Reusable Rockets to Orbit

GW Johnson · 2012-01-21 10:13:04

Rune:

Yeah, I kinda forgot about land landings. Old guy. Sorry. Most of the launch sites currently operating launch out over the oceans, except the Russians. I am more familiar with what we did than I am the Russians.

The Russians do land landings all the time with Soyuz capsules, and the Vostok/Voshkod series before Soyuz. They also developed a way to deliver tanks to the battlefield by air. It combines parachutes with last-second retro-rockets, and it has worked for decades.

That kind of thing is steerable by cg shift during re-entry hypersonics, not after. You're pretty well ballistic, coming down on a drogue, then some sort of main chute cluster, and those tend not to be steerable. Pinpoint landings are impossible, you will always be faced with recovery operations, and that's not cheap.

The sudden stop in the water at 20-40 mph is bad enough, on land it was unsurvivable, unless you did one of three things: (1) energy absorbing one-shot crush seats, (2) last second rocket braking, or (3) have the crew bail out on their own chutes before impact. Vostok and Voshkod did item (3). Mercury and Gemini simply ruled out land landings. Apollo and Soyuz now do item (1). I think some of the early Soyuz's might have used item (2), but I may not be remembering correctly after all these years.

But rocket braking works if you combine it properly with parachutes. The trick is to use the highest thrust you can tolerate for the shortest possible burn at the last seconds before impact. That's a serious control and altitude/descent rate measurement problem. Otherwise, you have to carry a huge kitty of heavy landing propellant, so you can just sort of "slop" your way through it. The Russians did this for parachute/rocket landing of heavy tanks to the battlefield from high-altitude transport aircraft.

Spacex is going for the same scheme with land landings of Dragon, I believe. Not sure whether the Dragon would have legs to protect the heat shield, it might. But it comes down on a chute, then fires the Draco thrusters to slow for touchdown, last second or so. You just have to plan on landing with enough thruster propellants to cover your needs with a margin of safety. Of course, those propellants are a crew hazard after landing. Hydrazine and some oxide of nitrogen, I believe.

GW

RGClark · 2012-01-22 08:54:04

JoshNH4H wrote:

Hop, actually I did do the math. I did the math, in fact, on several different stages, because obviously not all are capable of making orbit, in theory or otherwise. A short post does not necessarily indicate a small amount of work having gone into it. The astronautix article on the Saturn V listed information for the Saturn V second stage that is different from that in the article to which you linked. I suspect that my link is correct because of the following text in yours:
Final common second stage design for Saturn C-3, C-4 and C-5 (November 1961). Developed into Saturn V second stage.
Also, it is silly to use the sea level Isp to approximate the Isp for the entire trajectory. As I said, it could in theory function as an SSTO vehicle so long as the trajectory averaged Isp were at least 380 s. This is pretty reasonable, IMO.
By the way, I'm not "continually ignoring" anything. I made a correct statement. You went off on a tangent for no reason. I offered something that would allow us to get back on track. You want to continue this pointless tangent. Why?

You do have to keep in mind though that for these upper stages the nozzles were optimized for high altitude, near vacuum performance. So they are much longer than the nozzles used for launch from the ground. Vacuum optimized nozzles give much worse performance at sea level. In fact they could cause instabilities in the exhaust that can even damage the engine.
To use this stage for ground launch you would have to have some method of altitude compensation, or switch out the engines for the SSME's.
BTW, here is a discussion of using the Saturn V third stage, the S-IVB, as an SSTO with altitude compensation on the engines as an SSTO:

Douglas_SASSTO.
http://en.wikipedia.org/wiki/Douglas_SASSTO

Bob Clark

RobertDyck · 2012-01-23 12:34:30

Ok, landings.

Yes, Vostok had the pilot bail out. However, Soyuz used a rocket mounted on the parachute cords to slow descent last minute. The seats are designed to absorb impact because even with the last minute rockets it's still a hard landing. I saw an interview with a Western astronaut that landed in a Soyuz, he said it felt like a car crash. But without the rockets it wouldn't be survivable. And Soyuz still uses that system today.

The X-38 was designed to be the Crew Recovery Vehicle (CRV) for ISS. It used the same body shape as X-24A, but since that body shape has control problems at low speed (such as landing) they added a large steerable parafoil. They had to invent the largest parafoil ever to do that. X-38 had landing skids for the final impact with ground, skids that deployed beneath the craft. I don't know for sure, but I suspect those skids had some sort of shock absorber. After all, the skids deploy just before impact so it would be trivial to mount them on a shock absorber. That parafoil was developed jointly with the US military, they can now use it to air-drop a HMMWV (Humm-Vee). The HMMWV is mounted on a wooden pallet with significant cardboard between the pallet and the vehicle. The vehicle doesn't have anyone in it while it's air-dropped, so the cardboard is enough to ensure no damage to the vehicle. I believe they can drop an armored personnel carrier or light tank that way as well.

HL-20 was designed to land on wheels like the Shuttle. It was developed after the Challenger accident, in case Congress would not authorize Shuttle to return to flight. Development halted when Shuttle was. Dreamchaser is now designed after HL-20, consider it to be continuation of HL-20. Like all lifting bodies, it has control problems at a specific speed range; however this one is designed so it's control problems are between mach 1.1 and 0.9. That means when crossing the trans-sonic barrier, just fly straight without attempting a turn. It's control problem is while there's nothing but a cloud to collide with, not the ground.

The first public announcement of Orion included ground landing. This would be done with a rocket fired last minute, and attached to parachute cords just like Soyuz. However, instead of landing hard it would have an air bag between the heat shield and capsule. That was later deleted, returning to a water landing like Apollo.

Apollo never did land on ground. Apollo, like Mercury and Gemini, "landed" in the ocean. They had a navy aircraft carrier completely with it's entire battle group pick them up. The first Mercury missions had a helicopter pick the capsule out of the water, using a hook to grab a cable on top of the capsule. It was gently set down on the carrier deck, and only then was the hatch opened. Later they had navy divers "rescue" astronauts out of their capsule, and brought the empty capsule to the carrier separately.

The reason for a land rather than water landing is cost of the pick-up. If you land on ground, you only need a flat bed truck with a truck crane to pick up the capsule. Add a van or HMMWV to pick up astronauts, perhaps paramedics to be safe, and if you're paranoid a car/van/jeep/HMMWV with guards/marines to protect them. That costs a lot less than an entire aircraft carrier battle group.

Dragon has publicity videos showing it landing with rockets embedded within its sides, near the bottom edge just above the heat shield. It uses a parachute, then last minute uses its rockets, and deploys 4 legs just as the rockets fire. That's publicity and what they hope for in the future. Announcements through NASA state their one successful test (COTS 1) uses a water landing like Apollo. However, they didn't send an entire aircraft carrier group to pick it up, just a fishing trawler. They used the trawler's crane for nets to fish the capsule out of the sea. Its next test will also use an ocean landing. SpaceX has asked to combine test flights COTS 2 and COTS 3 into a single test. Test known as "Commercial Off The Shelf" (COTS) 2, has been further contracted to "C2". However, NASA has been playing with definition of "COTS" because they're actually paying for some development. Anyway, C2 would fly a Dragon within 2 miles of ISS then return to Earth. C3 would rendezvous with ISS, then the station's arm would grab it and dock it to a "Common Berthing Mechanism" port. SpaceX's request is if all goes well with C2, it would proceed with rendezvous and docking. It has been postponed because of unspecified work for "safety". However, the Russians have been complaining about combining these tests ever since their Progress cargo ship failed. And Boeing/Lockheed-Martin are working on MPCV (formerly known as Orion) which would also be a direct competitor of Dragon. They have also raised "safety" concerns about their competitor. So what are these "safety" concerns? Pardon me if I question their legitimacy.

Soyuz uses solid rockets just its landing, so no concerns about MMH, UDMH, or N2O4. Oh, the Apollo Service Module used UDMH (Unsymmetrical DiMethylHydrazine) as fuel and N2O4 (dinitrogen tetroxide) for oxidizer. Soyuz did in the 1960s and still does today. Shuttle used MMH (MonoMethylHydrazine) and N2O4 for its OMS and RCS. But proposals for Orion to land on ground also would have used solid rockets. Dragon uses MMH and N2O4. But it worked for Shuttle, so why not Dragon?

GW Johnson · 2012-01-23 14:48:51

Landings:

I quite agree with RobertDyck. I'd add these things to the discussion:

The main problem with lifting bodies, aside from that speed where control is "iffy", is the high landing speed due to very high wing loading or broadside ballistic coefficient. A vehicle designed to fly into space and back is inherently still a bit heavy on landing because of the heat shield and the control propellants. Most of the those lifting body designs had touchdown speeds at Edwards in the 200-300 mph range. Stability on the ground during roll-out (skids or wheels) is a serious issue, unless the vehicle is rather long, like X-15, which landed fairly well at 200 mph.

Gemini was originally a Rogallo wing-type parasail design, with skids for dry lake bed landing. My memory may be off, but I think the landing speed was near 200 mph like X-15. The Rogallo was steerable, but the short "wheelbase" of the skid system was unconditionally unstable on the ground at any speed. Test articles always flipped over and went bouncing down the dry lake. Good thing they weren't manned. It just never worked, so they went with a parachute and water landing. The capsule was already in production, so that's why the chute came out the side instead of the nose on Gemini.

I've got no real problem with MMH N2O4 residuals after landing. Just don't go sniffing the rocket nozzles for a while. I remember NASA did have concerns, when the shuttle first flew, but they got over being nervous about it after a while.

The Draco thrusters on Dragon are rather powerful, and it has a lot of delta-vee built into it. If one reserves a kitty of propellants, one could fire them last second to slow a chute landing on land to survivable levels for the crew. It would feel like a car crash, just like Soyuz. If it had shock-absorbing legs, the heat shield would survive for re-use, too. Of course, they're extra weight. Reduces payload or crew size a tad. But the water landing version is said to be able to carry a crew of 7.

GW

Last edited by GW Johnson (2012-01-23 14:51:07)

Rune · 2012-01-23 18:16:24

About recovery operations, parachutes, and Dragon:

First off, I agree with most of what you two say, let that be clear! We are so much in agreement, actually, that I feel compelled to point out something, if only to keep the discussion going.

So the first thing to point out is that you both assume Dragon is going to use the parachutes when/if the manned version flies (I'd say when, but you know). That doesn't really seems to be the stated intention of the company, if you go by their CEO's public claims. IIRC, of course, Mr. Musk has been saying in a couple of different webcasted press conferences (just google for one that mentions the reusable Falcon) that the manned version is going to have "pinpoint accuracy, on the order of meters". Now we all now that if you go in by chute alone, you get a dispersion parabola because of it, and no "last-second" engine firing is going to reduce that by any measurable fraction. That kind of just rules out the pad landing shown in the famous video, too.

Also, I distinctly remember him pointing out that a parachute, singular, would be retained to provide a backup and be used in the case of an aborted ascent. I remember, because that's when I found out the Dragon of today is designed for a double failure in the chute system. If that is the case and you read him literally, then that "backup" means the chute won't open during a nominal landing. Full Heinlein-like magic, which would cut the recovery cost to basically nothing. But a nasty environment for plants to grow near the pad, for certain.

Now I imagine I should explore a bit the practicality of it all, so there I go. First off, how much fuel can Dragon take, in excess of what it carries now, which I will assume is enough for a nominal mission plus reserves up to landing? Well, the downmass of the uncrewed one is 3,000kg. We should take away the weight of up to the seven stated crew members, so 700kg? Cal it one ton to account for life support equipment. The additional engine weight of the super-dracos that are under development is a full mystery right now, since I can find nothing on their T/W ratio. But 1mT seems enough, right? Even considering tankage, that's not asking much more than 30 out of it, since a fully loaded Dragon can't exceed ~10mT. First order of magnitude work here, expect it to be more than a bit off, of course, but I'm taking the pessimist side. For Isp, who the hell knows, but similar engines give about 300 vacuum, and data on sea-level isp is pretty much impossible to find, so I'll just pull 275 out of my... let's call it hat (Encyclopedia astronautica, you have failed me!). Straight good old Tsiolkovsky gives about 285m/s out of that, or a shade over 1000km/hour. So... yeah, if Dragon goes subsonic on it's own without chutes, then I call it plausible.

But then again, I have no freaking idea about that, so I take it all with a big grain of salt. And invoke the knowledge of my elders to fill the blanks I left, too.

Rune. I would kill for a look at those super-Draco designs, my curiosity is killing me on this. It would look sooo cool and sci-fi-ish.

GW Johnson · 2012-01-24 23:16:00

Rune:

I know nothing about any "super Draco" thrusters for Dragon. The ones they have are supposed to be the launch escape system for any manned Dragon, so capsule T/W is substantially larger than 1. For the Mars mission paper I gave at last August's convention in Dallas, I re-engineered some data from the Spacex website for Dragon as they had it posted last summer. I was showing 0.9 km/sec delta-vee capability with 6 suited astronauts on board. If you put more propellants in the unpressurized module, and connected that to the Draco system, it adds around 1.4 km/sec more to the capsule's delta-vee capability. I looked at that to get 2+ km/sec total delta-vee for Dragon as an emergency escape crew return vehicle from a very high-speed Mars return in an emergency, somewhere well above 15 km/sec. Maybe as much as 25 km/sec.

Stock Dragon post reentry has no unpressurized module. So the capsule's delta-vee capability is around 0.9 km/sec utter max, post re-entry. That's enough to land from considerable height without a parachute, but it's nowhere near enough to make the entire landing without a chute. Not on Earth. Moon? Not even there. Any "pinpoint" to the landing is during that final burn, but it's not all that much maneuvering. You know advertising flacks, they do tend to oversell stuff.

On the other hand, Dragon's true significance is that it is the very first capsule in history intended to be re-used, heat shield and all. It does need a new unpressurized module and a new nose cap for each flight. But, that heat shield was designed for a free return from Mars at about 50,000 feet/sec [15 km/sec] entry velocity. You could re-use it even after a return from the moon (36,000 feet/sec [same as 11 km/sec] for comparison). Coming back from LEO at around 26,000 feet/sec [8 km/sec] is by comparison "easy", since the heating is proportional to velocity squared. I'd guess they could fly the same heat shield to LEO about 4 times before replacing the pieces that make it up. And it is segmented.

None of Dragon's capsule-type competitors are intended to be reusable. Not Orbital's Cygnus, not Boeing's scaled-back Orion. While my contention is that Spacex achieved its lower costs per unit of payload to LEO by smaller logistical support tail, not actual reusability of anything, having a capsule you can fly several times is still going to help significantly. Plus, it's simply a technical capability that we need. And we have needed it for a long time. It is a major accomplishment.

I'd have to go find my original notes for that re-engineering I did on Dragon to find the Isp I used for the Dracos. But I think it was likely closer to 270 sec than 300 sec. There's quite a gap between what is theoretically-possible under perfect expansion, and what can actually be had with a real engine bell, even in vacuum.

GW

Last edited by GW Johnson (2012-01-24 23:25:26)

Rune · 2012-01-25 07:28:22

GW Johnson wrote:

I know nothing about any "super Draco" thrusters for Dragon. The ones they have are supposed to be the launch escape system for any manned Dragon, so capsule T/W is substantially larger than 1. For the Mars mission paper I gave at last August's convention in Dallas, I re-engineered some data from the Spacex website for Dragon as they had it posted last summer. I was showing 0.9 km/sec delta-vee capability with 6 suited astronauts on board. If you put more propellants in the unpressurized module, and connected that to the Draco system, it adds around 1.4 km/sec more to the capsule's delta-vee capability. I looked at that to get 2+ km/sec total delta-vee for Dragon as an emergency escape crew return vehicle from a very high-speed Mars return in an emergency, somewhere well above 15 km/sec. Maybe as much as 25 km/sec.

Well, they are the escape system under design, I believe. The current Dracos are tiny 400N thrusters optimized for vacuum for reaction control of the capsule and the second stage, so if they want to give the whole capsule T/W over 3 (and probably closer to six, to ensure escape) they need a new engine. I can't find the reference, but Mueller (the propulsion VP) has dubbed them "super-Dracos" and said they are under development. I think it was the conference where they presented the Merlin 1D?

GW Johnson wrote:

Stock Dragon post reentry has no unpressurized module. So the capsule's delta-vee capability is around 0.9 km/sec utter max, post re-entry. That's enough to land from considerable height without a parachute, but it's nowhere near enough to make the entire landing without a chute. Not on Earth. Moon? Not even there. Any "pinpoint" to the landing is during that final burn, but it's not all that much maneuvering. You know advertising flacks, they do tend to oversell stuff.

That was kind of my main question, what is Dragon's terminal speed at sea level? And I know I am actually one of the most equipped people here to work that out (or at least I should), actually, but I was feeling a bit lazy ^^'. Plus, I don't have a clue about what Cl to assign to a ballistic capsule to get lift, for starters. Maybe I will start a very extensive google search for facts one of this days.

GW Johnson wrote:

On the other hand, Dragon's true significance is that it is the very first capsule in history intended to be re-used, heat shield and all. It does need a new unpressurized module and a new nose cap for each flight. But, that heat shield was designed for a free return from Mars at about 50,000 feet/sec [15 km/sec] entry velocity. You could re-use it even after a return from the moon (36,000 feet/sec [same as 11 km/sec] for comparison). Coming back from LEO at around 26,000 feet/sec [8 km/sec] is by comparison "easy", since the heating is proportional to velocity squared. I'd guess they could fly the same heat shield to LEO about 4 times before replacing the pieces that make it up. And it is segmented.
None of Dragon's capsule-type competitors are intended to be reusable. Not Orbital's Cygnus, not Boeing's scaled-back Orion. While my contention is that Spacex achieved its lower costs per unit of payload to LEO by smaller logistical support tail, not actual reusability of anything, having a capsule you can fly several times is still going to help significantly. Plus, it's simply a technical capability that we need. And we have needed it for a long time. It is a major accomplishment.

Yup, same feeling here. Orbital-grade aerospace material is the most expensive part of a rocket, by weight. Just one tiny bit to point out, HL-20's older brother Dream Chaser is intended to be reusable, and runway landing at that. IMO it is the only realistic contender for crew transport besides Dragon and CST-100, if they can get funding and ULA doesn't screw them over when integrating on the Atlas.

GW Johnson wrote:

I'd have to go find my original notes for that re-engineering I did on Dragon to find the Isp I used for the Dracos. But I think it was likely closer to 270 sec than 300 sec. There's quite a gap between what is theoretically-possible under perfect expansion, and what can actually be had with a real engine bell, even in vacuum.

Good to know my hat-pulling wasn't that far from yours, at 275. I just saw the better-performing vacuum engines got around 300sec, and assumed the higher required T/W and sea-level expansion would put them closer to the lowest-performing, or even lower. I imagine I could have used the numbers form a first stage engine using UDMH/N2O4, similar enough, but I didn't want to fall into the trap of thinking a similar fuel means a similar engine when I know little about the differences. Plus, russian magic would have made as much as 285s isp possible at sea level at T/W:136 (RD-253-14D14: Glushko N2O4/UDMH rocket engine. 1746 kN. Proton KM-1. In production. Developed in 1990s. Isp=317s. First flight 1999.)

Rune. Good old legendary soviet design bureaus, they did wonders with engines, didn't they?

RGClark · 2012-01-26 00:39:24

RGClark wrote:

GW Johnson wrote:
To answer a question Josh asked earlier, I spent some time at what was then LTV Aerospace in Dallas working on the old "Scout" launcher. I used a combination of jiggered rocket equation stuff and motor manufacturer catalogue data to set up the real trajectory code stuff. The "gold standard" was (of course) the trajectory code. My job was to determine feasible advanced configurations for "Scout", and feasibility of some really unusual missions for it to do. "Scout" was a 4 stage solid propellant vehicle. They lost 1 of 4 in flight test, then never another one in 30-some years.
For Bob Clark: airbreather thrust, particularly ramjet, is very strongly (dominantly) dependent upon flight speed and altitude air density. The nozzle thrust is calculated same way as a rocket (chamber total pressure, gas properties, pressure ratio across the nozzle, and nozzle geometry), the pressure is just lower and the expansion ratio a lot less. You do need to worry about the difference between static and total chamber pressure, unlike most rockets.
The ram drag is the drag of decelerating the ingested stream of air into the vehicle. Its massflow multiplied by its freestream velocity (in appropriate units of measure) is the way that is done. But, nozzle force minus ram drag is only "net jet" thrust. There are several more propulsion-related drag items to account.
There is spillage drag for subcritical inlet operation (which also means reduced inlet massflow!), additive or pre-entry drag for ingested stream tubes in contact with the vehicle forebody, and the drag of boundary layer diverters or bleed slots, quite common with supersonic inlets. None of those are simple to calculate "from scratch" (we use wind tunnel test data to correlate empirically a coefficient for each as a function of Mach and vehicle attitude angles), and taken together they are often quite a significant force.
If you subtract that sum of drags from net jet thrust, you have the "local" or "installed" thrust, corresponding with just plain airframe drag. Most airframers work in that definition. If you don't, then you have to add that sum of propulsive drags to the airframe drag to get the corresponding proper drag for "net jet" thrust-drag accounting (not very popular outside the propulsion community).
Thanks for the detailed response. That's actually a little too much detail for what I need. I read your post on ramjet boosters:
Sunday, August 22, 2010
Two Ramjet Aircraft Booster Studies
http://exrocketman.blogspot.com/2010/08 … e-boe.html
I noted that you were able to get better payload with more shallow launch angle but it created a problem for retrieving the first stage booster, since it went so far downrange. If I'm reading it correctly you were able to double the payload mass with the shallow angle, presumably using aerodynamic lift.
What I'm trying to determine if I can increase my payload just going to the range turbojets can get to, ca. Mach 3+. I intend to use the jets to get to medium altitude for a turbojet, ca. 15,000 m. But I need to get to a good angle as well as reaching its max speed. Another problem is that I don't know if it can get to max speed while climbing.
I looked at the case of the SR-71 and the XB-70 Valkyrie. These had more thrust than I wanted but that added weight because jet engines are so heavy. In any case I noted the climbing rate. From that it seemed doable, considering the high effective Isp, that you could reduce propellant mass that way. The problem is this is for a SSTO application and I can't afford the weight. What I wanted was the engines to put out in the range of 1/7th the vehicle weight to reduce the jet engine mass. What I don't know is how will that effect the climb rate, and will it even be able to reach supersonic now.
Note that an advantage of the SSTO is that you can get the better payload by flying a shallow angle and not have to worry about recovering the booster stage.

Bob Clark

I discussed previously on NewMars the reasons why I think it should be possible to do a partially airbreathing SSTO with current jet engines in this post copied below from before the server crash.

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

...Looking at the numbers though I'm convinced now you can even make a single stage to orbit vehicle with a combined ramjet/rocket engine, and without having to use scramjets.
The idea is to combine the turbo-ramjet/rocket into a single engine. This is what Skylon wants to do with their Sabre engine. But the Sabre will use hypersonic airbreathing propulsion up to Mach 6.5 before the rockets take over. This will require complicated air-cooling methods using heat exchangers with flowing liquid hydrogen for the Skylon.

However, just being able to get to say the Mach 3.2 reached by the SR-71 would take a significant amount off the delta-V required for orbit. Of course if the ramjet could get to Mach 5 that would be even better but key this would be doable with the existing engines of the SR-71. Note too the engines of the XB-70 Valkyrie bomber could operate at Mach 3 and as far as I know they didn't have ramjet operation mode. So it might not even be necessary for the engines to have a ramjet mode, turbojet might be sufficient.

The problem with using jets for the early part of the flight of an SSTO has been they are so heavy for the thrust they produce, generally in the T/W range of around 5 to 10. While rocket engines might have a T/W ratio in the range of 50 to 100. But a key point is the jet engine will be operating during the aerodynamic lift portion of the flight where the L/D ratio of perhaps 7. The XB-70 for instance had a L/D of about 7 during cruise at Mach 3. So if we take the T/W of the jet engine to be say 7 and the L/D to be 7, then the thrust to lift-off weight ratio might be about 50 to 1 comparable to that of rockets.

BTW, it is surprising there has been so little research on this type of combination with the jet and rocket combined into one. You hear alot about turbine-based-combined-cycle (TBCC) where it combines turbo- and scram-jets and rocket-based-combined-cycle (RBCC) , where the exhaust from a rocket is used to provide the compression for a ramjet. But not this type of combined turbojet/rocket engine. It doesn't seem to have an accepted name for example. It would not seem to be too complicated. You just use the same combustion chamber for rocket as for the jet. Probably also you would want to close off the inlets when you switch to rocket mode.

For the calculation the delta-V and propellant load would be feasible, note that for a dense propellant SSTO might require as much as 300 m/s lower delta-V than a hydrogen fueled SSTO, in the range of about 8,900 m/s, so I'll use kerosene as the fuel. Hydrogen might have an advantage though in being light-weight if what you wanted was horizontal launch. Say you were able to get to Mach 3+ with the jets, 1,000 m/s. The delta-V to supplied by the rocket-mode is then 7,900 m/s. But note also you can get to high altitude say to 25,000 m. This might subtract another 300 m/s from the required rocket-mode delta-V, so now to 7,600 m/s.

A bigger advantage than this of the altitude is the fact that you get the full vacuum Isp during rocket-mode, call it an exhaust velocity of 3,600 m/s for kerosene rockets. Note this results in a mass-ratio for the rocket mode portion of e^(7,600/3,600) = 8.3, less than half that usually cited for a kerosene-fueled all rocket SSTO. Note the fuel required for the jet-powered portion would only be a fraction of the dry mass rather than multiples of it based on the fact the 1,000 m/s jet-powered speed is only a fraction of the 10,000 m/s or so effective exhaust speed of jet engines.

Note this brings the kerosene fuel load to be about that of hydrogen fueled SSTO's, except you still have the high density of kerosene. With modern lightweight materials this should be well doable.

Bob Clark
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Last edited by RGClark (2012-01-26 06:16:29)

Rune · 2012-01-26 11:35:22

RGClark wrote:

I discussed previously on NewMars the reasons why I think it should be possible to do a partially airbreathing SSTO with current jet engines in this post copied below from before the server crash.

I kinda remember that. My main point of contention is that you are trying to marry an air launch with a SSTO and still get SSTO structural fractions. I imagine it would be much, much easier (perhaps even doable) to keep the two components separated, and do a straight air-launch. Kind of like the new concept that has been flying around by Scaled Composites and SpaceX (and a few strange names thrown in for fun). As to why, here are my main disagreements:

RGClark wrote:

BTW, it is surprising there has been so little research on this type of combination with the jet and rocket combined into one. You hear alot about turbine-based-combined-cycle (TBCC) where it combines turbo- and scram-jets and rocket-based-combined-cycle (RBCC) , where the exhaust from a rocket is used to provide the compression for a ramjet. But not this type of combined turbojet/rocket engine. It doesn't seem to have an accepted name for example. It would not seem to be too complicated. You just use the same combustion chamber for rocket as for the jet. Probably also you would want to close off the inlets when you switch to rocket mode.

The reason for that is rockets don't have turbines placed right after the combustion chamber. Which, by the way, is the main limitation in jet engines, the fact that the first stage of the turbine can't melt. Neither do ramjets and scramjets have them, btw, but don't plan on having any turbojet combustion chamber double-duty as a rocket combustion chamber.

RGClark wrote:

The problem with using jets for the early part of the flight of an SSTO has been they are so heavy for the thrust they produce, generally in the T/W range of around 5 to 10. While rocket engines might have a T/W ratio in the range of 50 to 100. But a key point is the jet engine will be operating during the aerodynamic lift portion of the flight where the L/D ratio of perhaps 7. The XB-70 for instance had a L/D of about 7 during cruise at Mach 3. So if we take the T/W of the jet engine to be say 7 and the L/D to be 7, then the thrust to lift-off weight ratio might be about 50 to 1 comparable to that of rockets.

You want to add not only jet engines, but wings, undercarriage and the works, and still get SSTO mass ratios? Airplanes have structural fractions upwards of 30%, and there is a reason for that. No way you can do all that at 10% structural fraction.

RGClark wrote:

Note this brings the kerosene fuel load to be about that of hydrogen fueled SSTO's, except you still have the high density of kerosene. With modern lightweight materials this should be well doable.

You list every advantage that this "M3 airlaunch carried to orbit" gives you, and they are indeed great, but you gloss over the myriad of "inconveniences" that getting that boost entails, which is almost like carrying a SR-71 to orbit with you. Why not letting a lower stage carry the inefficiencies away at staging and get only the benefits in a true air-launched second stage? That would make much more sense than this. Get all the hardware necessary to do all the things you say, and instead of sticking it to the SSTO, build a plane our of it.

Rune. BTW, before you point to skylon's SABRE, that's different. Not the same cycle as a turbojet, and separate combustion chambers for the rocket part going into the same nozzle.

Terraformer · 2012-01-26 15:21:36

Didn't GW do a back-of-the-envelope design for a hypersonic airlaunch system using a ramjet, with the limitation being that you need the GLOW of a 747 to get several tonnes to orbit (I think this is the article)? Something along those lines, anyway. Mind, that's not such a problem if you can replace the upper stage for a (premium) passenger or cargo module, for when you need very quick transport, though whether the market is enough to make it viable I don't know. Also, what about the prospects for sea launch of very large (1-2000 tonnes?) hypersonic aircraft?

Could integrating the Ramjet and Rocket into one engine allow an SSTO? Or could you not get a high enough T/W ton justify it?

Oh, and when's that book coming out?

GW Johnson · 2012-01-27 12:33:25

Skylon's SABRE is a liquid-air cycle engine. They think they can solve the heat transfer problem to liquify the air as it comes in, with the cold hydrogen. Heat transfer is the slowest of the physical processes, which is why liquid air cycle engines were never attempted before. But, I hope they can do it.

Gas turbine engines can be built that have pretty good T/W. Isp looks good on subsonic fanjets, not so good on supersonic-capable low-bypass ratio turbojets, and really bad when you light the afterburner. The combustor can has to run very lean on fuel to hold turbine inlet temperatures under about 1800 F nor normal flights, under 2000 F for temporary high-power operation. Really exotic materials can add about 200 F more to those figures. But you hit those conditions with turbine somewhere between about M 3.2 to 3.6, almost no matter what design you use.

Most of the combined cycle proposals I have seen compromise performance of each component just to build one device instead of two. (Maybe SABRE can get around that, we'll see.) Most of these combined-cycle things end up pretty complex and heavy, as a geometry change is almost invariably involved. The best of these is the ejector ramjet my old friend Joe Bendot did at Marquardt. It adds a big rocket inside the duct of the ramjet.

On takeoff, the rocket thrust induces some airflow through the ramjet to generate static thrust. Isp is closer to rocket than ramjet. Once you've reached the speed at which the inlet functions properly (usually M1.5+), you can fly on ramjet alone at high speed. This will work as an accelerator up to around 60,000 feet altitude, where frontal thrust density drops too low due to low air density.

At that point you have to turn the rocket back on, and use both together at blended Isp, which floats down toward rocket levels as you climb higher into unusably-thin air. Run the rocket alone, on into space, at rocket Isp. SSTO.

But rocket Isp won't be as high as you are used to, because being inside the duct interferes with free plume expansion, especially in near-vacuum. Typically, lowered Isp raises GLOW of an SSTO no matter what kind of propulsion or airframe design you choose.

I actually like the parallel-burn approach better, because neither device is compromised, and you can run any blend of the two without variable geometry. It actually works out no heavier, and averages higher performance than the combined cycle designs. Rocket-ramjet (maybe to M6, and rocket-turbojet (M3.4+/- on turbo) would be two good candidates.

The air turboramjet could beat the M3.4+/- limit in turbine, if 100% air bypass from ahead of the compressor face was used. No one has ever built one, but I think it could be done without too much fuss and bother. If you can shut down the turbine and cut off all its airflow, bypassing straight to the afterburner, you can run the afterburner as a true ramjet. Capability could be as high as M6. M5 anyway. Hard, but do-able.

As for the book, I have several topics roughed out now, but only about 30% or so of them. None are in final form. The science is not too hard to write down, it's the art that's hard. I did some of that the other day, on really how to size ramjet system geometries for several different ramjet systems and weapon/launch applications. Still sweating. Or swearing. (Both, maybe.) I never wrote that stuff down before, I just did it, and noticed I was one of very few around the country able to do it that fast and well. Neither has anyone else ever written sizing procedures down, as near as I can tell.

GW

Last edited by GW Johnson (2012-01-27 12:41:10)

GW Johnson · 2012-01-27 16:54:10

For Bob Clark:

It makes a lot of sense for the ramjet and the rocket to burn a common fuel. My old article that Rune referenced just above does exactly that.

There are several kerosenes: JP-5/Jet-A/Jet-A-1, RP-1, and commercial K-1 kerosene for kerosene heaters and lanterns. These all have almost identical distillation curves and properties. K-1 is the dirtiest.

Clean is very important: fuel metering orifices clog easily. Liquid methane is the very cleanest, though. You have to rework your fuel system to handle a cryogenic fuel, but fundamentally, both ramjet and rocket can burn liquid methane.

Rocket and ramjet might be made to burn diesel, but vaporization is difficult. Both can be made to burn gasoline quite easily. The downside is the flammability and explosion hazards associated with gasoline. If you use gasoline (and both rocket and ramjet will burn it), there is no point to high octane rating in either engine: 30 MON drip gas is fine.

Anything a ramjet can burn, a gas turbine can be made to burn as a drop-in fuel, FAR's and type certificates notwithstanding.

There is an old ramjet fuel known variously as RJ-5 / Shelldyne-H. It is an artificial synthetic, a single substance, not a mixture of compounds like all distillates. It does behave exactly like kerosene in all the subsystems of a kerosene ramjet, it's just denser: sp.gr > 1. (It sinks in water, unlike all the petroleum liquid hydrocarbons.) But, it's expensive. Yet, propellant costs are but a tiny fraction of launch costs.

For first stage use, density x Isp is more important than Isp. That's why kerosene beats hydrogen for first stage fuel use in practical designs. RJ-5 is really good for that.

GW

Terraformer · 2012-01-27 17:13:38

What kind of mass, Isp, and T/W are we looking at for a Ramjet capable of taking a craft from say 0.5 to 2km/s, when burning Methane? I'm thinking of perhaps a Methane/LOX SSTO utilising parallel burning...

Rune · 2012-01-27 19:09:19

GW Johnson wrote:

I actually like the parallel-burn approach better, because neither device is compromised, and you can run any blend of the two without variable geometry. It actually works out no heavier, and averages higher performance than the combined cycle designs. Rocket-ramjet (maybe to M6, and rocket-turbojet (M3.4+/- on turbo) would be two good candidates.
The air turboramjet could beat the M3.4+/- limit in turbine, if 100% air bypass from ahead of the compressor face was used. No one has ever built one, but I think it could be done without too much fuss and bother. If you can shut down the turbine and cut off all its airflow, bypassing straight to the afterburner, you can run the afterburner as a true ramjet. Capability could be as high as M6. M5 anyway. Hard, but do-able.

Yeah, integrating a ramjet into a reaction engine (or, more like, vice-versa) would get around the problem of getting the ramjet supersonic and into operating speed. But "your" method of using the ramjet as a solid booster casing sounds a lot more simple, mechanically. And simpler means less maintenance, and more reliability which, fuel costs aside, eventually means cheaper and safer. Landing issues of a ramjet aside, that is. Definitely the way for expendables.

Rune. ¡Y ánimo con el libro!

Last edited by Rune (2012-01-27 19:09:54)

karov · 2012-01-28 06:56:18

http://en.wikipedia.org/wiki/Philip_Bono

I think this is the truth. Make it without expendable first stage boosters and running on cheaper and easier to handle fuel and oxidizer / say nox + kerosene /.

GW Johnson · 2012-01-28 09:38:22

If you do TSTO with a rocket+ramjet 1st stage, you have to decide fundamentally between vertical-takeoff/ballistic/wingless and horizontal-takeoff/lifting-flight/winged. The decisions on which stage does what with wings are secondary to that first level choice. At least, that's what my attempts at sizing systems and components says.

If you do HTO, I recommend a winged booster airplane carrying a second stage that could be either a small rocket plane or a ballistic rocket. My best approach so far uses a high-speed ramjet design as the fuselage (for the biggest cross-sectional area), with rockets and propellant tankage in the wing fillets and wings. The ramjet has a centerbody external compression spike in the nose, and the booster pilot rides there.

It takes off on rocket only, climbing and accelerating to M1.6-ish, then shifts to ramjet propulsion only. It climbs at M1.6-ish in ramjet, gradually pulling over level and accelerating, at about 60,000 feet. There it accelerates in ramjet to about M5.5 to 6, depending more on the drag of the configuration than ramjet thrust. Light the rockets too, pull up sharply, and stage at a path angle for ballistic flight of the second stage. The booster then cuts rocket and throttles back ramjet, does a decelerating pull-down 180-turn, and cruises back to base at M1.6-ish in ramjet. At base, it goes subsonic powerless for a glide landing, retaining a small kitty of rocket propellant for go-around or divert.

The 3 most important HTO staging variables, in order of importance, are velocity, path angle, and altitude. This kind of approach could get you close to M6 (about 1.8-1.9 km/sec) at staging, path angles 30-60 degrees, and probably no more than 60,000 feet (18 km) altitude.

The ramjet has an external- or mixed-compression inlet with a min takeover speed around M1.5 to 1.6, a slightly convergent-divergent nozzle with a throat area pretty near 65% of the combustor duct area, a sudden dump flameholder approach, probable perforated air-cooling sleeve, and Inconel-X skins like the X-15. (Shock-impingement heating damage becomes intolerable above M6 with all known materials, so I see no point to trying a scramjet for this.)

Rocket T/W at TO need not exceed the [EDIT: 0.x instead of the original 1.x, sorry, my mistake ] 0.3-0.5 range of any other airplane. The inert fraction of the booster will look more like an airplane because that is what it really is: 30-40%. If you try to push the staging altitude too high, the ramjet is unable to accelerate fast enough, and cannot fly the return range. Staging altitude is a tradeoff driven by that range. I don't have an exact figure for that yet.

The velocity requirement for the second stage is around 20,000 feet/sec (6.1 km/sec), which is high, but do-able with the one-shot technologies that we already have. Its nozzle pretty much always operates in near-vacuum conditions.

I know a lot less about the VTO case, except that it is more of a ramjet-assisted rocket vehicle. The idea is to shoulder some of the dead-weight loads onto as much higher-Isp / lower frontal-thrust ramjet as you can afford. I don't think the ramjet will exceed about 20% of the first stage thrust, maybe not even 10%.

Vehicles like this typically leave the sensible air (60,000 feet, 18 km) fairly slow: M 2-ish. Thus, a low-speed ramjet design is better: normal-shock / pitot inlet, convergent-only nozzle capable of unchoked operation, and usable thrust at high subsonic flight speeds (say, M 0.7+). This thing most likely would be a set of strap-on pods staged off around 60,000 feet, and perhaps flown back to launch site for recovery like a big model airplane, but a glider, of course. Nose wheel and tail skids, like X-15. Some sort of extendable wing? Scissor swing wing?

To solve the dead-weight problem between launch and ramjet ignition, I'd package a missile-type integral solid booster inside the ramjet case, which is lined with simple ablatives. Plain steel, no external heat protection needed. I'd make the ramjet pod modular, for easy re-insulation and reload by very few folks with very little effort. An inlet module, a concentric fuel tank and air duct module, a combustor/booster module, and a nozzle module.

Mix and match and stack, then hang these on the core vehicle as strap-ons. It'll shed a port cover and a booster nozzle at transition to ramjet. The core rocket is a typical TSTO with staging around 10,000 feet/sec (3 km/sec) well outside the sensible atmosphere. You just get to add a few % more payload than it would normally carry.

Those are my best guesses right now.

GW

Last edited by GW Johnson (2012-01-31 08:51:54)

Terraformer · 2012-01-28 10:40:48

What about shifting more of the velocity on to the first stage using the rocket motors, if you're using parallel burning? For example, closing the inlet at M6 and using rockets to push the entire thing to M10 before staging - or would that reduce the payload ability? Could we do an inert fraction of say 20% while still being structurally rugged enough for easy reusability, given the materials advances, and get maybe 2% total payload fraction?

How hard would building one for suborbital space tourism be, if possible - can we translate M6@60,000ft into an altitude gain of 350,000ft? That would require pulling up at 45 degrees, to result in a vertical velocity sufficient for this... depending on the demand for high speed transport, we may not need to do this to pay for the development of the first stage aircraft.

Rune · 2012-01-28 14:42:41

I am intrigued enough by the prospects to explore that vertical-launched ramjet variant a bit more. I especially like it, because a ramjet has an easier time flying back to the launch site under its own power, however it lands (parafoil gliding on skids?), than a rocket does. And I don't like ocean recovery, or having to transport stages across continents to reuse them.

First, I would like to ask how hard it would be to bend further the standard gravity turn of vertical launch trajectories to keep the ramjets operational for their full, up-to-M6, envelope inside atmosphere, then stage and pull up in rocket power. More aerodynamic heating, for sure, but if we are talking reusables, like we are, then they will have to endure worst on their way down. If we can get the staging done at those speeds and altitudes, then it should be no problem to solve the rest of the trip to orbit with dense fuel engines and reasonable enough mass ratios (7-8 with kerolox and a vacuum nozzle, for the ~7km/s remaining to orbit) to consider a single stage that can be reused. You know, since we make good expendable ones with MR 30, and just tens of flights out of the same airframe would be already a big revolution. Such a "second stage" would also be a single-stage moon and mars lander, by the way. Just saying.

Then, I would challenge the percentage of thrust assigned to the ramjets. In fact, I would challenge that doing a parallel burn with the rockets is a good idea in the first place. First off, we all agree the second stage(s) takes the lion's share of the delta-v to orbit, and makes most of its work outside the atmosphere. Therefore, if we don't want to get into the trouble of an altitude-compensating nozzle, the engine should only be used outside the atmosphere (or at least, the dense portion of it). Not only that, but at the speeds the ramjet is going to take over, the body of the rocket could give a significant amount of aerodynamic lift to counter the gravity losses of the slower accelerating ramjet at a very reasonable attack angle. That would pretty much solve the low thrust and frontal thrust issues inherent with ramjets. Also multiply the drag losses, but you know, nothing is free and right now they are in the order of hundreds of m/s. We are going to have a structure though enough to take it, if we plan on reusing the whole thing, after all.

And, the first stage is the one we are SURE of recovering. It's the biggest, most complex, most expensive part of any launch vehicle (ok, maybe the shuttle was more complex than it's launch vehicle... but only if you count the SSME's as part of the shuttle, not the launch vehicle). We should be trusting it to do as much as it can, so we can get the most benefit out of reusing it, hopefully, with airplane-like procedures and lifecycles. That way, we can build only a few, even if they are big, and still get a huge benefit in terms of final cost to orbit.

And takeoff thrust, thanks to your "solid engine inside the combustion chamber" method, GW, is certainly not the problem. Maybe keeping the G's low enough for humans is, instead? Handling the loaded booster prior to liftoff can be a problem (solids are HEAVY), but again, modularity and parallel staging is clearly the route around that as you said.

I ran into a problem here, however. The second stage also has to be reused, and reused fast and easy, so it has to land properly, not splash down in the ass-end of nowhere. And it's going to be strapped for mass ratio, so not a lot of room to add fancy landing systems like wings, undercarriage and that short of stuff. Those kinds of things belong in the ramjet booster(s), which is the airplane-like part of this vehicle. Increased structural weight hurts less in the lower stage, and all of that. A rocket is NOT an airplane.

Seeing as how it is big as hell (MR 7-8, remember? Even if it is a dense fuel...), maybe you could land it gliding as a lifting body with only the undercarriage and control surfaces required. Maybe a parafoil for good measure and reasonable approach speed. Throw is a cargo bay, probably, so you don't have the aerodynamics of hell to contend with, although carrying stuff on shrouds has worked well until now. But the really Heinlenian way to go about landing would be to turn back on the main propulsion system and land majestically on a plume of rocket awesomeness and hydraulic pistons. Problem is, I insisted on having the engines designed for vacuum. Thrust is not an issue here, since the stage is mostly empty, but the expansion ratio is going to mess up the engine seriously. Like backpressure choking the engine and blowing out the combustion chamber, right? Thing is, again you can design around it (me likes aerospikes, promise! They go well with VTVL) but you are going to take a mass/performance hit somewhere, for sure. And I can't help the nagging feeling that it's the thing that makes this architecture not the best one. Of course, I'm expecting the lot of you to polish it until it is!

Rune. So please go ahead. And have fun while you are at it!

Last edited by Rune (2012-01-28 14:46:28)

Terraformer · 2012-01-28 16:23:45

You could decide to not reuse the upper stage, I guess - if you can make it cheap enough, it might not make sense to do so for a 1st generation TSTO system. See if you can launch humans in some kind of capsule, perhaps with it's own, 2km/s delta-V if you want to go for a 2.5 stage to orbit system - with the first stage aircraft, you should be able to make the middle stage splash down in the ocean.

Or even land the upper, orbital, 7km/s delta-V stage in the ocean - horizontally. Make it a sea spaceplane. The problem is, of course, that your engines are still going to get splashed...

GW Johnson · 2012-01-28 21:51:23

Hi Rune:

The integral booster is a very old technology, much like the rest of ramjet. The first flying integral booster ramjet I know of is the SA-6 "Gainful" SAM that the Russians built, starting late 60's. It's an easy thing to do technologically, and very effective. Missiles like that are built cheap, too. Low logistical tail. Surprise, surprise.

I did the sustainer ramjet mechanical engineering exploitation on the -6 here in the US, back in the late 70's. I also got to work on ASALM-PTV and the original US VFDR in the 80's and 90's. All had integral boosters. I saw but did not work on LTV's ALVRJ integral booster ramjet in '74, when it first rolled out for test. I was working on LTV's Scout satellite launcher at that time.

The two kinds of ramjet overlap but are not interchangeable. The pitot inlet type is useful from about M0.7 to about M2 or 2.5. The supersonic inlet type is useful from about M1.6 to about M5 or possibly 6. ASALM accidentally went to M6 about 1980 ("accidentally"? thereby hangs a tale). The supersonic type of ramjet has the higher Isp and frontal thrust density. But its inlet doesn't work at all below about M1.5.

I'd rather increase the launch acceleration levels of a VTO rocket to reach about M5 at about 60,000 feet, than try to bend the trajectory to stay in the air to higher speeds. Trajectory-bending requires lift forces, meaning wings, and that puts you right back to the HTO TSTO airplane launcher. Use V^2 = 2as to estimate average gees for M5 at 60,000 feet, and you find it to be 6.5 gees. Peak might be twice that at 13 gees.

Apollo pulled 11 gees coming back from the moon. Quite survivable by humans. 45+ gees is not, as Paul Stapp proved at Holloman AFB on the rocket sled in the late 50's. There's a lot of roller coasters out there that pull 5 gees as brief transients in loops. They just don't tell anybody that's what they're pulling. Easy to figure from speed and radius, though.

Short transients to 5 gees under a few seconds are survivable by most members of the general public, no matter how unfit. Someone reasonably fit can pull 5 gees for several minutes without a gee suit. They did it all the time in WW2 fighters. Time to 10,000 fps (M10) at a constant 6.5 gees is about 50 seconds. Reasonably short.

GW

GW Johnson · 2012-01-28 22:07:58

Terraformer:

As near as I can tell, reusing upper stages would be very difficult. It would be easier and more practical to orbit the stage fully, and then de-orbit later to a place of your choosing.

I think an ablative heat shield on the front end, plus some ablative inflatables around the rockets engine(s) at the rear, would do nicely. The inflatables also protect the engines during ocean impact. It'll need a series of chutes deployed out the rear to slow it to under 100 mph at impact, and it'll need to be tougher than an old boot to survive 100-gee-class impact forces.

You just don't do that with any known materials at 5-8% inert fractions. But you might get it done at 15-20% inert fractions. To keep the same payload fraction overall, you'll need three stages instead of two.

GW

Rune · 2012-01-29 18:14:37

GW Johnson wrote:

The integral booster is a very old technology, much like the rest of ramjet. The first flying integral booster ramjet I know of is the SA-6 "Gainful" SAM that the Russians built, starting late 60's. It's an easy thing to do technologically, and very effective. Missiles like that are built cheap, too. Low logistical tail. Surprise, surprise.
I did the sustainer ramjet mechanical engineering exploitation on the -6 here in the US, back in the late 70's. I also got to work on ASALM-PTV and the original US VFDR in the 80's and 90's. All had integral boosters. I saw but did not work on LTV's ALVRJ integral booster ramjet in '74, when it first rolled out for test. I was working on LTV's Scout satellite launcher at that time.

Yeah, I knew you didn't invent the concept or anything, but you did introduce it to me, way back before the big crash. Sorry, but in my head, the credit for me knowing about it is yours.

GW Johnson wrote:

I'd rather increase the launch acceleration levels of a VTO rocket to reach about M5 at about 60,000 feet, than try to bend the trajectory to stay in the air to higher speeds. Trajectory-bending requires lift forces, meaning wings, and that puts you right back to the HTO TSTO airplane launcher. Use V^2 = 2as to estimate average gees for M5 at 60,000 feet, and you find it to be 6.5 gees. Peak might be twice that at 13 gees.
Apollo pulled 11 gees coming back from the moon. Quite survivable by humans. 45+ gees is not, as Paul Stapp proved at Holloman AFB on the rocket sled in the late 50's. There's a lot of roller coasters out there that pull 5 gees as brief transients in loops. They just don't tell anybody that's what they're pulling. Easy to figure from speed and radius, though.
Short transients to 5 gees under a few seconds are survivable by most members of the general public, no matter how unfit. Someone reasonably fit can pull 5 gees for several minutes without a gee suit. They did it all the time in WW2 fighters. Time to 10,000 fps (M10) at a constant 6.5 gees is about 50 seconds. Reasonably short.

Well, if I could get a ramjet to pull the second stage at 6.5G, then I would take it too, for sure. Gift-wrapped if you please . But seems to me (without any practical knowledge in the subject) that a ramjet isn't going to achieve the T/W required to do so, unless it's a very big first stage compared to the second one. They are basically just supersonic sustainers in missiles, right? And you don't really need "wings", just enough control surface to have control authority from the high hypersonic through the low supersonic/high subsonic. Which, in something that has to return to the vicinity of the launch site under ramjet power, is not a bad idea anyhow. The body is probably going to have enough surface (counting the second stage also) to provide enough lift to make do with much less thrust after the solid boosters run out, if you can stay using it for longer (even T/W<1?). I know your generation didn't like to count on the lift generated by fuselages, but this days a computer will crunch the numbers without much problems (...or accuracy. But it'll get you there). If you are forced to add real wings, then it's just a big airplane (rocket boosted ramjet plane, but plane), and therefore HTHL makes more sense.

Rune. I'll admit VTHL lifting bodies are intriguing also. Not much on their structural fractions though, compared to a capsule, for example.

GW Johnson · 2012-01-30 10:55:45

Hi Rune:

Yep, compared to rocket, ramjet has a low frontal thrust density, due in part to the lower chamber pressure, and in part due to ingested airflow ram drag. Unless the core rocket accelerates like a bandit, the ramjet will be incapable of accelerating like that, especially vertically.

Besides, most existing launch vehicles accelerate off the pad very gradually. T/W just barely > 1. I'd add strapons with integral boosters to help it do the same thing at higher payload weight, and then save on a little of rocket propellant by shouldering a piece of the load between about M0.7 low and M2 about 60,000 feet.

That low down, the trajectory is still essentially vertical, so it would be easy to recover the strap-on ramjet pods at the launch site. I kind of like the scissor-wing idea the Russians put on their "Baikal" proposed strap-on booster idea. Add a nosewheel and two tail skids, and land on a runway.

But, you're right. Most missiles need no wings at all to pull lots of lifting gees at M3+.

GW

Rune · 2012-02-01 14:22:48

Rune wrote:

GW Johnson wrote:
I know nothing about any "super Draco" thrusters for Dragon. The ones they have are supposed to be the launch escape system for any manned Dragon, so capsule T/W is substantially larger than 1. For the Mars mission paper I gave at last August's convention in Dallas, I re-engineered some data from the Spacex website for Dragon as they had it posted last summer. I was showing 0.9 km/sec delta-vee capability with 6 suited astronauts on board. If you put more propellants in the unpressurized module, and connected that to the Draco system, it adds around 1.4 km/sec more to the capsule's delta-vee capability. I looked at that to get 2+ km/sec total delta-vee for Dragon as an emergency escape crew return vehicle from a very high-speed Mars return in an emergency, somewhere well above 15 km/sec. Maybe as much as 25 km/sec.
Well, they are the escape system under design, I believe. The current Dracos are tiny 400N thrusters optimized for vacuum for reaction control of the capsule and the second stage, so if they want to give the whole capsule T/W over 3 (and probably closer to six, to ensure escape) they need a new engine. I can't find the reference, but Mueller (the propulsion VP) has dubbed them "super-Dracos" and said they are under development. I think it was the conference where they presented the Merlin 1D?

What do you know? Just a few days later, there is a press release dedicated to "SuperDracos". I comment it here.

Rune. Not a peep on T/W, though. But they are deep-throttleable.

GW Johnson · 2012-02-02 15:47:02

Rune:

I looked at Spacex's web page yesterday. You are right, the Super-Draco is a much more powerful MMH-NTO thruster system. What I saw indicated 8 units, angled though they are, sum up to 120,000 pounds of axial thrust.

It's a tad unclear, but I'd hazard a guess they will keep the little Draco's for attitude control, and use the big Dracos as a launch escape system, a delta-vee maneuver source, and as a landing system.

As a landing system, I'd bet it's combined chute and final-seconds thrusters.

GW

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#172 2012-01-29 18:14:37

Re: Reusable Rockets to Orbit

#173 2012-01-30 10:55:45

Re: Reusable Rockets to Orbit

#174 2012-02-01 14:22:48

Re: Reusable Rockets to Orbit

#175 2012-02-02 15:47:02

Re: Reusable Rockets to Orbit

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