PB666

I was going to say, people mistake the draw hardware on a Venetian blind for UFO's so its not surprising that something outside their house has more credibility to them.

I think the Italians tried it, they tried to inflate a parachute in space, the problem in space is that the particles are not cohesive, consequently there is no static pressure (uhmm I thin that is why its called space).

I know what caused this. Its not the Contrails per say. When the rocket reached a certain altitude it was in the sunlight, this is what caused the smoke tails to glow. You can clearly see the smoke even though its not glowing, just much less bright. The reason is that the sun was blocked from reaching it, it could be a cloud or some other obstruction on the horizon. That is to say the sun caused them to glow, the contrails are always there, too much moisture and too low of a pressure and temperature, when the sun could not reach them they could not glow.

What about a satellite made of 1" thick titanium designed to survive an explosion and reentry that documents UDE it could have little tiny force wires attached to the payload fairing documenting the forces of take-off, the event, temperature sensors that document the heat, three dimensional accelerometer, a small gyroscope that allows it to orient itself during a fall. Of course it would have to have a very durable battery and a transponder to tell everyone where it landed.

How about a satellite that's worth like $49.99.

Does it matter, both of them are going to be benthic substrate. That is off course if they did not close the kerosene tank valves, if they did they can just go pick up the first stage with a good crane.

They are going to land the first stage on the pacific ocean, that should be a big firey splash, lol.

Actually no, since initially there is no force acting on the rope, the ropes straitness will be governed by internal forces like tension that remains after it was last coiled, it would want to sort of coil back to that shape. Of course you could run the rope through a steamer to make sure all the tension is released. My idea is best, throw it backwards and let its w2r + g pull it down under the ship, as it entered the atmosphere the drag would pull it backwards. In this fairy-tail his atmosphere is not a grade it immediate, so chances are the 'knot board' at the end of the rope would bounce up and down each time tossing a little energy. The rope would have to be slacked out and then slowly pulled back otherwise the impact energy would cause it to pull the rope over the ship and slam back into the atmosphere.

If wishes were horses beggars would ride. Watching NASA over the last 50 years how many things NASA says it could do has it actually done.

1. The problem is that ISS is an international cooperation with multiple countries contributing. Thats what funny, suddenly the last 2 years many entities want a colony on the Moon. But I hear nothing about international cooperation. 2. The ISS is there at some point everyone will leave, but there is nothing to build, its already built. Getting back to DSG. IMO if its going to get done people really have to want it, and by and large if you were to Ask the 600 or so people who legislate from Washington maybe 10% may know what it is and of those 6% don't want to spend the money.

lol

Responses 1. Thermocouple efficiency is temperature dependent if you want to get away from mechanical systems then each phase needs to drop temperature. If you are going to use 973K the efficiency will be less, and it is highly unlikely that fusion would work at all. But if you had a 10% efficiency reactor you would need 2.5 times the reactor and many more kilometers of cooling. If you use a high temperature transfer temperature such as 973 k you will only benefit from 1 thermocouple and the power efficiency will be 10%, which means power needed it 2.5 times higher and waste heat. In my system the first therocouple is silicone germanium type with incoming T below 1380K and dropping out at 800K, the second thermocouple is lead alloy, and would drop the temperature to 450K the last thermocouple would be Bismuth/ lead telluride. Since no single TC is more than 10% efficient the 25 % comes from layering from hot to cold side of the radiator (The first one could be close to the reactor for liquid sodium transfer to the radiator). 2a. (do you see your 2 errors. lol) My temperature and calculations are based on 373K, as I properly stated and assume a three phase temp drop thermocouple system, you changed that. 0.7 GW of power = 0.5 x E x A x T^4. A = 1.4E9 / (5.8E-8 x 3734) = 1,243,000 sq. meters. . . .x 20 kg/meter = 24,000 t. This would decrease a from 0.0105 a to 0.000164, the dV would go from 9600 to 130. <--- Will not got to Mars, period. Conclusion is the same. 2b. However in your system you have a much lower power efficiency and since your efficiency is lower to higher radiator heat, you have 212,000,000 watts/ 0.10 = 2.12GW reactor output x 0.90 = 1.98GW. =---> 76,175 m2 x 20 kg/m2 = 1,523 t. This system would drop the a from 0.0105 to 0.002 and drop dV to 1828. <---- Will not go to Mars from LEO. Conclusion is pretty much the same. 2c. Streetwind said 1.7 terrawatts not gigawatts. He may be right, it might be impossible to have fusion in space for anything smaller, which means Fusion is summarily off the table for power plant. The system I created was proportioning weight with power need to avoid impossible scenarios like launching at 1 kT fusion reactor into space. Conclusion is pretty much the same. 2d. Agreed that current systems are too weighty, I was giving a best possible case scenario. All the more reason to reject Nuclear. Makes matters worse for proponents of Nuclear Electric propulsion. 3. A sphere is used as being the smallest possible structure that can support itself. Part of the weight issue for NASA's solar panels and radiators are the weight of the truss and the need to be extendable. A sphere could be support for much less the size, an accordian like truss would not work.

The power output of 20 of these reactors would not suffice to run a ship and 1 VASIMR. The system is designed to run on the ground not in space. Technically isn't a space reactor, its' a surface mounted colony support reactor. Abandon in place.

Some folks here have a hard time understanding that. Aside from that, even if you had to spiral out, it provides very little harm in terms of time and provides alot of benefit if the first halfof the orbits are Pe kicks. Originally I put up this example to show that if we had a Magic Wand fusion reactor, it would not grant us 39 days to mars, that the 39 days to Mars is a fantasy scenario. The point of the exercise is to show that player x is a space tug operator who carrys supplies to mars (200t to mars, 20t back), and these are the ways he has to get from LEO, where is picks up supplies from SpaceZ and carries them to Mars and back again. There is a huge connundrum for using ION Powered systems which you touched on but missed the Major point. In an ideal world you have a ship that burns in LEO all the energy it needs to get to Mars, but even through kicks during the last pass you need a system that can chug out 500 dV over a few minutes. KIcks can save most of the money by spiraling but not all. Given all this space tug operate has to choose amoungst the options for the most time economic and fuel economic way of get the cargo to Mars. Just to make the point that ventures that rely on ION drive systems (with no Magic wand power supply) are extremely wasteful and will not get the traveler back home. As I stipulated at the beginning, I don't believe Fusion power is an answer at all, it creates 2 problems for every problem it solves (see below). It does not violate the laws of physics per say, it disregards current human abilities to convert heat to energy. 1.7 TW is not a problem if the fusion reactor is 100% efficient in the power conversion, if it is only 99% efficient it generates 10 MW of power, as I have stated here recently that a fusion power reactor would need at least 100 times its mass in radiative cooling (see below). The engines themselves are rated at 24 kw per 0.3731 meters so we can generate 64.32 kw of power to thrust conversion per meter. So for a system that weighs 380 t at 0.0105 acceleration requires 3933 N of thrust These are the parameters, they are not as unrealistic as you think, save one (which itself is composed of many). 3933 x 81128.7 / 2 x 0.75 eff = 212,719,415 watts of power. This can be delivered by 3307.2 meters squared of ION drives (93 tons might be a credible weight). HiPEP with a little more investment could get the efficiency higher, its power density could increase, it did perform at 80% efficiency at ISP 9650 and 39 KW, it could be possible to get it to 85% efficiency at 8000 ISP at say a power density of 100 kw/meter. It is possible to get 93 tons of ION drive laid out on six spans that cover the desired area once they flip out. At the center could have a carriage that would insert a 50 ton reactor, all of that is credible, hard but credible. What is incredible is the fusion reactor. Scenario 1. we can get a fusion reactor that produces 212,719,415 watts of electric power at its destination. The best power conversion in space at the moment (idealistically) would be a three phase thermocouple that delivers 25% of the input heat as electric power. So thats ~ 1 Gigawatts of power. How do you start a fusion reactor, you need batteries that can generate 1 Gigawatts of power. Most of the power if not all of the power would go back into initiating the reactor. Suffice it to say we don't know how bad the waste problem will be with a fusion reactor because we do not currently have a working fusion reactor, but even in this situation a sizable fraction of the engines total mass will be tirgger batteries required to initiate the reaction. So if we started with a 50t reactor we can add another say 20 tons for batteries. Next we have to deal with waste heat. If all the above is true then there is going to be 0.7 GW of waste heat. If you set the maximum temperature of the radiator at 100'C then you can see (Stefan-Boltzman constant = 5.67067 x10−8) Where T = temperature epsilon-sigma-temperature^4 is Q the heat flux the specific Power. If we set emmissivity at 0.5 and temper at 373K we can calculate the area required to dissipate a give amount of heat to be about 50watts per meter. So if you don't see the problem now, its very simple to make it visible. .7 GW/ .00000005 GW = 12,755,00 square meters (12 sq.km) this could be possible down by a factor of 3 to about 4 km this produces a sphere with a radius of 1 km. Since in our sphere will be facing half its surface to the sun, that side cannot be used to transfer heat, then we would need 8 km of shell or a radius of 1.414 km in Assuming the cost is 1 kg per meter or radiative heating (very optimistic) then such a Fusion reactor would require approximate 8 million kg of radiative cooling. So now our fusion reactors power system dropped from 0.01 a to 0.00045 a. This is the best case scenario. Scenario 2. A probable scenario is that Fusion will work but only at the Terawatt range as you say, in which spectacular amounts of heat will be generated again we are talking about heat shells 10 to 100s of kilometers in radius. Scenario 3. Fusion electric power is never a thing. Replace Fusion reactor with a lower efficiency per mass Fission reactor and the cooling problem is the same as scenario 1. So if we disregard all nuclear electric power systems over 10 kw and look at solar. . . . . . . . . Alt. Scenario 1. Solar power at 212719415 / 0.4 KW = 531,798.537 kg Removing 50t for the reactor and adding 531t gives us acceleration of 0.0046 but 500,000 square meters is 0.5 square kilometers of panels. And so now we move to alt 2 Alt. Scenario 2. We split scenario 1 into say 20 ships of managable solar panel size and we link them in a long teather to carry the payload behind. (ION drives will need to point off angle).

I was being lazy, I did not want to have to give the math involved. VASIMR has not changed all that much in 6 years, its still heavy, it still cannot outperform lighter technology that is older than it is. My personal reason, even looking at the most recent specs, VASIMR still underperforms HiPEP with regard to Thrust to weight ratio, with regard to efficiency. HiPEP is scalable from as low as you want (for example using capacitors). IN FACT if you wanted the most stable system have twice as many HiPEP as you can power and alternate their use so that they run cooler and run them at the highest possible ISP, that achieves the best efficiency, per weight. Reason #1 Its not about how much KW you can use, its about how much acceleration you can produce per amount of fuel and per kilowatt. The 39 days to mars justifies VASIMR for the sole reason that if you have the power to burn from LEO to Escape velocity (3212 m/s) you only need 400 more dV to reach Mars, but in 100 to 150 days. So VASIMR is something you add to the craft to make it get there in 39 days once you have burned to or almost to escape velocity. The problem is that you could almost do the same with conventional power given the extra added weight of VASIMR and the power supply. But the craft is already going to have solar panels, so just plop a few lightweight HiPEP on the backside and maybe a few more panels and use waste electricity to shorten the trip time. We are talking about the difference between 20kg of thruster versus a couple of tons. Its a no brainer, go with HiPEP. Reason #2 I could not care less if humans went to Mars, if you want to send some joker there to die on Mars, thats fine 100 day trip aint going to kill him half as fast as living on Mars will. VASIMR will not get them back home. My only thing is get a sample return mission going, and that sample return mission is probably going to come on a conventional ION drive system that can shuttle materials back and forth from Earth. Reason #3 is that it will not pay off, as already stated you have a technology that is a fraction of the weight, you can have multiple redundancies and in the end if one breaks you flip on the next one for the cost of one VASIMR. There is no scenario were VASIMR can pay off because you have this already Thrust = 2 * 0.8 * KW / 95000 or Thrust = 2 * 0.72 * KW /54000 This equation is tyrannical and VASIMR does not change or alter the equation, it never can, that's that. The best VASIMR can do is get efficiency in the 80 to 90 range. The most you can ever get is 100% efficiency. Power comes at the expense of something, and that something we do not have. We do not have a nuclear reactor and as far as I know, NASA is not planning one. Second there is no situation were radiative cooling and the reactor is going to produce 200kw of power unless the nuclear reactor exceeds 1 Megawatt, in which case the ship needs to be able to cool 800 MW of radiative heat, this involves mass also. For HiPEP we can have a system, light weight, essentially flat, small enough to run on solar power small enough and light enough to cool itself. Another example of this is the Cannae drive, everyone talked about how it was going to change the world being 10 time more efficient than a photon drive. TPhoton drive = 2 * Power /3E8 = 0.000007 N/kw versus Tcannae = 2 * Power/3E7 = 0.00007N/KW <----- not changing fundementally the thrust issue with electric propulsion systems. Reason #4, Build to the power, and not power that needs to add mass just to cool itself. If it is not self cooling, leave it near a cooling tower on earth. The future is better solar panels, they can be lightened and they can produce more power, but they can be cooled simply by changing their angle to the sun. VASIMR is doing the exact opposite, it is building to the power we do not have in space. IF we go by current specs at 34 kw per 14tons to get a ship with 200 kw would require 100t just for solar panels, not intelligent. The only reason ever to use nuclear power in space is to travel beyond the asteroid belt, for missions to Jupiter this may be neccesary. That justifies nuclear power, but it does not justify VASIMR. SLS can get beat when Space X proves itself with the heavy project, but they have to prove themselves.

Are we talking SpaceX rockets that have been launched in the last year or the last 5 years. In the last year launches have been extremely reliable. In fact their overall debris status is better than the space shuttle over the last year. Most of the first stages have been returned to a secured location. For the space shuttle, not counting in flight disasters, but two booster cores coming back to earth. In the last year has space X dumped a single space craft part into the water under the launch corridor.? Second thing is, Boca Chica is relatively stark terrain there is not alot of ship traffic crossing the border. Most of the ship traffic I have seen are boats weighting north of the N. Port Isabel Jetty. There is a Sea Way there but its about 5 miles north of the launch site Finally, when they start launching, i'm gonna take my 4x4 down there and camp on the beach. The more the merrier.

You don't understand really how badly VASIMR performs compared to other lighter electric propulsions. Take a look at this page. " The pre-prototype HiPEP produced 670 mN of thrust at a power level of 39.3 kW using 7.0 mg/s of fuel giving a specific impulse of 9620 s.[2][4] Downrated to 24.4 kW, the HiPEP used 5.6 mg/s of fuel giving a specific impulse of 8270 s and 460 mN of thrust. " and this page. https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20040139476.pdf After reading those compare that stats with VASIMR. http://spacenews.com/vasimr-hoax/ Thrust = 2 * power * efficiency / Vexhaust Accleration = Thrust/Mass

6423 - 3211 = 3212 you think wasting half the fuel is acceptable!!!!!! ROFL. Don't apply for the captain-ship of a space tug. The 3500 you quote is to escape Earth and intercept Mars, to just escape earth by bruning at at 160 km orbit only requires 3211.888. This is considering the effects of the sun and the hillsphere. See other thread, a spiraling orbit is actually slower to break orbit than one that takes a rest for 90' per orbit (270 burn span versus 360' burn span). Two things that you believe to be true are unncessary. 1st, that with an abundance of power (Such as in the fictitous fusion electric system . . .again weight is not established, you may well need to use the highest ISP) that you should not use high ISP. Actually, if you want to take advantage of low ISP, the only portion of the system leaving where it would be advantageous is on the last passes, and in particular close to the periapsis, the lower the better. This is because if you have established a low ISP from previous kicks, you can now take advantage of an 10,000 m/s, If you cut your ISP by a factor of 3 but your moving 4 times faster, the energy you gain is still 25% higher than if you had spiralled out an using the highest ISP. Using a lower than optimal ISP before the last pass is simply wasteful and wont save much time. During the last pass it will save both energy and time. 2nd. For most of the burn out. . . 3/4ths of it at least your can save 900 dV with no loss of time by using highest ISP available. The reason for this is that burns at apogee are very wasteful. by burning at 3/4ths of the time you save a periapsis burn meaning that you increase the amount of energy put into the orbit versus at the wasteful apoapsis. Then paper you quote apparently did not try to optomize for burn span, must not be a very good paper, since in a few hour of work I did.

This is a statement whose foundational logic is in error. The ISS is at an inclination of 51.6 degrees and Cape Canaveral is at 28.6N Boca chica is at ~26'N. During every orbit the ISS crosses 28.6'N and 25'N once as it travels north and once as it travels south. At a semi-major axis of 404km in altitude (a = 6776000 m, v = 7669.76 m/s, P = 5550 sec, f = 31/day) the ISS, on average, crosses either launch site and any launch site between 56.0'N and 56.0'S passes within 5.8 degrees along the east west (or north as one approaches 51.6 degrees north). The reality however is the 51.6 - 28.6 = 23' or 51.6 - 26' = 25.6 means that the distance in absolute degrees is ~ 2.26' at CC and ~2.59' at BC. Given the earth is 40,000 km in diameter this is a distance from ISS course (DfIC) of 251 km for CC and 288 km for BC. However if one looked at windows every 2nd day or 4th day the DfIC would be half or a quarter. But lets just use the daily value. Lets say that 251 km puts the two vessels at 0.351' and 0.412' orbital inclination to each other. To correct this in orbit (the worst case scenarios) one would need a burn dV of 46.9 for CC 55.72 m/s for BC, a difference of only 8.73 dV higher for Boca Chica. However, the western launch from Boca chica has a higher contribution from Earth's rotation. 40,000,000/86400 = 462.6 meters per second on any east west departure at the equator. For CC its 406 m/s and for BC its 415 m/s, since both intercept trajectories will be traveling ENE or ESE to intercept there are cosine losses 373.3 for CC and 374.3 (and yes I am aware of the east launch restriction, but also don't forget due east is only restricted to 320 km downwind). Consequently we can calculate the dV required to reach orbit for any average day if we stipulate that drag and gravity losses equate to 1200 dV. Boca chica launch site is about 8 miles from the Mexican border so the maximum variation is 2.92' South which means at 200 miles it would only need to turn 23' South. SQRT(2 * (7669.7^2/2 + u/6371000 - u/6776000) ) + 1500 = 8142.3 + 1500 = 9642.3 dv. For CC it would be that + 46.9 - 373.4 = 9315.4 m/s for BC it would be 9323.72 a difference of only 8.32 dV. Remembering that we have to make an orbital inclination burn anyway of 0.412' orbital inclination burn, but we also use more momentum from the Earth. One could after traveling 183 miles eastbound heading south at 25.6 degrees, the horizontal orbital velocity at that point would be 3000 m/s so that the turn would cost 1296 m/s however the cost could be lowered if and agreement could be reached with Mexico allowing it to change course over the commercial zone. However in choosing an east to west launch one gets back 30 dV of the cost. The other alternative is to simply catch the ISS on the rise, since the first stage would be captured in the gulf of Mexico, the only down range risk is the failure of the second stage (which it will either fail at firing or quickly be out of risk for the continental US.). Space Law applies to any vessel above 100 km; however, and this has major importance on such a course change. Because the ISS is so high now, 404 km, one can choose a vertical course steeper, and essentially fly over Mexico's internationally recognized airspace of 100km with a surface velocity of 600 m/s (and 415 = 1015) Since. this will occur pretty close offshore, you could launch to the east get about 300 or so horizontal velocity and turn to 120' the cost of the 30' change would be a loss of 200 dV and (increased because of the loss of some of earths rotation that contributed with east to west take-off). In the grand scheme of things that would only be an increase in Boca-chica versus Canaveral of 2.2%. Ok that is cleared up. The same is also true for GSO orbits, since they are so high, there is no general benefit to try to cling to the Earth, However, Boca Chica is better launch point for Equitorial GSO than any launch site in the continental US. However because Space X has a barge technology they can simply barge a rocket from the Port of Brownsville, haul it to a launch site and launch it, this could be done from any port that has access to the rocket assembly technology. Finally, the Sun, the sun (tropic of cancer) approaches Boca-chica from the south, Once June 21st this orbit 3 33'86'' from the Boca Chica site at Midday, it can be also reached at Midnight on December the 21st. Both can be accessed from a due east flight and this allows for launches to other targets outside the solar system. During spring it is as great as 25.8 degrees to the south (or north if launched at midnight). Boca chica is a better launch site for orbits that travel out of the planetary system, it captures more of the Earths momentum (9 dV more and needs to make a smaller dV change to achieve a decent orbit for escape from our system, however because of the dynamics of the Earth moon system this needs to be precisely calculated. Boca chica is not suitable for polar orbits, those are best launched from Siberia or Alaska. Biggest problem with Boca Chica. The launch site they are building on is the flood plain of the Rio Grande river, its not particularly stable soil (I know this from very unpleasant personal experience). They will have to spend alot of time stabilizing the soil because of the large amount of sulfides and other unstable organic materials that have made the sand spongey. It is hard to imagine the amount of sediment at the site, but just consider the length of the rio grande and the terrain it has cut through.

not bury, it does not seem to me that anyone is seriously planning to build or launch it. If you want to launch something you don't announce that you are going to have a half-fit group to study it. If DSG is supposed to be a cryogenics depot as stated, where is the launch vehicle that is going to bring fuel to refill it. You see any bulk liquid hydrogen or oxygen carriers out there. Certainly we are not go to have a cryogenic refueling station that is loaded with 4000 kg of fuel at a time? And frankly I don't think they have the cryogenic storage worked out either.

It assumes that his business model is expansive, that is to say that this years number of flights are greater than last years, the number of customers have increased and the costs have fallen.

You could just ask scotty to beam it over.

Let me start with the following disclamer. The results I present are for an ideal ship weighing approximately 335 t with 39t of fuel (Xenon or Argon). The ship was designed to shuttle 200t to mars and bring 20t back. The fuel that is provided is ideal fuel and assumes that burns are instantaneous. In reality more fuel would need to be brought along. The drive is assumed to have a nuclear powersupply, no such power supply exists AND if it did the cooling systems for the power supply would be much greater in mass than the mass provided for engines, power-supply, and radiative cooling. The vessel produces in excess of 0.0105 acceleration and has a total dV in the 8.7 km/sec range. The first set up assumes that the burn angle occurs around the periapsis of an opposing exit vector. So if a burn span is 270' then the burn begins 45' after the exit vector and continues until 45' bfore the exit vector, resting as it travels through the apoapsis. The first set up gets the ship to Earths average hill sphere with zero velocity relative to the Earth. SpanBurn TBurn (sec)(cumm.) T (to escape) Dv (m/s) Waste(m/s) 360 ' = 2π 604413 >7 days 6633 3422 270 = 3π/2 524607 >7 days 5726 2514 240 = 4π/3 452699 >13 days 4917 1705 180 = π 364657 >60 days 3937 725 120 = 2π/3 324279 >200 days 3491 280 90 = π/2 313283 >400 days 3370 159 60 = π/3 306224 ~1000 days 3293 81 Burning to Mars. Adding the 6471769 J/kg required to reach Mars at its apogee. SpanBurn TBurn (sec)(cumm.) T (to escape) Dv (m/s) Waste(m/s) 360 800005 >9 days 8902 5018 270 719818 >8 days 7964 4080 240 634915 >13 days 6863 3099 180 527567 >60 days 5759 1875 120 491702 >200 days 5355 1471 90 483796 >400 days 5265 1381 60 476635 ~1000 days 5185 1301 Note that some waste in the final burn to mars is unavoidable, because once spefici orbital energy >=0 one cannot 'kick' past the periapsis and so once the final burn that has SME_>0 has also be a parto of the burn that goes to mars despite the fact that most of the burn occurs at low velocity (and thus low dE/dt during the burn). There are several ways to improve this. 1. Burn down 5 to 10' more than prograde to keep Pe low (from 270 to to 200) 2. On the last jump, jump to 270 and also burn down at 30' 3. Switch the ION drive to burn at a lower ISP and higher thrust. 4. Have an auxillary chemical based engine on board. The dV difference between escape, and mars intercept is only about 400 at burn from 6531 km and so a reasonably small fuel tank with powerful engine would suffice, the fuel tank can be ditched after the burn, it will go into a circumsolar orbit. Lessons to be learned: In a burn to Mars from Earth orbit, you do not want to use a spiral (360 burn span) orbit. It wastes both fuel and time. Since a Mars mission would be at least 90 days the loss of 5 days getting a better fuel economy is worth it. This means that blocking off 120' of the orbit as a no burn zone is a reasonable move. For a space tug that has no real cost of time except for capital cost of inventory. blocking the burn from a half to 2/3rds of the orbit is a reasonably good choice.

Gravity is a faux force, just as centripetal acceleration is a faux force, in a circular orbit the two faux forces cancel each other. Lets sat you wanted something to fall say 30 feet in space, like throwing an anchor. You throw the anchor behind the craft, not under it, provide it with enough rope and it will fall because now gravity is stronger than the centripetal force. As it falls it will speed up and eventually be under the craft. If you though the anchor down it will end up in front of the craft, if you through it in front of the craft it will end up over. Here is the equation you want. V2= u (2/r - 1/a). If you throw something down you don't change a, you change e. If you throw it behind, you change a and since you would be at forced change at apoapsis then periapsis must change. By the same token your ship will gain a, which means it would be harder for it to enter the atmosphere, as the object is falling your ship is speeding up, so the rope should be very long. The foolproof way of causing the ship to enter the atmosphere is the way I describe, toss some weight faster than the ship and the ship will slow down and fall into the atmosphere. The reason that very high altitudes are hot is that they are being bombarded with energetic particles from space. You really don't care about the heat of space because heat density is so low. What you care about is the heat the velocity your craft creates. However there are ways to control this heat. 1. Reentry angle is oblique, this means craft slows down at higher atmosphere. 2. The shielded areas of the spacecraft can be tilted to provide lift. Just as in number 1 reducing that -verticle velocity can reduce the rate at which the space craft encounters denser gas at higher speed. 3. Thicker shields or wider shields. 4. retrograde thrust.

I have to agree somewhat with the space argument for the sheer reason that if there is enough space, say 10s of meters the collision residue would have ample incidence angle x distance to avoid the second collision with the hull. The problem with objects traveling at near light speed is that you have other energies to also consider (for example fusion of nuclei) and the resulting atomized particles would best be to spread out. However structurally, a thin plate being hit with an artillery shell energy rocket, you are going to have unexpected oscillations in the plate caused by the collision, so that the structure of the plate needs to be reinforced. So it might no be a bad idea to have two thin plates with an aerogel in between.