circular orbit velocities are 71% of escape. Something higher than escape is the landing delta vee required for all airless destinations.
Location escape speed surface gravity notes
Km/s gees
Mercury 4.17 0.365
Earth’s moon 2.37 0.165
Mars 5.03 0.381 aerobraking: 0.7 km/s
Io (Jupiter) 2.46 0.180
Europa 2.07 0.147
Ganymede 0.90 0.163
Callisto 2.01 0.0857
Titan (Saturn) 0.82 0.143 aerobraking: very low
Ceres (asteroid) 0.463 0.0295
Pallas 0.317 0.0216
Juno 0.161 0.0128
Vesta 0.261 0.0178
These figures suggest to me that Red Dragon as it is (without an additional propulsion stage) can visit Mars and Titan, and the asteroids, but not any of the outer planet moons except possibly Ganymede, and not our own moon. Crewed Dragon performance is similar but lower, if operated the same way.
GW
Last edited by GW Johnson (2017-02-18 12:41:37)
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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Great posts and thank you GW.
I started to look at what do we have for the engines as Oldfart1939 post #33 fuel types and was some what taken in that some of what we wanted was undertaken under the CEV Constellation that was cancelled.
Methane Lox Engines
http://www.astronautix.com/l/loxlch4.html
Raptor is not listed at the site and then we get into the Manned rated question if we are Nasa and not Space X....
https://en.wikipedia.org/wiki/Hypergolic_propellant
MMH + NTO
Aerojet Rocketdyne
http://www.rocket.com/propulsion-system … nt-rockets
http://www.projectrho.com/public_html/r … nelist.php
http://www.thespacerace.com/forum/index … pic=2583.0
Rocket Propulsion Elements a 764 page reference by Rocket GEORGE P. SUTTON
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A few other LCH4/LOX engines:
Aerojet Rocketdyne: Common Extensible Cryogenic Engine - current
ATK: ATK Successfully Test Fires 3,500 lbf Liquid Oxygen/Methane Rocket Engine in Vacuum - announcement of a test in December 2007. Don't know what they have now.
XCOR: XR5M15 - from the Astronautix website you linked.
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NASA has the most current experience with LOX/LH2, so upper stages should use LOX/LH2. The upper stage fires for a handful of minutes over the course of a handful of hours. Then it's just space garbage. There's no need to introduce a new, untested, lower-Isp propellant and engine combination for TLI.
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GW-
You may wish to revisit your source for the Ganymede escape velocity, which should be similar to that of Callisto; Ganymede is slightly larger (diameter), but slightly less dense, and only a tiny bit more massive. Their gravity relative to Earth is around 0.12 to 0.14 Earth G.
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Boeing is I think the only company making any upper stages that are fueled with Lox/LH2 which are the RL-10b engine.
http://www.astronautix.com/l/loxlh2.html
Notably the J-2 engine has come out of mothballs for serveral versions to which they were all not used in favor of others for the SLS.
https://engineering.purdue.edu/~propuls … quids.html
Nasa really has lost most of its edge with anything as they are not making anything for production but are contracting it all out.....
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Oldfart1939:
You are quite correct about the escape velocity for Ganymede. Titan is also wrong. I got these data from an old CRC handbook from 1972 and didn't notice the disparity between the listing in the reference and my back-calculated values. Everything else checks out, masses, radii, centrifugal force at 45 deg latitude. My surface gravity estimates are correct. The escape velocity estimates revise to 2.85 km/s for Ganymede and 2.58 km/s for Titan.
To land a Dragon on Ganymede will require a big rocket stage like the other big Jovian moons. Titan will not, because aerobraking is available in a dense atmosphere.
Here's something for Spacenut: current Spacex thinking for Red Dragon on Mars deletes the chutes on Mars entirely. Stripped out for weight savings as otherwise useless there. They plan on direct propulsive landing right out of the hypersonics, no chute at all. It's what they're already doing landing boosters. No chutes there, either.
Now that I've corrected the escape velocity data, what do y'all think of my weight statement and velocity capability analysis of the Dragon configurations?
Does anyone have an estimate of attitude control propellant budget versus delta-vee propellant budget for any of these configurations?
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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Yes no parachutes for the Red Dragon but for the crewed Dragon there is still parachutes slated for use I believe as it needs to dock to the ISS which the Red Dragon can not do.
I have found a few more articles for the Red Dragon Sample return topic to post...
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For crewed Dragon, the chutes are retained only as an emergency water landing backup to the no-chute propulsive landing on land. NASA is uncomfortable with this, and demands that the first few landings are chute landings in the ocean, like cargo Dragon. After that, Spacex is "free" to do propulsive landings on land with crews, assuming it works at all. This does require demonstration before risking crews.
It does not matter whether Red Dragon can dock at ISS, as the design concept calls for direct launch of Red Dragon to interplanetary trajectory to Mars on Falcon-Heavy, which has absolutely nothing to do with ISS. That's why they delete the chutes from Red Dragon, not to mention that at 7.5 metric tons entry mass, chutes are pretty much useless. As JPL will confirm (see Justus and Braun, which results I posted over at "exrocketman" some time ago).
GW
Last edited by GW Johnson (2017-02-21 18:05:45)
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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Why won't a forced inflation ring sail work? If HIAD and ADEPT slows the payload to just over supersonic speed, then an inflatable torus can unfurl a ring sail chute. It's just like a supersonic ring sail, except the outer portion / band gap is a CO2 or helium pressurized torus instead of normal chute fabric. If it works, it's absolutely the simplest and lightest option. It may even be possible to get rid of the mortar and pyro that MSL uses. Testing should not be terribly expensive, either. It could be packed into a sounding rocket, just like HIAD was.
If initial tests are successful, then use Falcon Heavy to send a Cygnus module equipped with HIAD + supersonic inflatable torus ring sail to Mars. If it lands in one piece, then the first consumables depot (water, solar panels, and batteries) is waiting on Mars for our initial exploration teams.
If it's not absolutely necessary for a particular mission evolution, then keep the crew separated from rockets and rocket fuel. We shouldn't need rockets to land on a planet with a usable atmosphere.
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Kbd512:
In a word, because you will reach deploy speeds closer to 5 km altitude than 25 km altitude. I really doubt that adept or hiad will slow you to Mach 1 at a reasonable altitude. Mach 2 is more likely, and that will be at the lower altitude nearer 5 km. That speed is pretty near 0.5 km/s.
Assume for the sake of argument that trajectory path angle at that point is 45 degrees downward. That means the path length to impact is 7 km. At 0.5 km/s undecelerated speed, impact is only 14 seconds away.
Just how in hell is a drag coefficient of around 1-ish going to supply significant deceleration in ~ 4+ sec (no matter how big Aref is !!!!), when you need around 10+ sec of rocket thrust to land, from just about any conceivable speed?
As I have often said, the "air" on Mars is just too damned thin for significant deceleration of objects >1 ton below just under Mach 1 (around 0.2+ km/s) no matter how long your deceleration interval is. The Justus and Braun EDL report confirms this (the EDL experts from JPL). Terminal velocities are just too damned high. You just haven't got the time budget for chutes or ballutes or anything else to do very much at all for you, if you exceed 1 ton at entry.
Surface density on Mars is around 7/1000 that of Earth at the surface, and very much lower still at altitude. It's time to face that ugly little fact.
That last is precisely why I think helicopters and fixed-wing aircraft designs have very little potential for general application on mars. There might be some very specialized exceptions, but these will not be craft that can take off or land without rocket boost.
GW
Last edited by GW Johnson (2017-02-22 19:39:28)
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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Adept umbrella heatshield is about as close as we can get to lowering the retro propulsion fuels to be used to land with.
How would one size up an Earth departure stage for sending basically 3 dragon truck masses plus supplies to survive to the moon and back for what would be a week to 2 weeks time duration?
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GW,
That's not what JPL's models for HIAD show. Mars Mach 2 is reached at 25km and subsonic at 15km. The DGB chutes can be deployed at Mach 2.5. Still impossible?
With respect to fixed wing or rotary wing aircraft, I agree. That won't work. That's why supersonic CO2 compressors require further investigation.
Edit:
An all-propulsive landing requires 20t IMLEO for every 1t of payload soft landed on Mars. HIAD or ADEPT lower that to 5t-6t IMLEO for every 1t of payload soft landed on Mars. Subsonic retro-propulsion burns are still required. If the retro-propulsion requirement is eliminated entirely with forced inflation chutes, then the IMLEO requirement drops even further.
Last edited by kbd512 (2017-02-22 22:05:46)
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Here's a link to the proposed manned EM-1 Mission option.
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There are too many acronyms being slung around here that only an insider would know. I have no clue what "IMLEO" means. Don't know what a "DGB chute" is either, except that it is some kind of parachute. I am ASSUMING that HIAD is the extendible heat shield, not the inflatable (I think, at least I hope, that ADEPT is the inflatable).
As for "That's not what JPL's models for HIAD show. Mars Mach 2 is reached at 25km and subsonic at 15km. The DGB chutes can be deployed at Mach 2.5. Still impossible?" -- I'll be more receptive when those models are validated by real tests and flight experience. Real data trumps calculation, always, even somebody with a track record as good as JPL's. Especially in an atmosphere with a density profile as variable as is Mars's. So far, there just isn't any such experience, other than an early exploratory shot or two with ADEPT.
I'm not at all sure about the tonnage figures either, as I have never, ever seen how such things are calculated. People lie with statistics all the time for a lot of reasons, and not all of them are politicians. Some are just trying to get their projects funded. I can't blame them for that, but you have to allow for the deception, and realize that an awful lot of published publicity stuff is, shall we say, "overstated".
Also, as for making SLS/Orion EM-1 manned, that depends on how much risk an administration can arm-twist NASA management into taking. Making Apollo 8 into a manned trip to lunar orbit and back was an enormous risk in 1968, when that was supposed to an unmanned flight test for the Saturn 5, which had already screwed-up non-fatally on the previous two flights.
Trump's tweets have more arm-twisting power than anything I've seen in ages, but NASA is so risk-averse after killing two shuttle crews with stupid management decisions, that I don't believe they would actually do it. They’ll drag their feet really hard, in any event.
Just doing rocket-equation estimates with bogus numbers, I sized a dangerously-minimalist ascent vehicle for 4 suited astronauts that could take them from Mars’s surface to low Mars orbit without any reserve for maneuver, and only 2 days' packed life support. There was room for only a few hundred kg of retrieved samples and equipment. It uses NTO-MMH storables. Ignition mass about 12 metric tons.
I added a descent stage to that along the lines of the Apollo LM, using the same propellants. The landing legs, frame/heat shield, tankage and engines totalled about 8 metric tons, for a combined mass at entry interface of about 20 metric tons. I looked at retropropulsive landing from M3 at 5 km, using a 1 km/s landing dV budget. No room for a rover! If they miss the designated landing point by more than walking distance in a suit, they are dead. Period. Which is why I personally HATE minimalist designs!
This thing had a heat shield, but no aero backshell or lateral heat resistant structure. The ride down would be risky at best. Your propellant tanks and other exposed structures would have to heat-sink their way through it. I did not add mass for that. RISKY!!!
Now, maybe, some sort of extendible or inflatable heat shield could reduce the ballistic coefficient and increase the M2 altitude, so that the use of chutes would be enabled for part of the descent, but you still need around 0.15 km/s worth of thruster burn in order not to crash from high subsonic. I have very serious doubts whether my 20 ton thing would be significantly smaller going that route.
So, the extendible heat shield and its structural stiffeners and operating equipment, plus a few hundred kg of chutes, plus a descent engine and some propellant, is that significantly lighter than the same descent engine and much more propellant? Good question! I only need about 6 times as much propellant in the retropropulsive version as in the final touchdown-only burn version. I only needed 8 additional tons for the landing stage.
20 tons is too big to fling there with Falcon-Heavy. I really doubt I could squeeze this below 13.6 metric tons. I really, really don’t like how dangerous this minimalist thing is, as it is now.
As for cost, this minimalist stuff is all throwaway. I don’t like that, either! Here on Earth, throwaway flight vehicles are called “missiles”, and they are quite expensive, much more so than "artillery shells". The other sort are called “airplanes”, and are much cheaper, but ONLY if you fly them many thousands of times!
Thrusters are ready "now". Retropropulsive landings are being done by two companies “now”, so far. Extendible heat shields are not ready to use yet. If they become so, I’ll use them.
GW
Last edited by GW Johnson (2017-02-25 15:20:17)
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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