Orbital Access Methodologies Part III: Pop-up TSTO

This third installation in my Orbital Access Methodologies series (parts I can be found here, and part II here) has been a long time in the coming. It has taken so long, not because I’ve been spending months researching and analyzing the topic (I knew most of what I wanted to say back in January), but mostly because I was surprised by how much favorable attention the first part received, and I’ve been worried about not meeting expectations. A good part of the reason why that first article was so good was that I was able to lean heavily on help provided by Dan DeLong and Antonio Elias, both of who had been analyzing air-launched orbital access methodologies since I was still in gradeschool. I now have a bit more empathy for movie directors trying to make a sequel or a prequel to a first movie that had been far more successful than they had ever thought.

In the previous installation, I discussed approaches to incrementally make ELVs more reusable (or at least recoverable/refurbishable). I discussed why I think that while making ELVs recoverable will be an improvement over the state of the art, such incremental improvements may actually be on a different evolutionary path from high-flight rate capable, truly reusable launch systems. I then discussed the key challenge for TSTO RLVs: how to get the first stage back after a mission, and I outlined the benefit of having the first stage be able to return itself to the original launch site without having to land downrange. This article and the next several in the series will focus on TSTO approaches that provide for return to launch site capabilities.

The first of these approaches, what I like to call “Pop-up TSTO”, has gained quite a bit of attention over the last several years, particularly due to Patrick Stiennon and David Hoerr’s book “The Rocket Company” (which they had me review here, and here). The basic concept is to have a TSTO vehicle, where the first stage flies up purely vertically (John Carmack, who is a fan of the approach has likened the first stage in this concept to a freight elevator) with an apogee of around 100km, the second stage separates from the first stage, and then the second stage provides all of the horizontal acceleration to reach orbital velocity. The first stage reenters and lands vertically like the suborbital vehicles that we at MSS, as well as our friends at Armadillo Aerospace, TGV, and Blue Origin are trying to do. The upper stage after delivering its passengers or payload, reenters and also lands at the launch site.

Benefits of the Pop-Up TSTO Approach
There have been several benefits posited for this TSTO approach:

  1. The vehicle is very operationally simple. The first stage goes straight up, the second stage straight over. You have at most four important engine ignition events (liftoff, 2nd stage ignition, 1st stage landing, and upper stage landing if the upper stage uses powered landing).
  2. If the upper stage T/W ratio is high enough (approximately 1.4) or if the first stage staging altitude is high enough, the first stage ends up soaking up most or all of the typically 1600m/s of losses that an SSTO design would face. This means that the upper stage only has to provide the ~7800m/s needed for orbital velocity, minus ~325-465m/s for the rotational velocity of the earth depending on launch site latitude, yielding a required delta-V of around 7400m/s for most US launch sites.
  3. The upper stage main propulsion system only has to operate in vacuum, so all of the engines can be vacuum optimized, giving much higher mission averaged-Isp.
  4. The upper stage also doesn’t operate for the most part inside the atmosphere, so it might not need slosh baffles (or if it does, they probably don’t have to be as heavy as baffles needed on a lower stage). It also probably doesn’t need anywhere near as much gimbal authority as a 1st stage would.
  5. Staging can be done at high enough altitude that it is a very low dynamic pressure event. Part of what caused the loss of the last Falcon I flight was that the staging ended up occurring at a lower altitude than planned, which imparted higher aerodynamic forces on the stages, which caused a collision between the 2nd stage nozzle and the first stage.
  6. The first stage ends up having performance requirements more like a suborbital launch vehicle than a typical orbital first stage. This means that it’s easier to make it robust and simple, costs can be lowered at times by throwing weight at problems (since the first stage is very weight insensitive). This also means that the first stage could be either evolved from a future suborbital launch vehicle, or at least could possibly be developed by a team that has worked out the challenges of a VTVL suborbital vehicle.
  7. Since the upper stage has such a high delta-V requirement, it will end up having a relatively high propellant mass fraction, which means that when it reenters, it will be mostly empty and will thus be very fluffy. Having a low ballistic coefficient (ie a low mass per unit frontal area) means that you decelerate quicker, higher in the atmosphere where the density is lower–this yields both a lower peak g-loading, but also a lower heat flux, thus making the TPS material challenge somewhat easier than for a dense reentry vehicle like the shuttle or most capsules.
  8. Since the first stage has no downrange velocity, it’s Instantaneous Impact Point stays right around the launch site throughout the flight. This makes it easier to launch over land, out of more populated areas (instead of having to launch along the coasts or from islands or sea platforms out in the ocean). Most of the high-risk phases of flight (ignition, max-Q, staging, upper stage ignition, etc.) happen when the IIP is within spaceport grounds, and thus away from the uninvolved public. This should make it easier to get licenses for the vehicle to operate out of less traditional launch facilities, which may be a key to lowering some of the cost of space access–and to being able to get more customers for said vehicle.

Now, there are probably other advantages, but those are some of the primary ones as I see it.

Challenges, Constraints, Limitations and Drawbacks
Like with the Air-Launched “Assisted SSTO” concept I discussed in Part I, the Pop-up TSTO approach does not come without its own set of problems. There are always both pluses and minuses to all approaches, and the key to good engineering is to make sure you understand what those limitations really are so they can be dealt with properly. Here are a few of the main drawbacks that stick out to me:

  1. Much like the air-launched SSTO rocket stage, the upper stage for a Pop-up TSTO vehicle still faces a nearly-SSTO level of delta-V requirements. Due to the non-linearity of the rocket equation, knocking off 1600m/s vs. a ground launched SSTO makes a huge difference, but providing 7400m/s in a single, reusable stage is still challenging.

    As an aside, many commenters on my air launched SSTO concept seemed to think that such a vehicle wasn’t really technologically doable, but that a Pop-up TSTO stage would be relatively easy to build. I stayed up till 2am doing the math last night, and the fact is that the two are not as different as you might think (I can provide some of the math and explanations if people are interested). The Air-launched SSTO stage needs about 8000m/s (maybe 100-150m/s less for a stage using a more dense propellant combination, or one that has a high thrust to weight at ignition due to using Thrust Augmented Nozzles), compared to 7400m/s for the Pop-up TSTO upper stage. What this equates out to is that for two stages using similar propellant types and similar propellant loads, the pop-up upper stage would only have 20-25% more mass to play with than the air-launched SSTO stage. Specifically for a LOX/LH2 upper stage, you’re talking about propellant mass fractions (the propellant mass divided by the stage plus payload mass) in the range of 0.81-0.82 for the pop-up stage, and around 0.84 for the air launched stage. For LOX/HC, the numbers are around 0.89-0.91 for the pop-up stage, and and 0.9-0.92 for the air launched stage. While that 20-25% more dry mass is nothing to sneeze at, it’s a lot closer than most people would seem to believe.

  2. The upper stage needs a relatively high stage thrust to weight ratio at ignition in order to avoid incurring drag losses (around 1.4 being ideal according to The Rocket Company). While you could theoretically loft the first stage a bit higher to give more time, this quickly starts putting your abort g-loads in the range that is problematic for manned flights. So, you either end up taking a small delta-V hit (thus pushing you closer to the air-launched SSTO case), or you end up taking a mass ratio hit for larger engines.
  3. The upper stage ends up being very sensitive to weight growth. Adding 1 pound to the upper stage could require an additional 20-30lb worth of hardware and propellants on the first stage. This either means designing in lots of performance margin on the first stage, taking a hit to payload, having to spend a lot more money on weight control on the upper stage, or possibly all of the above.
  4. The high delta-V requirements, and the sensitivity of first stage weight to upper stage weight growth push you towards LOX/LH2 or at least LOX and one of the lighter hydrocarbons (cryogenic methane or subcooled propane) for the upper stage. This is typically done by the ELV people as well, but the complexity of adding a cryogenic fuel on-board is annoying.
  5. The typical configuration for a pop-up TSTO is going to be two serially stacked stages, which now requires ground handling equipment for stacking stages. This costs money and makes it harder for a given location to setup a launch site.
  6. Because the delta-V split on the stages is less than optimal, this results in very big first stages (depending on the achievable propellant mass fractions). Which means that as you scale up, at some point you’ll wind up with a stage that’s too big for normal ground transportation. And because RLVs will typically have a much lower payload to GLOW ratio than ELVs, you’ll run into this road/rail transportability limit at much smaller payloads than ELVs do. For instance, if you don’t go with a LOX/LH2 upper stage, even a very light RLV (1-2klb payload) could end up having a first stage that’s as big as a Falcon IX first stage.

    There is one possible work-around to that problem–and that’s having the first stage be modularly assembleable. While I think John takes the modularity concept way too far (I’d never go more than 7, and would generally try to keep it to 3-4 parallel units), and while I’d definitely go with a more aerodynamic module configuration with higher aspect ratio modules than he has, modularity could possibly help with getting around this problem. Think Saturn-IB first stage except having the separate tanks modularly assembleable, instead of preassembled. Sure, it’ll cost you a lot more integration, and a lot more mass for the mechanical, fluid, and electrical interconnects, but your first stage is already fairly weight insensitive. This would allow you to scale up by at least another half order of magnitude, and by that point you’re probably up into the light EELV range–which RLVs won’t be approaching in the near term anyway.

  7. You’ve still got to deal with TPS for the orbital stage.
  8. Because the most likely configuration for a pop-up vehicle is two vertically-stacked stages, the upper stage may need to be able to separate itself from the lower stage in some abort modes. While HTHL vehicles can more readily survive propulsion failures at most points in their flight, VTVL vehicles like the pop-up TSTO would likely be don’t have the option of just dumping most propellants and gliding down to an emergency landing. If you have a full propulsion failure of the first stage, it may require separating the upper stage in a hurry. Since this a reusable stage though, typical expendable launch towers aren’t a practical answer, which involves some sort of reusable escape engines (possibly an aggressive TAN extension to the upper stage primary propulsion system). Testing these and making these abort modes safe and graceful is going to be non-trivial.

Enabling Technologies
Being a less aggressive design approach than the Air Launched SSTO, there aren’t as many enabling technologies that aren’t already on the shelf. Thrust augmentation could possibly be helpful (especially for emergency abort operations), but aren’t necessarily required. Composite propellant tanks and structures could reduce the weight of the upper stage a bit, making it easier to hit mass targets, but the upper stage is probably within the realm of feasibility even using metal tanks and existing manufacturing processes. The first stage development and operations would benefit from the existence and flight experience provided by suborbital VTVL RLVs.

The main enabling technology for this style of RLV is going to be the TPS system (and possibly the reentry technique). There are a couple of interesting options out there that might be doable with such a fluffy reentry stage, such as metallic TPS like was planned for Dynosoar or X-33. And there are some more exotic ideas I’ve heard such as Joe Carroll’s “spike” idea. But the reality is that none of these have been proven out yet, and that’s the only real enabling technology for Pop-up TSTOs that isn’t already on the shelf. It’s important to note that this is the case for all of the RLV techniques I’ll be talking about. There are tons of good ideas, but very limited flight data.

Also, looking back at what I said in Part I, all RLVs could benefit from commercially available prox-ops tugs.

Remaining Unknowns and the Path Forward
Unlike Air-launched SSTOs, there are far fewer unknowns that I can see for this approach. The upper stage is still fairly aggressive, so there’s some questions about if we can make a highly reusable stage with the required performance. There’s still the questions about the TPS. And the other big unknown is going to be how to handle aborts throughout the flight regime. In order for an RLV to make economic sense, you can’t be losing it frequently. Just getting the crew, passengers, and/or cargo out isn’t enough if you can help it. Figuring out how to design a reliable VTVL vehicle that can survive reasonable failures is going to be a challenging task. And figuring out how to perform a rapid separation in possibly adverse conditions without adding so much mass or complexity to your upper stage that you make the vehicle less reliable or unworkable is also going to take work.

The key to moving forward though I think is pretty clear. VTVL RLV companies like us at MSS and our friends at Armadillo and the others need to keep plugging along until we are actually reaching 100km on a repeatable and affordable basis. We need to keep working our way up the learning curve, and hopefully finding businesses along the way to make that possible. Once we’re there (or possibly sooner if XCOR or Virgin beats us to space–which is actually fairly likely), subscale TPS experiments need to be done using suborbital vehicles. This can be done using the “nanosat launcher” suborbital RLV upper stage I mentioned in Part I. By decreasing the cost of actually getting real flight data into the hundreds of thousands range might allow for enough iterations to work out some of the bugs on the small scale before trying to build a full-scale prototype. Also, once a VTVL suborbital vehicle is there, most of us in the industry plan on trying to use our vehicles as a first stage for launching nano-sats. This should help work out the challenges of stage integration, staging, and could even provide an environment for testing out subscale launch escape systems and techniques.

Once all of the subscale work has been proven out with suborbital vehicles, it should be much easier to start into developing a prototype orbital vehicle. There’ll still be a lot of work involved, and there will still be some scaling risks, but by using suborbital vehicles to prove out the various concepts, a lot of the important risks can be retired before its time to start work on a full-up orbital RLV.

The following two tabs change content below.
Jonathan Goff

Jonathan Goff

President/CEO at Altius Space Machines
Jonathan Goff is a space technologist, inventor, and serial space entrepreneur who created the Selenian Boondocks blog. Jon was a co-founder of Masten Space Systems, and is the founder and CEO of Altius Space Machines, a space robotics startup in Broomfield, CO. His family includes his wife, Tiffany, and five boys: Jarom (deceased), Jonathan, James, Peter, and Andrew. Jon has a BS in Manufacturing Engineering (1999) and an MS in Mechanical Engineering (2007) from Brigham Young University, and served an LDS proselytizing mission in Olongapo, Philippines from 2000-2002.
This entry was posted in Launch Vehicles, Orbital Access Methodologies, Space Transportation, Technology. Bookmark the permalink.

24 Responses to Orbital Access Methodologies Part III: Pop-up TSTO

  1. FC says:

    Who prefers Tiesto to Orbital?

  2. gravityloss says:

    1)Where do you stage? You didn’t handle that at all. It’s of course optimal to do that much before the apogee, it’s a question of dynamic pressure I guess.

    2)And how much can you raise the apogee of the “elevator” as that’s the most straightforward way of reducing second stage requirements, before getting too much gee loads during entry? How much would high drag devices help? You want them anyway for low terminal velocity but there they don’t have to take heat, anyway I see them as one of the important neglected VTVL technologies. Unless you have a hydrogen SSTO which is a high drag device by itself. 😉

    3) Earth rotation and pop up launch. How much does the first stage have to maneuver to return exactly at the pad? I’d hunch it’s not a completely trivial thing but involves some math. Ie the ground track of the launch site is a part of a circle that is in the plane of the latitude but the launch vehicle travels an ellipse which has the earth center as one focus so it’s in an inclined plane and thus re-enters at different latitude. Or then not?

    4) And additionally, It’d be cool if there was somewhere some nice physics based explanation of that “fluffy high altitude deceleration gives a low thermal load” rule of tumb (I know it’s very common), because it isn’t completely intuitive, ie a first physics based assumption would give same total thermal load for both cases. Of course the load per area and the peak heating are different.

  3. gravityloss says:

    And by these questions I mean, an excellent article! It is, if it raises so many questions.

  4. Dave Salt says:

    Excellent post (again) Jon!

    One more thing that this architecture may lend itself to better than most others is some form of technology augmentation (e.g. just imagine if Jordin Kare’s laser modules were used to “improve” the Isp of the booster).

    Dave

  5. redneck says:

    >>If the upper stage T/W ratio is high enough (approximately 1.4) or if the first stage staging altitude is high enough, the first stage ends up soaking up most or all of the typically 1600m/s of losses that an SSTO design would face. < <
    ==================================
    Gravityloss mentioned staging location which I think might alleviate some of the T/W requirements for the upper stage. If you are heading to 100 km, you have about 1,000 m/s velocity as you pass 50 km altitude. If the upper stage fires from there, it has 100 seconds of upward travel before it starts falling. The T/W ratio after burning 100 seconds of propellant should put it toward the 1.4 you suggest. Even a modest vertical angle can stretch that 100 seconds to 200 seconds. With hydrogen starting from T/W of 1, that is 44% of initial mass burned before it starts falling back. A designed trajectory should do much better than that. If you stage from 100 km and 0x0 velocity, then you will certainly need the 1.4 T/W.

    ===================================
    >>Adding 1 pound to the upper stage could require an additional 20-30lb worth of hardware and propellants on the first stage.< <
    ==============================

    I’m not following your math here. It would seem that a conservative design for the pop up stage would have a dry mass of 15% and a M/R of 4. That leaves 10% for the upper which would be half to a third of the mass you suggest.

    I see my mistake here. I pound of upper stage ‘hardware’ times upper stage mass ratio times lower stage hardware and mass ratio. Very sensitive and I missed it on the first two readings.

    =================================
    >>The high delta-V requirements, and the sensitivity of first stage weight to upper stage weight growth push you towards LOX/LH2 or at least LOX and one of the lighter hydrocarbons (cryogenic methane or subcooled propane) for the upper stage. This is typically done by the ELV people as well, but the complexity of adding a cryogenic fuel on-board is annoying.< <<
    ============================
    You made me get out my calculator. I am opposed to hydrogen for the most part. Using 4,400 and 3,400 m/s as the exhaust velosity for H2 vs Kero, I get M/R of 5.38 and 8.8respectively. To get one ton in orbit that is not engine or tanks, I get 8,877 kg for the H2/Lox stage vs 12,144 kg for the Kero/Lox stage. That is using T/W of 75 and 125 for the two engine types, and 20 kg/cubic meter for all tankage. I think it might be an easier engineering problem to make the first stage 37% larger to avoid dealing with hydrogen in the second. YMMV

    Good post, except it made me have to think.

  6. redneck says:

    To further my anti-hydrogen bias I resurrected an arguement from years back. If you take the hardware mass of the H2 upper stage and use it for a Kerosine upper stage, you can get enough additional upper stage performance to shrink the lower stage for a lower over all mass.

    Check any numbers I use that interest you. It has been a while since I did this one. I.m using your 7,400 m/s number.

    On the second pass I get 8800 kg for an upper stage using hydrogen getting one ton of non propulsion mass to orbit. A 164 kg engine producing 12,300 kg of thrust. 7164 kg of propellant in 23.64 cubic meters of tank massing 472 kg.

    If you use the 636 kg of propulsion system mass for kerosine engines and tanks, you can have a 254 kg engine producing 31,750 kg of thrust. 382 kg of tanks holding 19 tons of propellant.Though the stage now masses 11,936 kg more than the hydrogen stage, the hardware mass is the same, and the available V now exceeds 8,600 m/s.

    Off load 11,936 kg of propellant from the first stage and ask it to give you an ideal 400 m/s including losses instead of the 3,000 m/s for the hydrogen varient.

    It would be easier to balance them better than this, perhaps 2,000 m/s from the pop up to make the upper more feasible.

    If the absolute maximum mass of an upper stage is fixed, then hydrogen wins. I believe that any stage to LEO that is not hard fixed by maximum mass, will best be served by denser fuels.

  7. Anonymous says:

    Hey Jon its Aaron, I think we talked about this last summer, but what is wrong w/ the idea of bypassing the vertical landing part(at least for now) and using a gps guided ram air parachute to bring the first stage back near the pad(possibly w/ airbags for easy/cushy landing). Seems like that would bypass a major hurdle of trying to get a first stage to hover and let you move on w/ getting a vehicle flying, then you can come back to it as your flight rates increase, which is when vertical landings really see their benifits. Of course one problem is that it would take a freakin big parachute, but are there others?

    Aaron

  8. Mike Puckett says:

    Why not have the 1st stage give it some horizontal delta V and recover it downrange 50 to 100km away?

    The 1st stage could be self-ferrying, just add fuel and fly it back or use a Sky Crane to ferry it?

    It it really that big an operation advantage to immeadiately return to the launch site?

  9. Michael Mealling says:

    Mike,
    Given re-entry insurance requirements and how MPL is calculated based on population, etc plus the logistics requirements, yes, it probably is. Maybe once the actuarial tables are sufficiently backed up with real work data you could convince the insurance companies to lower the price but its still the logistics that cost money.

  10. Anonymous says:

    Jon:

    Great post again!!

    Why not add stubby wings, a TPS, and some additional RL-10 engines (higher T/W) to a Centaur upper-stage, and see if you have the mass fraction to get this to orbit. You can probably buy all the hardware for much cheaper than you think, and maybe even do drop tests from White Knight 1 if the unfueled weight is under 9,000 lbs.

    I think that a 40,000 lb Centaur upper-stage imparts around 8 km/s of delta-V to a 10,000 lb payload when it takes those payloads from first-stage cut-off to GTO (i.e. 6 km/s + 2km/s = 8 km/s). This 8 km/s appears to be enough by your calculations for an air-launched RLV or a vertical platform assisted RLV to reach orbit. The Centaur weighs 4,000 lbs unfueled and the additional hardware should be under the 9,000 lb carrying limit of White Knight 1.

    It appears that all the parts are available and cheap for you to do an X-37 type of landing test from White Knight 1 of a full-scale model of your concept.

    What do you think about having DARPA pay for it like they did with X-37 drop tests?

  11. Jon Goff says:

    Gravityloss,
    1) Exactly, it’s a tradeoff. The sooner after shutoff you stage, the more time you have to accelerate to orbital velocity, but the higher the staging dynamic pressure. So, it’s a tradeoff.

    2) I can’t remember the exact numbers, but there’s a reason why none of the suborbital tourism guys are planning on going much higher than about 170km (IIRC). Basically, for a pure vertical trajectory, almost all the decelleration starts at a certain altitude (can’t remember if it was 50km or 50kft). And the higher you reach, the faster you’re going when you hit that point. You can get a rough guess at peak g’s just by treating it as pure vacuum above that point, figuring out how fast you’d reach dropping from apogee, then calculating the dynamic pressure times your frontal area.

    There are some tricks you can use, but IIRC, by the time you’re getting up to 200km or so the G loads are already quite high (something like 8-10g)…but that’s something you could probably calculate easier than I could.

    3)I’m not entirely sure if this will be an issue or not. Our vehicle simulator so far has been focused on our early hovering vehicles. But it includes all the physics, so if there is such an issue, it should become apparent when we try simulating higher apogees. I’ll let you know if we find anything interesting.

    4)One of the textbooks we have at the office explains it. But yeah, lower peak heat flux and lower load per unit area are both quite helpful even if the overall heat load ends up not being too different. But I’ll look it up.

    ~Jon

  12. Jon Goff says:

    redneck,
    Any “vertical angle” you give the stage starts reintroducing gravity losses. Ie some component of the velocity is now counteracting gravity, and not going to accelerating the stage. So you’re not actually gaining anything. I’m not entirely sure how much time you gain by raising the first stage apogee (thus allowing a lower upper stage T/W and longer acceleration to orbital velocity). I’ll have to see if I can figure out an approximation of the relationship.

    As for the hydrogen vs kerosene argument, it’s worth noting that AFAIK, the best LOX/LH2 propellant mass fraction is only a tiny hair worse than the best LOX/Kero pmf. Ie, it at least appears historically that the density differences between the two propellants might not help as much as seems to be commonly argued in our circles. It could be that they just aren’t trying as hard with the LOX/Kero stages, but the more likely answer is that there are other effects that are being glossed over…not sure.

    ~Jon

  13. Jon Goff says:

    Aaron,
    Vertical landing isn’t that much harder than stable controlled flight in general (unless you’re talking going with an HPR styled vehicle that uses fins, launch rails, and high liftoff T/W). It isn’t that hovering is so tough that we could just bypass it, it’s that our vehicle wasn’t flying stably at all anyway. Once we have enough stability for the thing to fly without a launch rail or fins, the extra challenge of hovering and powered landing is relatively straightforward. Plus, using a parachute means you have to reach at least a certain altitude first…for our purposes, working out powered vertical landing seems to be a better trade for now than doing basically a liquid HPR then adding powered landing later.

    ~Jon

  14. Jon Goff says:

    Mike,
    I talked about that a little in part 2 of the series. But basically, especially if you’re in a high-flight rate operations mode, down-range recovery implies additional ground crews, additional areas setup as launch sites, additional launch licenses, additional insurance as Michael points out, etc. There are ways for the first stage to impart downrange velocity while still landing back at the launch site. But that’s the topic of the next two or three (or maybe four) parts of the series.

    ~Jon

  15. john hare says:

    Jon Goff said…
    redneck,
    Any “vertical angle” you give the stage starts reintroducing gravity losses. Ie some component of the velocity is now counteracting gravity, and not going to accelerating the stage. So you’re not actually gaining anything. I’m not entirely sure how much time you gain by raising the first stage apogee (thus allowing a lower upper stage T/W and longer acceleration to orbital velocity). I’ll have to see if I can figure out an approximation of the relationship.

    ===========================
    I will try to figure out a coherent relationship in a few days for the velocity split. I believe the gravity losses are less if you begin thrusting earlier with a bit less vertical angle. I will try to organize that also when I feel like it. Just read about Len.

    =================================

    As for the hydrogen vs kerosene argument, it’s worth noting that AFAIK, the best LOX/LH2 propellant mass fraction is only a tiny hair worse than the best LOX/Kero pmf. Ie, it at least appears historically that the density differences between the two propellants might not help as much as seems to be commonly argued in our circles. It could be that they just aren’t trying as hard with the LOX/Kero stages, but the more likely answer is that there are other effects that are being glossed over…not sure.

    ==================================
    You are one of the people that can get me to reexamine my assumptions with a single sentence.
    =============================

    ~Jon

  16. paradoxolbers says:

    Wonderful set of posts on OrbAccMethods! A trivial text change for u. In your first set of bulleted points, under Benefits, number 8 should read:

    Since the first stage has *no* downrange velocity, its

    best,
    Paradox Olbers in Second Life

  17. Jon Goff says:

    Paradox,
    Good catch! Glad you’re enjoying the series.

    ~Jon

  18. john hare says:

    Jon,

    I think my main assumption problem was that your pop up stage would fire from zero-zero horizontal and vertical velocity at 100 km altitude. From that assumption I can’t get under 400-500 m/s gravity losses even with T/W of 1.4.

    Firing from 80+ km with 500+ m/s vertical, it is easy to get under 200 M/S gravity losses with the 1.4 T/W. To eliminate the gravity losses entirely seems more difficult that staging at 50 km with 1,000 m/s vertical velocity and <100 m/s gravity losses.

    Your observation that hydrogen stages nearly match kerosine stages make it a bit silly for me to attempt to figure an ideal velocity split. I don’t have enough knowledge to spin this portion into a cost arguement.

  19. SpaceCadet says:

    Nice article (and series)! TSTO popup has been my preferred option for a while.

    Have you considered expendable drop tanks on the second stage? It’s a mature technology for military aircraft. I wouldn’t use them for LH2, too much like the Shuttle with ice & foam issues.

    But for HC, maybe 30 to 40% of the fuel tankage could be outside the RLV, making it smaller & lighter. That plus losing part of the tank mass halfway to orbit, should give a fairly substantial gain.

    Drop tanks should be quite cheap, compared to any other part of the dry mass.

  20. Jon Goff says:

    Spacecadet,
    While you could add drop tanks, I think the best way to use them would be to use them the way military fighters do–as a way of enhancing capability. I don’t think there’s any military fighter craft that has to have drop tanks to perform its bare minimum mission. So, you make an orbital vehicle that can put a small payload up without drop tanks, but you use drop tanks for the occasional much bigger payload, or for giving you new capabilities you couldn’t have had with drop tanks. Even if they’re relatively cheap (lightweight tanks aren’t really that cheap, especially in orbital LV sizes), you don’t want to be tossing them every time if you can avoid it. But every once in a while? Sure. Of course that might require you to oversize your first stage, but that’s not necessarily the end of the world–nothing says you need to always load a full load of propellant if it isn’t needed.

    ~Jon

  21. Ivan Vuletich says:

    Hi,
    Great series. One idea to help with the reentry TPS challenge that I’ve been kicking around is a sacrifical “reentry drag chute”.

    This would be some sort of parachute that you would deploy at the beginning of reentry and whose function is to provide extra drag as high as possible during rentry before buring up, thereby allowing a lighter & simpler main TPS.

    In form it might look more like a solar sail than a conventional parachute. See

    http://www.spacedaily.com/reports/NASA_To_Attempt_Historic_Solar_Sail_Deployment_999.html
    In that article they mention one of its uses would be to de-orbit LEO satelites.

    Alternatively maybe a fine mesh of stands of an ablative material might work better.

    If it works as advertised it could make for an easier TPS problem, reduce g loads on re-entry and help maintain the proper orientation on reentry. It might be able to help with some about cases as well.

    Its also something that could be tested with sub-orbital vehicles 🙂

    Key challenges would be to:

    1) Getting enough usable drag before it burnt up.

    2) Getting it deployed and keep it deployed in near vaccuum conditions as it would need a lot of surface area to be usefull.

    3) Making it light and cheap enough to make using it worthwhile.

  22. Jon Goff says:

    Ivan,
    I honestly don’t know whether your idea would work or not. I think Henry Spencer has suggested such an idea at one point, so it isn’t silly. I just don’t have the tools to easily tell if it really makes sense. That said, I’m a fan of avoiding expendable elements, particularly ones that may be fairly expensive to replace like this. One thing to remember though, the g-loading on reentry is almost entirely dependent on L/D. So unless the parachute somehow gave you a high L/D, it wouldn’t really affect the g-loading that much. Also, I wonder how effective it would be if the upper stage is far ahead of the parachutes (as is often the case).

    Just some thoughts.

    ~Jon

  23. Anonymous says:

    I agree, L/D drives the peak g’s, but some sort of high drag device at initial reentry could reduce the peak heating even if it’s done in a ballistic fashion.

  24. Pingback: Random Thoughts: New Shepard for Pop-Up TSTO NanoSat Launch | Selenian Boondocks

Leave a Reply

Your email address will not be published. Required fields are marked *