Heatshield for Mars return

[keermalec]

<p>I'm working on a Mars mission design where a heatshield will be used twice: once for Mars aerobraking, the other on the return trip for Earth aerobraking.&nbsp; This seems possible using current technology but I am unclear as to whether a non-ablative space-shuttle type heatshield would be adequate on the Earth entry part.</p><p>In both cases, atmospheric entry is used to simply slow down the vehicle to a parking orbit, not bring it down to landing.</p><p>Here are the numbers:</p><p>Mars:<br />------<br />Atmospheric entry velocity:7.32 km/s<br />Atmospheric exit velocity: 4.76 km/s<br />Delta-v:2.56 km/s</p><p>Earth:<br />------<br />Atmospheric entry velocity:11.38 km/s<br />Atmospheric exit velocity: 10.73 km/s<br />Delta-v:0.65 km/s</p><p>As you can see, the entry velocities are high but the actual delta-vs are relatively low. From the literature I get that&nbsp;peak heating would occur&nbsp;somewhere between 9.2 and&nbsp;10.2 km/s on the Earth entry part.&nbsp;As we exit&nbsp;at 10.73 km/s we&nbsp;do not attain this&nbsp;peak heating. The Space Shuttle typically enters the atmosphere at 7.9 km/s and undergoes peak heating at something like 6.7 km/s. Are its heatshield tiles adequate for a Mars return of the type described above?</p><p>Thanks for your input.</p> <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[Boris_Badenov]

<p>While I don't know the math (I'm just a simple Chef <img src="http://sitelife.space.com/ver1.0/content/scripts/tinymce/plugins/emotions/images/smiley-smile.gif" border="0" alt="Smile" title="Smile" />) it sure seems to me like you're headed in the right direction. </p><p>Do the entry & exit velocities at the Earth end&nbsp;represent just the first part of the reentry maneuver?&nbsp;If so, I would think an ablative shield would be prohibitively large, & a Shuttle type radiative shield would be appropriate. </p><p>If the entire ship is reusable, & you enter the atmosphere in a taxi vehicle (Dragon, DreamChaser, Orion) then a multi-use ablative shield that can be replaced in orbit&nbsp;after X number of uses might be appropriate.</p> <div class="Discussion_UserSignature"> <font color="#993300"><span class="body"><font size="2" color="#3366ff"><div align="center">. </div><div align="center">Never roll in the mud with a pig. You'll both get dirty & the pig likes it.</div></font></span></font> </div>

 

[keermalec]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>If the entire ship is reusable, & you enter the atmosphere in a taxi vehicle (Dragon, DreamChaser, Orion) ...<br />Posted by boris1961</DIV><br /><br />That is what I was leading to, a single habitat for both interplanetary flights. Maybe the heatshield could be replaced as you suggest in a later evolution.</p><p>At the moment I question the feasibility of using the same heatshield for two&nbsp;aerobrakes with substantially different entry velocities and delta-vs. Is it possible?</p> <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[qso1]

The shuttle heat shield tiles are good for a maximum of 2,200 degrees F IIRC. The higher temps generated at the nose and wing leading edges being handled by RCC which handles up to 3,000 F. I listed these temps because that may help determine if the shuttle TPS (Particularly black tiles) is up to the task. That is, at the entry velocities you mentioned, what are the peak temps? <div class="Discussion_UserSignature"> <p><strong>My borrowed quote for the time being:</strong></p><p><em>There are three kinds of people in life. Those who make it happen, those who watch it happen...and those who do not know what happened.</em></p> </div>

 

[aphh]

<p>For aerobraking blunt-end-first is optimum shape? So your vehicle would look a bit like a reversed Apollo command module with the main propulsion residing at the coned end.</p><p>On the other hand, for the first Trans Mars Injection and aerobrake you could have a service module with propulsion unit, that is discarded before entering Mars orbit revealing the blunt end and the heatshield. For return you would have 2nd propulsion unit at the coned end.&nbsp;</p><p>Would it work?&nbsp;</p>

 

[scottb50]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>That is what I was leading to, a single habitat for both interplanetary flights. Maybe the heatshield could be replaced as you suggest in a later evolution.At the moment I question the feasibility of using the same heatshield for two&nbsp;aerobrakes with substantially different entry velocities and delta-vs. Is it possible? <br /> Posted by keermalec</DIV></p><p>This goes to the Cycler vehicle I have brought up. They go from LEO to LMO and back entering orbits at both planets. Actual landing and ascent vehicles, or containers, would be transported by the Cyclers.&nbsp;</p><p>To aerobrake into an elliptical orbit and then circularize it with engine burns would be well within the temperatures the Shuttle tiles are designed for, and they could be serviced at either end as needed, though it seems the actual descent of the Shuttle isn't any problem, it's the ascent. If you are simply braking into an orbit the Silica tiles would work very well.</p><p>What I have posted before is a core vehicle that various Modules attach to for cargo or passenger accomodations as needed. The actual departing vehicle could be a variable number of either. What I have also brought up is the heatshield is combined with Solar panels on one side and the shhield on the other. In transit the Solar panels are tethered to the vehicle, for orbital braking the vehicle docks to the panel side. This would provide Sun shielding both ways for the vehicle, which would reduce radiation hazards as well as produce power.</p><p>My plan also is based on taking a large amount of water and having double hull Modules with water between them to provide shielding for the crew. With a very large Solar array, and a lot of water, it would be easy to hydrolize water enroute for use braking into or out of orbit. Since a hydrolizer and a fuel cell are interchangable you could also provide power to the vehicle and draw off and cryo-cool the Hydrogen and Oxygen gas as needed for orbital entry or departure, allowing simpler storage of the liquid propellants. </p><p>If usable water is accessable on Mars you could refuel at both at both ends of the trip.&nbsp;</p> <div class="Discussion_UserSignature"> </div>

 

[keermalec]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>The shuttle heat shield tiles are good for a maximum of 2,200 degrees F IIRC. The higher temps generated at the nose and wing leading edges being handled by RCC which handles up to 3,000 F. I listed these temps because that may help determine if the shuttle TPS (Particularly black tiles) is up to the task. That is, at the entry velocities you mentioned, what are the peak temps? <br />Posted by qso1</DIV><br /><br />qso, I do not know the peak temperatures: that seems to be about the hardest thing to work out.</p><p>Here are all the specs in case anyone does know how to work&nbsp;it out (remmember, we&nbsp;only decelerate by 0.65 km/s):</p><p>Entry velocity: 11.38 km/s<br />Exit velocity: 10.73 km/s<br />Vehicle mass: 25 tons<br />Diameter: 5.4 m<br />Emissivity: same as Reinforced carbon-carbon (0.9?)<br />Drag coefficient: same as Apollo CM (1.1?)<br />Ballistic coefficient: 1100 kg/m2 <br />Lift/Drag ratio: same as Apollo CM (0.3&nbsp;for 25&deg; entry angle)</p> <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[qso1]

Peal temps for shuttle at 17,500 mph entry or thereabouts is 3,000 degrees F. For Apollo, slamming into the atmosphere at 24,500 mph...peak temps of 5,000 degrees F were sustained. I'm not sure how to calculate temperatures because in both cases, different heat dissipation methods are used. <div class="Discussion_UserSignature"> <p><strong>My borrowed quote for the time being:</strong></p><p><em>There are three kinds of people in life. Those who make it happen, those who watch it happen...and those who do not know what happened.</em></p> </div>

 

[DrRocket]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>Peal temps for shuttle at 17,500 mph entry or thereabouts is 3,000 degrees F. For Apollo, slamming into the atmosphere at 24,500 mph...peak temps of 5,000 degrees F were sustained. I'm not sure how to calculate temperatures because in both cases, different heat dissipation methods are used. <br />Posted by qso1</DIV></p><p>I doubt that you have all of the necessary data or computational capability to handle these sorts of questions.&nbsp; The factors involved include the speed of the body, the density of the atmosphere, the detailed geometry of the body, and the material of the heat shield.&nbsp; It is a fairly difficult analysis to determine the heat input from the flow of the atmospheric gasses (which requires a good model of atmospheric density as a&nbsp;function of altitude out to pretty high altitudes)&nbsp;over the body due to skin friction and stagnation of the supersonic gas flow, and then couple that with a heat transfer analysis involving conduction within the heat shield material and radiation from the surface.&nbsp; The 3000 degree figure for the shuttle is a result of all of these factors and could be different for a different vehicle with a different trajectory profile (as noted for the Apollo for instance).&nbsp; This&nbsp; problem is not particularly amenable to hand calculation.&nbsp; It is not an easy problem even with big fluid dynamics models and heat transfer codes, and high-end engineering work stations.&nbsp; Even then some sophisticated test capability is needed to verify material behavior.</p><p>Just these sorts of issues were involved in the development of some of the materials and processes identified to be used for in-orbit emergency repair of damage to shuttle tiles.&nbsp; </p> <div class="Discussion_UserSignature"> </div>

 

[qso1]

<p>Oh yeh, I'm sure I don't have that data. Another thing too is that in the early days of the space program, they actually tested various methods for re-entry...in part to determine what the temperatures would be for certain payload carrier shapes and warheads in the case of ICBMs. Which end would be the best end for re-entry factoring in cross range capability vs thermal loads.</p><p>What data I have seen seems to indicate the temperature is somewhat velocity dependant as well as dependant on the factors you mentioned. But because I don't have enough data, I don't know for sure. For the OP, looking at Viking data might be useful.</p><p>I should also mention for correction purposes to my previous post. The velocities mentioned are atmospheric entry velocities rather than the velocity at which maximum heating is occuring.&nbsp;</p> <div class="Discussion_UserSignature"> <p><strong>My borrowed quote for the time being:</strong></p><p><em>There are three kinds of people in life. Those who make it happen, those who watch it happen...and those who do not know what happened.</em></p> </div>

 

[DrRocket]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>Oh ...What data I have seen seems to indicate the <strong>temperature is somewhat velocity dependant</strong> as well as dependant on the factors you mentioned. But because I don't have enough data, I don't know for sure. For the OP, looking at Viking data might be useful.I should also mention for correction purposes to my previous post. The velocities mentioned are atmospheric entry velocities rather than the velocity at which maximum heating is occuring.&nbsp; <br />Posted by qso1</DIV></p><p>That is what I meant by speed of the body.&nbsp; In part what is going on is this.&nbsp; Think of the air as being in motion relative to the body of the craft (which it is).&nbsp; It is moving really fast, high mach numbers.&nbsp; So the molecules in the ari have a high kinetic energy due to that velocity.&nbsp; Now, when the air hits the body of the craft, it can slow down, and perhaps stop.&nbsp; If that process is done adiabatically, the air heats up to what is called the stagnation temperature, and it can be quite high.&nbsp; From this one can see that the dependence on velocity is very strong.</p><p>Thermodynamically this process is just the opposite of what is done in a turbine, or in a rocket.&nbsp; In those situations you start out with a hot gas.&nbsp; It is then expanded, allowing the gas to do work and in the process the gas achieves high velocity and cools off.&nbsp; In a stagnation process&nbsp; you do work on the gas which slows it down and compresses it, achieving high temperatures.&nbsp; The details of the process are dependent on the geometry of the body as well as the properties of the gas and involve some fairly sophisticated fluid dynamics.&nbsp; When you add in the complexities of heat transfer and the chemical process of ablation, it gets evern more difficult to handle analytically.&nbsp; The ablation factor is often fudged into the analysis by inserting a Cp&nbsp;(heat capacity at constant pressure) spike by hand into the material properties -- not very elegant nor traceable to first principles but expedient and accurate enough if one has a lot of empirical data.<br /></p> <div class="Discussion_UserSignature"> </div>

 

[qso1]

<p>Excellent posts on your part, I hope the OP reads your posts. He/she was designing a mission to Mars in which the heat shield is apparently used twice. I don't know if he/she is employed in the aerospace industry or just doing it as a mental excercise. I did a mars mission for a graphic novel I wrote but my solution was to have the mother ship and upper stage of the mars base return to earth orbit where the smaller upper stage (Slightly larger than an Apollo CM) would be transported back by means of a private enterprise aerospace plane.</p><p>The idea being to refurbish the Apollo CM like vehicle for reuse on a subsequent mars base. But since I was designing for a graphic novel, and for an as yet, uncertain future. I didn't need to know anything more than some general information on re-entry.&nbsp;</p> <div class="Discussion_UserSignature"> <p><strong>My borrowed quote for the time being:</strong></p><p><em>There are three kinds of people in life. Those who make it happen, those who watch it happen...and those who do not know what happened.</em></p> </div>

 

[keermalec]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>...it gets evern more difficult to handle analytically.</DIV></p><p>Difficult? Or impossible?</p><p>I'm not looking for an exact qualitative answer here, but just an order of magnitude. If the peak temperature is below 1600 K then a space-shuttle type re-usable heatshield can be considered. If the peak temperature is above 1600 K, then an ablative heatshield must be used (though I suppose this also could be re-used several times if the ablation layer thickness is designed for multiple ablations).</p><p>From what I can gather, the following relationships are true for a given vehicle, of given mass, shape, diameter, emissivity, L/D ratio and ballistic coefficient, evacuating heat by radiation:</p><p>Delta-v is proportional to air density * velocity squared, dv&nbsp;~ d.v^2</p><p>Heat transfer is proportional to air density * velocity cubed,&nbsp;q ~ d.v^3</p><p>Temperature is proportional to the 4th root of heat transfer,&nbsp;q ~ t^4</p><p>From relationship 1 I work out that the average air density over the vehicle's entry path&nbsp;need&nbsp;be only 5.7% of the average air density over the entry path of&nbsp;the same vehicle coming in to land (d = (0.65/11.38)/1), as I decelerate by only 0.65 km/s instead of 11.38.</p><p>Assuming that an Apollo-shaped capsule has a similiar entry profile to the space shuttle, <u>during the very first part of its entry path</u>, from relationship&nbsp;2 I work out that, in order to&nbsp;experience the same heat load, &nbsp;the&nbsp;<u>peak</u> air density over the vehicle's entry path&nbsp;must be 25% or less of the&nbsp;air density at which the space shuttle attains its peak heating (d = 1/(11.38/7.19)^3). NB the space shuttle enters the atmosphere at 7.19 km/s and attains peak heating at 6.7 km/s, 70 km up, which corresponds to an air density of 8.28x10E-5 kg/m3. Therefore&nbsp;our entry capsule's peak heating&nbsp;must occur at an air density of 2.07x10E-5 kg/dm3 or less,&nbsp;that is,&nbsp;about 79 km up or higher.</p><p>Now will the average air density above 79 km be sufficient to slow&nbsp;our vehicle down by 0.65 km/s? If yes, then peak heating will be below 1600 K and a re-useable space-shuttle-type heatshield can be used. If not, an ablative single-use heatshield must be used. From what I can work out, we are very close to a borderline situation which requires a very precise calculation. My impression is that with a lifting body we can and will attain the deceleration required without going below 79 km altitude: therefore a re-useable heatshield can be used.</p><p>See anything wrong with this rationale?</p><p>&nbsp;</p> <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>

 

[scottb50]

<p><BR/>Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>Difficult? Or impossible?I'm not looking for an exact qualitative answer here, but just an order of magnitude. If the peak temperature is below 1600 K then a space-shuttle type re-usable heatshield can be considered. If the peak temperature is above 1600 K, then an ablative heatshield must be used (though I suppose this also could be re-used several times if the ablation layer thickness is designed for multiple ablations).From what I can gather, the following relationships are true for a given vehicle, of given mass, shape, diameter, emissivity, L/D ratio and ballistic coefficient, evacuating heat by radiationelta-v is proportional to air density * velocity squared, dv&nbsp;~ d.v^2Heat transfer is proportional to air density * velocity cubed,&nbsp;q ~ d.v^3Temperature is proportional to the 4th root of heat transfer,&nbsp;q ~ t^4From relationship 1 I work out that the average air density over the vehicle's entry path&nbsp;need&nbsp;be only 5.7% of the average air density over the entry path of&nbsp;the same vehicle coming in to land (d = (0.65/11.38)/1), as I decelerate by only 0.65 km/s instead of 11.38.Assuming that an Apollo-shaped capsule has a similiar entry profile to the space shuttle, during the very first part of its entry path, from relationship&nbsp;2 I work out that, in order to&nbsp;experience the same heat load, &nbsp;the average air density over the vehicle's entry path&nbsp;must be 25% of the average air density over the early entry path of&nbsp;the space shuttle (d = 1/(11.38/7.19)^3). NB the space shuttle enters the atmosphere at 7.19 km/s.5.7% being less than 25% I&nbsp;conclude that the upper reaches of the atmosphere are sufficiently dense&nbsp;to slow my vehicle down by 0.65 km/s without letting it heat up to the space shuttle's peak temperature.See anything wrong with this rational?&nbsp; <br /> Posted by keermalec</DIV></p><p>I think it's more a matter of descending from LEO or plowing in at at speeds well above orbital. Then you have to deal with the reality of bringing something to Earth that might cause a lot of problems. Maybe keeping research on alien materials in LEO until safety is assured, makes sense to me.</p><p>We needed a vehicle, or vehicles, that can get variable payloads to LEO, over and over again. Next we need a vehicle that can go from LEO to other places in LEO. We then need a Common Plaform that can be put in higher orbit or used as moon or Mars vehicles. A Huge antenna to reduce user power requirements and focused transmitters to deliver faster responses.</p><p>Water is the oil of tomorrow, once you get it to LEO.&nbsp;</p> <div class="Discussion_UserSignature"> </div>

 

[keermalec]

Replying to:<BR/><DIV CLASS='Discussion_PostQuote'>I think it's more a matter of descending from LEO or plowing in at at speeds well above orbital. Then you have to deal with the reality of bringing something to Earth that might cause a lot of problems. Maybe keeping research on alien materials should be kept in LEO until safety is assured.We needed a vehicle, or vehicles, that can get variable payloads to LEO, over and over again. Next we need a vehicle that can go from LEO to other places. We then need a Common Plaform that can be put in higher orbit and assure World-Wide-Coverage. A Huge antenna to reduce user power requirements and focused transmitters to deliver faster responses.Water is the oil of tomorrow, once you get it to LEO.&nbsp; <br />Posted by scottb50</DIV><br /><br />Last post corrected. <div class="Discussion_UserSignature"> <p><em>“An error does not become a mistake until you refuse to correct it.” John F. Kennedy</em></p> </div>