The 80/20 rule

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barrykirk

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I've been following the debate between a lot of people about RLV vehicles versus Expendable vehicles.<br /><br />It should be all about the 80/20 rule. If you can divide your cost into two components, one responsible for 80% or your costs and the other responsible for 20% of your costs. You should make sure that the 80% component is minimized before tackling the 20% component.<br /><br />I'll use the spacex falcon 5 rocket as an example.<br /><br />The first stage of the falcon 5 rocket is responsible for about 80% of the total cost of the rocket. 5 Merlin Engines for the first stage versus 1 Merlin engine for the upper stage. I'm assuming that the other components will scale about the same.<br /><br />Note that the first stage is completely reusable. It comes down on parachutes. There is no fancy TPS required because it never reaches a really high velocity and altitude.<br /><br />The second stage is not reusable, but it's not a big portion of the cost, only 20%.<br /><br />Let's look at NASA's new Heavy Launch Vehicle, the one with two SRB and 5 SSME. The SRB are cheap and re-usable, the SSME are expensive and not reusable.<br /><br />So here we are re-using the 20% cost and not re-using the 80% cost.<br /><br />Sorry, but I think they got it backwords.
 
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barrykirk

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The whole point of a first stage is it is very easy to make it re-usable, as long as it doesn't go too fast or too high. I don't think strapping on SRB's to a falcon would improve things much.<br /><br />All they would do would be to increase the initial acceleration, which would increase the velocity at the point of Max Q. The framework of the rocket, including the upper stage would need to be heavier to handle the added stress.<br /><br />Also, the lower stage might be too high/fast after burnout and require a TPS to make it re-usable.<br /><br />How about modifying the Heavy lift vehicle as follows.<br /><br />Reduce the number of SSME from 5 to 1 or 2 SSME. This will reduce the largest portion of the expendable cost. The LOX/LH2 tank can be reduced in size.<br /><br />Increase the number of SRB from 2 to 4. The new SRB will be 6 or 7 segment. So you have 2 SRB with 5 segments and 2 SRB with 6 or 7 segments.<br /><br />What you've gained is, the total cost of the throw away portion is reduced.<br /><br />Now to be honest, I've done no calculations here. I have no idea how this will affect mass to orbit or total cost. Or even if this will work at all.<br /><br />It is based on the assumption that <br /><br />1) SRB are cheap.<br />2) SSME are expensive.<br />3) SRB are very heavy, but have high thrust and burn out quick.<br /><br />The general idea is total mass at liftoff will be slightly higher ( more solids less liquids ).<br /><br />The two shorter SRBs are reusable. The extra segment SRBs are ejected too high/fast to be re-usable, but by the time they are ejected, the upper stage is high/fast enough, that 1 or 2 SSME can finish the job of reaching orbit.<br /><br />Does anybody have a spreadsheet that can be used to make these calcs to find a more optimal solution? I would be happy to run the numbers, but I need a starting point.
 
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vogon13

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SRBs are very heavy.<br /><br />You aren't kidding. They make them out of iron. An unusual choice in the aerospace engineering world.<br /><br />A different material was discussed (fibreglass sort of) and maybe even one was built and tested but NASA never persued the idea, as far as I know, not even after the Challenger mishap.<br /><br /> <div class="Discussion_UserSignature"> <p><font color="#ff0000"><strong>TPTB went to Dallas and all I got was Plucked !!</strong></font></p><p><font color="#339966"><strong>So many people, so few recipes !!</strong></font></p><p><font color="#0000ff"><strong>Let's clean up this stinkhole !!</strong></font> </p> </div>
 
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barrykirk

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Actually, I believe that the SRB's use Ammonium Perchlorate as an oxidizer and aluminum as a fuel.<br /><br />The steel is in the casing, so it is only a small portion of the weight.<br /><br />It is heavy, but it produces a great deal of thrust... for a short period of time. The ISP is low, but not incredibly so.
 
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vogon13

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Sorry I wasn't clear, of course the casing is the iron part. Did not mean to imply fuel was iron too. (if I did).<br /><br />Every pound you shave off the casing is energy that can be imparted into the vehicle instead of the splat when the SRB casing hits the Atlantic.<br /><br />Iron really is an unusual material for the aerospace world though. Only other big thing I can think of is the XB-70. It's wings were made of a brazed stainless steel honey-comb material. Needed for the stress and heat of Mach 3 flight in the atmosphere.<br /><br />Soviets might have used steel in some fighters too, seems like a pilot of their's defected with his jet and the CIA exam of it was rather surprising in just how heavy and (for lack of a better word) crude it was.<br /><br /> <div class="Discussion_UserSignature"> <p><font color="#ff0000"><strong>TPTB went to Dallas and all I got was Plucked !!</strong></font></p><p><font color="#339966"><strong>So many people, so few recipes !!</strong></font></p><p><font color="#0000ff"><strong>Let's clean up this stinkhole !!</strong></font> </p> </div>
 
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drwayne

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"Actually, I believe that the SRB's use Ammonium Perchlorate as an oxidizer and aluminum as a fuel."<br /><br />The fuel is all in a binder of HTPB, which is also a fuel element itself. There are other odds and ends, such as some iron oxide which serve to provide extra kick.<br /><br />Wayne <div class="Discussion_UserSignature"> <p>"1) Give no quarter; 2) Take no prisoners; 3) Sink everything."  Admiral Jackie Fisher</p> </div>
 
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drwayne

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I believe you are talking about the MiG-25, which was the boogey man for a while there, and provided the impetus for the USAF to develop the F-15.<br /><br />Wayne <div class="Discussion_UserSignature"> <p>"1) Give no quarter; 2) Take no prisoners; 3) Sink everything."  Admiral Jackie Fisher</p> </div>
 
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barrykirk

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Well a steel casing in the first stage isn't as big a deal as it sounds. Sure it could be made lighter by using composites or aluminum alloys, but as one poster said a while back and i'm paraphrasing.<br /><br />Where the pounds are is more important than how many pounds there are.<br /><br />If you increase the weight of the first stage by one pound, you decrease the payload by a small fraction of a pound.<br /><br />If you waste one pound of weight on the payload, you decrease the usable payload by one pound.<br /><br />If they were using steel in the final stage to orbit, that would be significant.<br /><br />Also, isn't steel better at handling heat one re-entry than composites or aluminum alloys?<br /><br />That might be the difference between needing a TPS and not needing one.
 
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propforce

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In this case, I'd say 80% of critics and 20% of workers on <i>any</i> new launch vehicle designs. <img src="/images/icons/smile.gif" /><br /><br />I think we're locked-in on the SSME for a couple of reasons:<br /><br />1) We don't have any other similar engines made by the U.S., thanks to a lack of any real engine development effort in this country for the last 30 yrs. The RS-68 was a privately funded effort. <br /><br />2) We need the SSME type of performance in order for the CLV to work, thanks to the poor choice of SRB as the first stage booster. It's low Isp means the ET/SSME has to carry more than its share of total velocity (delta-vee).<br /><br />3) Unfortunately, an alternative liquid booster to the SRB goes back to #1, no existing engines available. Brining back the F-1 would essentially means embarking a new engine development program which would drawn out the timeline beyond 2010.<br /><br />So if you look at the 'whole pie' as oppose to a slice of pie, optimizing the mission objective as oppose to a subsystem, surely you'd realize the constraints imposed by the decision makers. <br /> <div class="Discussion_UserSignature"> </div>
 
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mikejz

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Mind you that the currently no SSMEs are being produced, and i'm not sure if all the equipment to build more still is in operation.
 
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wdobner

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I know this has undoubtedly been asked and I apologize for whatever repetition it causes (if you could just point me toward the argument where this was worked out, I'd be appreciative). <br /><br />Why can't we use a few more Delta IVH missions rather than try to develop a brand new heavy lift launcher? I thought the whole idea behind EELV was the eliminate this kind of thing, why suddenly does NASA need it's own launcher? Even if the EELV derived booster cannot match the lifting capability of the proposed heavy lift shuttle derived launcher with it's SSMEs and such, couldn't the crew at least ride into orbit on an EELV derived booster? For years it seems we've been told that liquid boosters are generally safer than solids, probably a byproduct of the Challenger disaster and the blame placed on the SRBs. Why now when we're going to build a new booster do we go back to solids to handle not just a strap-on role, but the entire first stage of the booster? Correct me if I'm wrong, but I believe the Delta IVH uses all liquid boosters in the first stage, so that the booster could be shut down if something went awry within the last few seconds before launch. Admittedly the Delta IVH is an entirely throw-away design, but wouldn't the savings versus developing a brand new booster which only reuses 20% of it's mass make up for that slightly greater expense per mission. Put the money saved toward more useful things, like making expendible rockets obsolete altogether. <img src="/images/icons/smile.gif" />
 
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barrykirk

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Well, that is true. If they had to engine choices to pick from, SSME and SRB than they had a problem. I was talking mostly about the Cargo Heavy launch vehicle since that was the more expensive vehicle. The CLV has one SRB and one SSME so it's cost is what about 20% of the cost of the cargo vehicle. There is the 20% number again.<br /><br />Actually, using an SRB as a first stage isn't such a bad idea. The further you have to lift/accelerate the fuel before it's used, the more expensive it becomes. You want to use your high ISP expensive fuel at the upper stages where it counts. At the first stage, you haven't invested much energy into the fuel so it's cheap. Note that I'm saying cheap/expensive in terms of energy not cost.<br /><br />Next factor. Gravity and air drag loses are highest at the beginning of the launch. So you want to minimize the time spent at that stage. This means that high thrust is critical at the beginning. SRB's have got high thrust in spades.<br /><br />The SSME is a realatively low thrust engine. Sure it's got high ISP, but it's thrust is puny compared to an SRB. This isn't a problem when your close to orbit and you have almost no gravity loses. At that point ISP is the most important factor.<br /><br />Also, you don't want the first stage going too high or fast before burnout. Otherwise, it gets expensive too reuse it. Upper stages will always be more expensive to re-use.
 
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barrykirk

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Actually, Werner Von Braun said something along the lines of and I'm paraphrasing here, men shouldn't be launched by solids.<br /><br />Liquid fueled rockets do blow up. So do solids. Which is safer? I would imagine it all depends on the engineering and the production and implementation.<br /><br />A good solid can be safe enough for people. The military has been using JATO solid rockets for aircraft for years. Aircraft ejector seats also use solids.<br /><br />As a general rule the lower the ISP the higher the thrust.<br /><br />SRB's have an ISP of about 237 seconds at sea level. But they have a huge thrust to weight ratio<br /><br />LOX/RP-1 engines have an ISP of about 255 at sea level. But a medium thrust to weight ratio<br /><br />SSME engines have an ISP of about 363 at sea level, but they have a low thrust to weight ratio.<br /><br />Using an SRB for the lowest stage makes perfect sense.<br /><br />High thrust to weight ratio is necessary at the launch where gravity and aero losses are high. Also, at sea level rocket engines aren't that efficient even SSME.<br /><br />For a second stage, solids only make sense in a very limited way as I suggested in another post. You wouldn't want to take them all the way to orbit.<br /><br />For an upper stage or for a stage used in orbit only, SSME does make sense.
 
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propforce

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<i>".... Why can't we use a few more Delta IVH missions rather than try to develop a brand new heavy lift launcher?..."</i><br /><br />You mean other than political reasons? <img src="/images/icons/wink.gif" /><br /><br />1) The existing EELV payload capacity is limited to 50K lbm to LEO, lower to the ISS. The only existing EELV capable of lifting that 50K payload is the Delta IV-Heavy (DIV-H) and is marginal at best, while NASA wants something as big as 30MT (66K lbm) as crew transport to ISS. Why? I have no idea <img src="/images/icons/rolleyes.gif" /><br /><br />2) The current EELV fleet was designed for <i>unmanned</i> flights. This means that they cut back on lots of stuff, including making tank wall as thin as possible to save weight. To embark a 'man-rating' exercise would effectively design 2 brand new vehicles, one for Delta and one for Atlas. Some question the cost effectiveness of this exercise.<br /><br />3) Both EELVs do not address the 'heavy lift' capability needs of NASA's cargo carrier ~ 300K lbm payload. Even at that payload, it is still less than what the old Saturn V could carry, and yet NASA has a far more ambitious objective than just another 'photo opportunity' on the moon surface this time. Some would raise serious doubt about needing a much bigger heavy-lift than the current concept of SDHLV .<br /><br /> <div class="Discussion_UserSignature"> </div>
 
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josh_simonson

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I'd like to point out that the more segments you put on a SRB, the higher the pressure is generated inside, requiring the casing to be re-designed in a heavier fashion. The goal of the program is to do as little re-design as possible, so more than the already tested 5 segments is out.
 
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propforce

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<i>"...As a general rule the lower the ISP the higher the thrust. ..."</i><br /><br />Oh no way !! <img src="/images/icons/shocked.gif" /><br /><br />That is a common misconception. <br /><br />As a general rule, engine efficiency, e.g., Isp, has more to do with <i>propellant of choice</i> and <i>engine cycle</i>, as oppose to thrust level. Where as any propellant system can be sized to any thrust you want.<br /><br />Therefore an engine's thrust level is independent of its efficiency, Isp.<br /><br />Take the SSME and RS-68 for example, they both use LH2/LO2 as propellant, so they are inherently more efficient in their combustion (C*) efficiency than, say, kerosene, hypergolic, or solid propellant, engines. But because the SSME is a staged combustion cycle (think of it as turbo-charged) therefore achieving a higher chamber pressure (Pc) than RS-68, therefore producing higher thrust, F, where F = Pc * At * Cf (where At = chamber throat area, and Cf = nozzle thrust coefficient).<br /><br />Also, the propellant used to turn SSME turbopumps have enough energy left to be re-injected into its chamber which means 100% of propellant flow is accounted for in the thrust equation. Whereas IN the RS-68, the turbine exhaust does not have enough energy to be re-injected back to the main chamber therefore is 'less efficient'.<br /><br />Since Isp = F/Wp; <br /><br />where Wp = total propellant flowrate <br /><br />In the case of SSME, the thrust gain requires less additional propellant flow rate than the RS-68. Therefore the SSME has a higher Isp than the RS-68. <div class="Discussion_UserSignature"> </div>
 
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barrykirk

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Well, If that's the case then go to plan B.<br /><br />4 ( 4 or 5 segment ) SRB as Stage one.<br /><br />Stage 2 is 1 or 2 SSME with Inline External Tank, got to love that one. and 2 SRB with 1 or 2 segments. The second stage SRB would not be re-usable.
 
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barrykirk

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I've heard $40Mil for an SSME. Non re-usable version.<br /><br />What is the cost of an SRB?<br /><br />Also, to be fair. What is the cost of an External Tank?
 
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propforce

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<i>"...I'd like to point out that the more segments you put on a SRB, the higher the pressure is generated inside, requiring the casing to be re-designed in a heavier fashion. The goal of the program is to do as little re-design as possible, so more than the already tested 5 segments is out. ..."</i><br /><br />I've heard the same thing. The 5-segment SRB is out for the CLV, but is still in for the Cargo version which is not expected to start designing until 1014. <div class="Discussion_UserSignature"> </div>
 
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barrykirk

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I misstated. I should have said thrust to weight ratio rather than thrust in general.<br /><br />I'm not a rocket engineer so I'm stepping out on a limb of assumptions here.<br /><br />Assumption: The thrust to weight ratio of a solid can be made really high by making it short and fat with a low burn time.<br /><br />Any case, I agree that the SSME is the highest thrust to weight ratio engine in existance in it's ISP range.<br /><br />But, I'm assuming that a limiting factor for it is the low density of LH2 which requires large heavy turbo pumps.<br /><br />Having said that, those turbo pumps are most likely an engineering marvel of super high pump capacity in a small lightweight package.<br /><br />Now I just said they are large and heavy and than I said that they are lightweight and small. What I mean is this. Could those pumps be smaller and lighter if they pumped RP-1 instead of LH2.<br /><br />In other words, could you build an engine with a higher thrust to weight ratio than the SSME if it used RP-1 instead of LH2?<br /><br />Other than the space shuttle, I know of no other vehicle which uses LOX/LH2 at the launch pad. Is this correct?<br /><br />But the shuttle's SSME couldn't even get it off the ground without the SRB's!!!<br /><br />Correct me if I'm wrong, but at the moment of SRB seperation on the shuttle, the three SSME don't even generate enough thrust to accelerate the shuttle. It actually slows down a little until it burns off enough fuel to get light enough for their to be sufficient thrust.<br /><br />I don't know what the tradeoff is. Would the shuttle have more capacity to orbit if it had the weight of a 4th SSME?<br />
 
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barrykirk

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Propforce. I just re-read your post and I missed something. Yes the efficiency is more related to ISP of fuel than thrust.... But......<br /><br />At liftoff efficiency includes more than just ISP of the rocket engine.<br /><br />You take the thrust from the rockets and subtract out the following.<br /><br />1) Gravity losses.<br />2) Drag losses.<br /><br />If I have a rocket that develops 3 LBs thrust for every pound of weight, well the people inside feel 3 G's. But the rocket is only accelerating up at 2 G's.<br /><br />33% of my energy is lost to gravity. That's a 33% loss of efficiency.<br /><br />That is why we don't use Ion engines to lift off the earth. They have an amazingly high ISP but the loses due to gravity would keep it from leaving the launch pad.<br /><br />The idea rocket would have<br /><br />1) Highest possible ISP.<br />2) Lowest possible gravity losses.<br />3) Lowest possible drag losses.<br /><br />Somewhere there is a balance.
 
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rubicondsrv

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"Other than the space shuttle, I know of no other vehicle which uses LOX/LH2 at the launch pad. Is this correct?"<br /><br />No the delta IV medium (1 cbc no srb's) and delta IV heavy (3cbc's no srb's) use nothing but LOX/LH2 on their liquid stages. <br /> <div class="Discussion_UserSignature"> </div>
 
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barrykirk

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Thank you,<br /><br />I was not aware of that.... Cool....<br /><br />From http://www.astronautix.com<br /><br />Delta IV Medium<br /><br />Stage Number: 1. 1 x Delta RS-68 Gross Mass: 226,400 kg. Empty Mass: 26,760 kg. Thrust (vac): 337,807 kgf. Isp: 420 sec. Burn time: 249 sec. Isp(sl): 365 sec. Diameter: 5.10 m. Span: 5.10 m. Length: 40.80 m. Propellants: Lox/LH2 No Engines: 1. RS-68 Status: In production. Comments: Low cost expendable stage using lower performance engine. Used in Delta 4, Boeing EELV. Engine can be throttled to 60%. <br /><br /><br />Stage Number: 2. 1 x Delta 4 - 2 Gross Mass: 24,170 kg. Empty Mass: 2,850 kg. Thrust (vac): 11,222 kgf. Isp: 462 sec. Burn time: 850 sec. Diameter: 2.44 m. Span: 4.00 m. Length: 12.00 m. Propellants: Lox/LH2 No Engines: 1. RL-10B-2 Status: In production. Comments: Delta 3 second stage with hydrogen tank stretch.<br /><br />Interesting <br /><br />The Delta IV first stage has Gross Mass of 226,400Kg with a Vacuum Thrust of 337,807Kgf. Don't know what the sea level thrust is but it should be lower<br /><br />The second stage has a Gross Mass of 24,170 Kg<br /><br />Total Vehicle Gross mass of 250,570 Kg <br /><br />Acceleration should be 0.34 G's at liftoff using vacuum thrust. It will be lower since the engine is at sea level.<br /><br />The gravity losses must be huge!!! Until it burns off some fuel, it really isn't going to move, just hover above the pad. I'm exagerating but not by much.
 
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propforce

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<i>"... Acceleration should be 0.34 G's at liftoff using vacuum thrust...."</i><br /><br />Huh?<br /><br />As a rule of thumb, if the acceleration, a, at lift off is less than the gravity, g, imposed; generally it means the vehicle will not get off the ground.<br /><br />The RS-68 sea level thrust is 650K lbf. I don't know what that is in communist (metric) unit <img src="/images/icons/tongue.gif" />. <div class="Discussion_UserSignature"> </div>
 
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