Reusable Rockets to Orbit

RGClark · 2012-08-29 10:26:02

I was discussing two different topics with those last two posts of mine. The first was about the question of producing a lightweight structure when it only had to withstand a 300 psf, 2psi, dynamic pressure as the shuttle does on re-entry.
The second had to do with the likelihood of Elon Musk succeeding at cutting the cost to orbit with reusability when it would require reusable rocket engines at low maintenance costs.

Bob Clark

GW Johnson · 2012-10-16 10:23:15

News story seen today on NBC news "space" topic: Krunichev in Russia is pushing ahead with a reusable flyback booster strap-on. This is for vertical launch rockets. The booster strap-on is a liquid rocket (LOX and kerosene or methane) unit with tail fins and tricycle landing gear. One version has a swivel straight wing. It stages off about 30 km up at around M7, and falls back to winged lifting flight at about 12 km (if memory serves). Then it cruises back on turbojet propulsion to the launch site, and lands horizontally as an airplane.

This is a concept seen at airshows as a mockup for some years now. It was also a topic of conversation in multiple threads here before the last server crash. One of the variants of this that I have explored is an integral rocket-ramjet strapon that stages around 70,000-80,000 feet and M2.5-toM3, based on simple pitot-inlet ramjet technology. Such a thing would integrate well with current acceleration practices for vertical launch rockets.

The article said NASA had looked at similar ideas, using the rocket engines to cruise back, instead of a turbojet package.

The fallback does require proper treatment of nose shape and construction for hypersonic but well-suborbital re-entry. And perhaps several other surfaces as well, but that was not mentioned.

If sufficiently simplified so as to require a small logistical tail on the order of a battlefield weapon, this could lead to significant launch cost reductions. We already know small logistical tail has a huge effect. Here is a chance to see if lower manufacturing cost could provide similar benefit.

GW

RGClark · 2012-12-27 01:43:42

RGClark wrote:

Hop wrote:
...
It seems to me one of the problems is achieving an FMR of 16:1 and having a spacecraft strong and temperature resistant enough to endure re-entry.
Watching Musk's Grasshopper video it looks like he hopes to use reaction mass to shed re-entry velocity in addition to aerobraking. If he hopes to achieve some re-entry delta V with propellant, this makes his FMR even more challenging. Get the FMR too high and you have a very tenuous, fragile vehicle even less able to endure re-entry. I'm not giving Musk's TSTO RLV even odds.
However, given propellant in orbit, I believe it's quite doable to decelerate the upper stage and land it intact on the launch pad. Thus I believe lunar supplied propellant depots would enable TSTO RLVs.
And not just TSTO's, and not just to orbit. But in fact SSTO's all the way to the Moon and back. It is a well known fact that if you have SSTO's, then with orbital refueling you can travel all the way to the Moon, land, lift off, and travel all the way back to Earth on that one single refueling. Another one of the many advantages of SSTO's. Note that this is not true for TSTO's whose second, orbital stage might get a delta v of, say, 6,000 m/s, insufficient for the round-trip to the Moon even with refueling in LEO.
I've been arguing that SSTO's are actually easy because how to achieve them is perfectly obvious: use the most weight optimized stages and most Isp efficient engines at the same time, i.e., optimize both components of the rocket equation. But I've recently found it's even easier than that! It turns out you don't even need the engines to be of particularly high efficiency.
SpaceX is moving rapidly towards testing its Grasshopper scaled-down version of a reusable Falcon 9 first stage:
Reusable rocket prototype almost ready for first liftoff.
BY STEPHEN CLARK
SPACEFLIGHT NOW
Posted: July 9, 2012
http://www.spaceflightnow.com/news/n1207/10grasshopper/
SpaceX deserves kudos for achieving a highly weight optimized Falcon 9 first stage at a 20 to 1 mass ratio. However, the Merlin 1C engine has an Isp no better than the engines we had in the early sixties at 304 s, and the Merlin 1D is only slightly better on the Isp scale at 310 s. This is well below the highest efficiency kerosene engines (Russian) we have now whose Isp's are in the 330's. So I thought that closed the door on the Falcon 9 first stage being SSTO.
However, I was surprised when I did the calculation that because of the Merlin 1D's lower weight the Falcon 9 first stage could indeed be SSTO. I'll use GW Johnson's estimates for the Falcon 9 specs here:
WEDNESDAY, DECEMBER 14, 2011
Reusability in Launch Rockets.
http://exrocketman.blogspot.com/2011/12 … ckets.html
The first stage propellant load is given as 553,000 lbs, 250,000 kg, and the dry weight as 30,000 lbs, 13,600 kg. The Merlin 1C mass hasn't been released, but I'll estimate it as 650 kg, from its reported thrust and thrust/weight ratio. The Merlin 1D mass has been estimated to be 450 kg. Then on replacing the 1C with the 1D we save 9*200 = 1,800kg from the dry weight to bring it to 11,800 kg.
The required delta v to orbit is frequently estimated as 30,000 feet per second for kerosene-fueled vehicles, 9,144 m/s. When calculating the delta v your rocket can achieve, you can just use your engines vacuum Isp since the loss of Isp at sea level is taken into account in the 30,000 fps number. Then this version of the Falcon 9 first stage could lift 1,200 kg to orbit:
310*9.81ln(1 + 250/(11.8 + 1.2)) = 9,145 m/s.
Then the Falcon 9 first stage could serve as a proof of principle SSTO on the switch to the Merlin 1D.

Bob Clark

Experimental Private Rocket Makes Highest Test Hop Yet.
by Miriam Kramer, SPACE.com Staff WriterDate: 26 December 2012 Time:
11:04 AM ET
"In the latest test at SpaceX's proving grounds in MacGregor, Texas,
the Grasshopper rocket flew for 29 seconds and reached a height of
more than 130 feet (40 meters). A video of the Grasshopper test flight
shows the rocket soaring up into the Texas sky, then smoothly
descending to land on four spindly legs."
http://www.space.com/19039-spacex-priva … -test.html

With reduced weight of the Merlin 1D engine while at increased efficiency, the Falcon 9 v1.1 first stage will have SSTO capability. Then ironically Elon is emulating the original purpose of the DC-X program in testing the Grasshopper VTVL stage, without realizing it.

Bob Clark

JoshNH4H · 2012-12-27 09:27:18

Just wondering, do you have a source on your assertion that the F9 first stage will SSTO capacity?

RGClark · 2012-12-29 13:00:06

JoshNH4H wrote:

Just wondering, do you have a source on your assertion that the F9 first stage will SSTO capacity?

It's theoretical since SpaceX hasn't released yet the specifications for the F9 v1.1. The justification was given in post #188 in this thread.

Bob Clark

Last edited by RGClark (2012-12-29 13:03:17)

JoshNH4H · 2012-12-30 07:56:57

Ah yes, thank you.

RGClark · 2013-07-22 02:00:08

GW Johnson wrote:

News story seen today on NBC news "space" topic: Krunichev in Russia is pushing ahead with a reusable flyback booster strap-on. This is for vertical launch rockets. The booster strap-on is a liquid rocket (LOX and kerosene or methane) unit with tail fins and tricycle landing gear. One version has a swivel straight wing. It stages off about 30 km up at around M7, and falls back to winged lifting flight at about 12 km (if memory serves). Then it cruises back on turbojet propulsion to the launch site, and lands horizontally as an airplane.
This is a concept seen at airshows as a mockup for some years now. It was also a topic of conversation in multiple threads here before the last server crash. One of the variants of this that I have explored is an integral rocket-ramjet strapon that stages around 70,000-80,000 feet and M2.5-toM3, based on simple pitot-inlet ramjet technology. Such a thing would integrate well with current acceleration practices for vertical launch rockets.
The article said NASA had looked at similar ideas, using the rocket engines to cruise back, instead of a turbojet package.
...
GW

Just saw this mentioned on another forum:

Reusable Ram Booster Launch Design Emphasizes Use of Existing Components to Achieve Space Transport for Satellites and Spacecraft.
http://www.nasa.gov/offices/ipp/centers … oster.html

It proposes separate turbojet and ramjet stages for an orbital launcher, with the final stage to orbit powered by rockets. However, I'm wondering if it might be better for simplicity and cost to combine the turbojet and ramjet stages into a single stage whose engines can work in both turbojet and ramjet modes, a la the SR-71 Blackbird.
I remember reading the primary impediment against the SR-71 reaching Mach 4+ in its ramjet mode, was cooling the engine. Then since this is for an orbital vehicle, we might borrow a concept from rockets even for this air-breathing stage of using regenerative cooling.

Bob Clark

GW Johnson · 2013-07-22 09:37:39

Hi Bob:

This is interesting stuff. 3-stage design, taking off on turbojet. Looks to be VTO, but must arc over and fly more of a HTO flight profile. Another patent for NASA, I guess. But, there's a fundamental problem with Mach 4 at 140-150 thousand feet with an airbreather of any kind.

This is not a lot different from some stuff I did and posted a couple of years ago over at "exrocketman". In the best form I found, I went two-stage for HTO, using rockets to take off and reach ram takeover speed, which for a supersonic inlet design is around Mach 1.5 to 1.6 min. You climb at near-takeover speed on ramjet, then pull-over and accelerate to high speed (Mach 5 to 6, actually). Then you must parallel-burn both rocket and ramjet to pull up sharply to release the second stage.

The first stage cruises back to base on ramjet at near-takeover speed, and glides to landing with reserve rocket propellant on board for go-around capability. The second stage is a straight rocket pod or plane. There are very severe frontal thrust density problems with ramjet above about 60 or 70 thousand feet. You can't climb or accelerate on airbreather alone, which is why combined-cycle airbreathing engines make little sense to me for launch applications.

For VTO, the ramjet assist is a strap-on, on a rocket core, intended to add a little thrust to the mix, just at much higher Isp. But, it's a simple pitot inlet, since most VTO designs leave the sensible atmosphere at about Mach 2-ish at around 80 thousand feet. Pitot inlets have subsonic ram takeover speeds. It's going to be a small effect, if it is worth it all. I'm not yet sure whether that scenario is really worthwhile. But I do know roughly what the fly-back strap-on pod ought to look like.

The problem is low frontal thrust density in the thin air above 60 thousand feet. All airbreathers are afflicted by that. Not even flying super fast with scramjet overcomes it. You either stay low and lose all your impulse advantage in drag, or you have to burn rocket and airbreather in parallel to achieve enough frontal thrust to climb/accelerate in the thin air. I just don't see any way around that dilemma. I see the potential for a lot of "gravy train" R&D programs, but I don't see much potential for anything we might actually fly.

The UK Skylon engine faces the same problem. They're pretty much done the airbreather by 80 thousand feet, gone to rocket only mode. For them, that's about Mach 4. Surprise, surprise!

GW

Last edited by GW Johnson (2013-07-22 09:41:06)

RGClark · 2013-07-22 23:53:41

GW Johnson wrote:

?..
The problem is low frontal thrust density in the thin air above 60 thousand feet. All airbreathers are afflicted by that. Not even flying super fast with scramjet overcomes it. You either stay low and lose all your impulse advantage in drag, or you have to burn rocket and airbreather in parallel to achieve enough frontal thrust to climb/accelerate in the thin air. I just don't see any way around that dilemma. I see the potential for a lot of "gravy train" R&D programs, but I don't see much potential for anything we might actually fly.
The UK Skylon engine faces the same problem. They're pretty much done the airbreather by 80 thousand feet, gone to rocket only mode. For them, that's about Mach 4. Surprise, surprise!
GW

Air launch is not perfect. It won't remove the need for a final rocket stage. But the question is does it improve your payload capability? Several studies suggest that it can:

Air Launching Earth-to-Orbit Vehicles: Delta V gains from Launch Conditions and Vehicle Aerodynamics.
Nesrin Sarigul-Klijn University of California, Davis, CA, UNITED STATES; Chris Noel University of California, Davis, CA, UNITED STATES; Marti Sarigul-Klijn University of California, Davis, CA, UNITED STATES
AIAA-2004-872
42nd AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, Jan. 5-8, 2004
http://pdf.aiaa.org/preview/CDReadyMASM … 04_872.pdf [first page only]

The conclusions are summarized in this online lecture:

A.4.2.1 Launch Method Analysis (Air Launch).

For a launch from a carrier aircraft, the aircraft speed will directly reduce the Δv required to attain LEO. However, the majority of the Δv benefit from an air launch results
from the angle of attack of the vehicle during the release of the rocket. An
ideal angle is somewhere of the order of 25° to 30°.
A study by Klijn et al. concluded that at an altitude of 15250m, a rocket launch with the
carrier vehicle having a zero launch velocity at an angle of attack of 0° to
the horizontal experienced a Δv benefit of approximately 600 m/s while a launch
at a velocity of 340m/s at the same altitude and angle of attack resulted in a
Δv benefit of approximately 900m/s. The zero launch velocity situations can
be used to represent the launch from a balloon as it has no horizontal velocity.
Furthermore, by increasing the angle of attack of the carrier vehicle to
30° and launching at 340m/s, a Δv gain of approximately 1100m/s
was obtained. Increasing the launch velocity to 681m/s and 1021m/s produced a
Δv gain of 1600m/s and 2000m/s respectively.
From this comparison, it can be seen that in terms of the Δv gain, an airlaunch is
superior to a ground launch. As the size of the vehicle decreases, this superiority
will have a larger effect due to the increased effective drag on the vehicle.

https://engineering.purdue.edu/AAE/Acad … aunch).doc

A speed of 340 m/s is a little more than Mach 1, while subsonic transport aircraft typically cruise slightly below Mach 1. So the delta-V saving could still be in the range of 1,000 m/s with air launch, a significant savings by the rocket equation.
And this study found by using a supersonic carrier aircraft you could double the payload of the Falcon 1:

Conceptual Design of a Supersonic Air-launch System.
43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit
8 - 11 July 2007, Cincinnati, OH
http://www.ae.illinois.edu/m-selig/pubs … Launch.pdf

Bob Clark

Last edited by RGClark (2013-07-22 23:59:02)

GW Johnson · 2013-07-23 10:35:18

A couple of years ago, I visited XCOR in Mojave, CA. They are looking at a future orbital craft. The idea is 2-stage HTO. First stage could be a rocket, but is a lot smaller if an airbreather. They wanted to look at ramjet, which is why they were interested in me. I am a full-capability ramjet expert.

I don't have a copy of their internal report, but I do have the verbal ranking of the important issues at staging. Most important is velocity, the faster the better. Second is path angle, as near to 40 degrees as possible (which relieves the second stage of the hardware required to pull up by lift or thrust, it just flies a gravity-drag ballistic turn). Third is altitude at staging: the higher the better.

Of the airbreathers, ramjet offers the highest staging velocity with something that can be reliably applied right now. A clean, low-drag vehicle design can reach Mach 5 fairly easily, even Mach 6. ASALM-PTV accidentally reached Mach 6 back in 1980, and it was a very clean, low-drag wingless dart shape.

None of the airbreathers has much frontal thrust density above about 60,000 feet, simply because densities and ambient pressures are so low. Compression ratio is limited to what the inlet can do at that Mach. So thrust depends upon incoming pressure, but weight does not. There is no way around these physics. That is why to pull up sharply at staging in the thin air requires more thrust than the airbreather can deliver, or else the vehicle decelerates sharply as you pull up. The most practical way to achieve pull-up thrust is to fire up some rockets in addition to your airbreather; i.e.; parallel burn, not combined-cycle.

Since altitude is the weakest of three effects at staging, one can stage at 60,000 feet instead of 100+ thousand feet, and get way-to-hell-and-gone better thrust results from the airbreather. This means both velocity and pull-up angle can be fully achievable. Plus, at only 60,000 feet, the time and range to accelerate (on the airbreather alone) to max velocity are a whole lot shorter than those at the higher altitudes in the too-thin air. This can have an overwhelming impact on your first stage's design size and weight.

All in all, this suggests to me a parallel-burn rocket and ramjet first stage, boosting to takeover on rocket, flying to staging on ramjet. Staging is in the vicinity of M5 to M6 at somewhere near 60,000 feet, pulling up to somewhere right around 40 degrees at release. The necessarily-supersonic inlet system for stage speeds that high will have a min takeover speed around M1.6 or so, maybe as high as 2.

Any variable geometry ought to be on the inlet, and designed to maintain shock-on-lip, similar to that on the SR-71, but controlled to a different objective (turbine inlets are operated completely differently than ramjet inlets, even though both are made with exactly the same components). Any other variable geometry is just going to be too heavy, and too likely to be unreliable. That's the flying state-of-the-art.

GW

Last edited by GW Johnson (2013-07-23 10:38:59)

RGClark · 2013-07-31 10:37:54

GW Johnson wrote:

...
None of the airbreathers has much frontal thrust density above about 60,000 feet, simply because densities and ambient pressures are so low. Compression ratio is limited to what the inlet can do at that Mach. So thrust depends upon incoming pressure, but weight does not. There is no way around these physics. That is why to pull up sharply at staging in the thin air requires more thrust than the airbreather can deliver, or else the vehicle decelerates sharply as you pull up. The most practical way to achieve pull-up thrust is to fire up some rockets in addition to your airbreather; i.e.; parallel burn, not combined-cycle.
Since altitude is the weakest of three effects at staging, one can stage at 60,000 feet instead of 100+ thousand feet, and get way-to-hell-and-gone better thrust results from the airbreather. This means both velocity and pull-up angle can be fully achievable. Plus, at only 60,000 feet, the time and range to accelerate (on the airbreather alone) to max velocity are a whole lot shorter than those at the higher altitudes in the too-thin air. This can have an overwhelming impact on your first stage's design size and weight.
...
GW

Thanks for that. That report by Sarigul et.al. I cited suggests a launch altitude of 15,000 meters at an angle of 25 to 30 degrees. And the "Conceptual Design of a Supersonic Air-launch System" report suggests 51,000 ft. at a 25 degree angle.
Perhaps these lower altitudes and more shallow angles are something a ramjet could manage without the rocket boost. Still you could imagine for higher thrust the rocket on the orbital stage could supply the extra thrust while still attached to the aircraft with the propellant during this parallel burn phase being supplied from the aircraft.

Bob Clark

GW Johnson · 2013-08-05 09:31:58

Hi Bob:

Once you accelerate an airbreather to very high speed, you are pretty close to thrust = drag, leaving no way to pull-up without decelerating. The force balance just isn't there. That's why I suggested the very easy-to-implement parallel-burn airbreather-rocket approach for the pull-up staging transient, and it is a short transient. If you are using a ramjet for your airbreather, you already have rocket available, which has been the most practical method of accelerating a supersonic ramjet design to its takeover speed.

Missile-style integral rocket-ramjet is not very practical for a reusable design, and neither is a staged-off booster (way too bloody big, to be inexpensive). That suggests parallel-burn, possibly using a liquid rocket engine that uses the same fuel as the ramjet. That's usually a kerosene or a kerosene-like synthetic. So, you're looking at kerosene-LOX rocket engines packaged somewhere in an airframe with one or more kerosene ramjet ducts, and a set of kerosene and LOX tanks. This is all very doable, well-proven, off-the-shelf stuff. No gravy-train technology-development programs here. Those almost never lead to real flying machines.

If you did do an integral solid booster for the ramjet, then you would need to carry along some sort of JATO bottles to get the extra thrust for the pull-up transient. I don't really recommend that design approach, as it has too many one-shot components for a reusable/inexpensive design. The common-fuel all-liquid approach seems to me to have more potential.

The rocket and ramjet might use liquid methane instead of kerosene. This gets you a higher-performing booster for a little less propellant weight. The ramjet can burn it, although vaporization from the cryogenic requires more care in design than vaporization from a room-temperature liquid. At least, the inlet air approaching the combustor is hot (that helps a lot), especially at Mach 5+ as we approach staging. You don't get much of that effect at all in a subsonic/transonic pitot-type design. But you don't get much speed either: about Mach 2 tops.

That being said, I'd recommend a supersonic inlet design with external compression features, a C-D nozzle, a dump-stabilized combustor, and a takeover Mach in the 1.6-to-2 range. I'd also recommend a minimal variable-geometry feature on the inlet to maintain shock-on-lip throughout the flight Mach range. That sort of thing has a top speed potential in the Mach 5-to-6 range, more dependent on vehicle drag than on ramjet design. I'd operate it "full rich" as an accelerator-thruster all the way up, and only lean it back for a Mach 2-ish lean-burn cruise, back to base after staging. I'd also retain just enough rocket propellants on board to support at least one go-around for what is otherwise a dead-stick glide landing.

How to carry the second stage is the hardest part of such a design. But the same basic idea, operated single stage as an edge-of-the-atmosphere skip-glider at Mach 4+ to maybe Mach 6, becomes a possibility for a transoceanic-range replacement for the SR-71. The "air turboramjets" that pushed the old SR-71 had max 25% air bypass to the afterburner duct, and that came from compressor stage 3 or 4, not the supersonic inlet duct. It was limited to about Mach 3.5-3.8 max, because of hot-air temperature limitations on the turbomachinery that the ramjet does not face.

GW

Terraformer · 2013-08-09 17:10:12

I'm still wondering about a ramjet-rocket SSTO, dual engined... you're going to be including the rocket engines into your ramjet craft anyway. Could you build a spaceplane that takes off under rocket power, accelerates to M1.5 with them before switching to a ramjet to reach M6, and then switches back to rockets for the rest of the flight to orbit - and make it reusable, much like the scramjets were supposed to be doing? What mass ratio would it have?

It would save on a lot of problems involved in staging, and allow for people/organisations/governments to buy and operate their own orbital capability (but then again, so would a TSTO system I guess). Such a craft could probably pay for a lot of the development costs in space tourism and suborbital flights, once you've got a suborbital craft. From there, you've got to make it bigger, faster, and a lot more heat resistant...

GW Johnson · 2013-08-09 22:28:11

All the numbers I have run point to two stages with all known non-nuclear rocket and ramjet technologies and materials that we have. The numbers just aren't there for SSTO, not at practical structural fractions, and payloads big enough to be worthwhile.

Fundamentally, there's no reason why both stages of a TSTO cannot be reusable, and this is true whether you design for HTO or VTO. A practical SSTO will require some sort of propulsion breakthrough (yes, I know it can technically be done right now, just with impractically-small payload fractions). I have a lot of hopes pinned on Skylon with its Sabre engine for such a breakthrough, but I'm not betting the farm on it.

GW

Terraformer · 2013-08-13 19:45:31

Would you consider it best to stage at say 2km/s, or use the rockets to boost to 3km/s before staging? I'm thinking that, in the latter case, the upper stage can have a lower mass ratio (so more mass to make it reusable) and the staging can be done in a vacuum (so no aerodynamic issues). It shouldn't shift the mass ratio of the lower craft so much that it becomes hard to reuse...

Just give me a few years to become a billionaire, and I'll hire you to help build such a TSTO system Cererean Labs Aerospace Division.

RGClark · 2013-08-14 02:11:23

GW Johnson wrote:

All the numbers I have run point to two stages with all known non-nuclear rocket and ramjet technologies and materials that we have. The numbers just aren't there for SSTO, not at practical structural fractions, and payloads big enough to be worthwhile.
Fundamentally, there's no reason why both stages of a TSTO cannot be reusable, and this is true whether you design for HTO or VTO. A practical SSTO will require some sort of propulsion breakthrough (yes, I know it can technically be done right now, just with impractically-small payload fractions). I have a lot of hopes pinned on Skylon with its Sabre engine for such a breakthrough, but I'm not betting the farm on it.
GW

As you know, I don't agree with that. As stated in my sig file it's not even hard. The breakthroughs in engines and lightweight stages were all made in the 70's. I'd like to see your numbers that say you can't carry significant payload.
Sure you can carry more with two stages, but that's not the same thing as saying you can't carry significant payload with one. Here's an analogy: you can carry more payload with three stages than two. But nobody would conclude from that, that you can't carry significant payload with two stages, or that it's not worthwhile to make two stage systems. But that is exactly what people say in regards to single stage compared to two stage systems.

Bob Clark

Terraformer · 2013-08-14 21:05:24

Given the requirement for inerts in the 20-30% range, we're going to have to make the lower stage responsible for more of the delta-V - there is no way we have inerts like that on a stage if it has to give 7-7.5km/s delta-V. 2km/s for staging is far too low. 3-4km/s is much more reasonable - and will allow for a pretty nifty suborbital spaceplane for tourism, hypersonic transport, and using off the shelf upper stages to put small satellites (and other payloads) into orbit. If SS1 is any indication, such a spaceplane could be developed for $1-200 million. Outside my current assets, but maybe in a decade I'll be able to pay for it out of my own pocket. From there, you need a bigger version capable of carrying the orbiter, plus of course the orbiter, so maybe $300 million? Or is half a billion dollars for a reusable space transportation system too optimistic (but see SpaceX - not reusable, but capable of launching a lot more)?

Whoever develops it though should remember that Airbus and Boeing don't run airlines. Build them to order, and sell them to whoever orders, be they companies, governments or individuals.

Oh, and in that case, don't develop it inside the US. ITAR would mess it up pretty bad (which I will add to my mental list of why my US citizenship is looking like a liability).

Terraformer · 2013-08-14 21:19:27

GW, I'd like to hear your response to this post on Selenian Boondocks from a while back. My first few thoughts are that the Kerolox lower stage is still going to need some thermal protection anyway, and the ramjet is going to be using the same fuel tanks as the booster rocket.

GW Johnson · 2013-08-20 09:23:07

Hi Terraformer:

I took a look at the blog posting via the link in your post just above. Here's my response:

Well, I think arguments based on mass ratio and “average” Isp are too crude to get you anything useful for ramjet (or any other airbreather). You need a real cycle analysis, which should be a subroutine in a real trajectory code, or which you can use for point performance calculations over a flight envelope for empirical correlation. (You still need a real trajectory code.) I’ve done both, they’re both effective approaches to first order. Fixed averages are not. Sorry, that’s a simple fact-of-life.

You do need to understand thrust and drag accounting, because if you don’t, it is really easy to leave out some very important drag forces in your force balance. I’m not talking about basic ram drag here (the “airbreather’s burden”), I’m talking about things like additive drag, spillage drag, diverter drag, and bleed drag. These are quite important, both at takeover, and at very high flight speeds. These are neither trivial to understand, nor trivial in their effects. You need some training in propulsion aerodynamics for these. This isn’t basic physics textbook stuff, and never will be.

All this stuff will be in my ramjet book, which is not yet ready for publication. Its intended audience is engineers working in ramjet propulsion, whether for missiles, or for launch vehicle work. I’m still trying to “rough-write”-down all the topics, but I think I have most of them documented in rough form, just not all of them. All this stuff is currently very rough first-draft stuff, and will need extensive re-organization and re-write, before it is book-ready. But, I really am working on it.

There are two speed ranges for ramjet design, “low” and “high”. Low speed range designs have simple pitot (normal-shock) inlets, convergent-only nozzles, and can be ignited at subsonic speeds. They will show nacelle thrust greater than drag down to very low speeds, but will have specific impulse lower than composite solid rocket, below about half a Mach number. Peak specific impulse potential is at about Mach 1.1 or so, at about half or 2/3 the max Isp potential of supersonic designs. Max useful speed is about Mach 2, or maybe Mach 2.5 at the very outside. With hydrocarbon fuels of almost all types, about the biggest nozzle throat/combustor area ratio is 0.65, limited by flame-holding considerations. Performance at lower area ratios is inherently lower.

High speed-range designs feature external compression features like ramps or spikes that protrude ahead of the inlet cowl lip. They also have almost-zero thrust potential below about Mach 1.6 to 2. But, they work just fine to about Mach 5-or-6, depending far more on vehicle drag characteristics, than anything about the ramjet engine design. With kerosene fuels, peak Isp potential is around 1200-1300 sec at about Mach 2.5-ish, lower slower, and lower faster. Nozzles are C-D, but exit “bell” area ratios are closer to 1.5-max, than anything to do with the expansion ratios one sees in rockets. With hydrocarbon fuels of almost all types, about the biggest nozzle throat/combustor area ratio is 0.65, limited by flame-holding considerations. Performance at lower area ratios is inherently lower.

These things can be very lightweight, depending upon whether it has to be re-usable or not. The “best” designs have been one-shot missile designs, with an ablative combustor liner, for missile speeds up to about Mach 4. External heat protection is also an issue, from about Mach 3 on up for reusable designs, even with steel construction. There are air-cooled perforated liner designs from the 1940’s and 1950’s that would actually work to Mach 6 on a transient, exclusive of external heat protection problems. There are ablatives that would work externally to Mach 6 on a transient, but these have replacement issues. Missiles generally always use ablatives inside, and maybe outside, if needed.

There is my oddball ceramic-ceramic composite combustor liner material, which offers considerable potential for a re-usable combustor. It might also serve as external heat protection, for a fully-re-usable design. This is still an experimental material, though.

Ramjets require boosters to reach takeover speed: about Mach 0.5 to 0.8 for “low speed” designs, and about Mach 1.6-to-2 for “high speed” designs. For one-shot missile applications, the best choice has proven to be the “integral rocket-ramjet”, wherein a solid rocket booster is cast or loaded within the ramjet combustor. This requires an appropriate ejectable booster nozzle nested within the ramjet nozzle, and some sort of inlet duct obturator, usually ejectable or frangible port covers. Re-usable launch applications might well be “best” with parallel-burn rocket and ramjet engines in the same airframe. It really helps if the rocket and the ramjet use a common fuel.

From a flame-holding standpoint, I think the dump combustor has “way-to-hell-and-gone” more potential than the V-gutter, or can, or “colander” (or any other type of obstruction-type) flameholder. Dump combustors have very little sensitivity to dump plane speeds, compared to any of the blockage-element types. Variable speeds at the dump are inherent with launch accelerators, whether vertical-launch or horizontal takeoff. Almost no textbooks describe dump combustors. My book will.

Ramjet liquid fuels can be any kerosene (or kerosene-like synthetic), or any liquifiable hydrocarbon. The early engines with subsonic ignition used mainly low-grade gasoline. Today, in supersonic-inlet designs, JP-4, JP-5, JP-7, Jet-A, Jet-B, Jet-A1, RP-1, K-1 kerosene, a synthetic variously known as RJ-5 or Shelldyne-H, and even liquefied methane, are all very attractive candidates.

I have even used propane, but it and LPG are not all that attractive, for their inherently-heavy fuel storage considerations. LCH4 will require extra care to insure full vaporization, and extra care with flame-holding issues. RJ-5 is a synthetic that resembles kerosene, except that its density is substantially higher. It was used in ASALM-PTV, with one test that reached Mach 6.

I hope the book might be available in a year or two. It’ll be the ramjet analog to the famous (or infamous) “drag bible” written long ago by Hoerner.

GW

Terraformer · 2013-08-21 15:12:51

What do you think the inert fraction could be on the craft? Presumably it would be built from titanium alloy - even at $22k per metric tonne, I don't think the materials cost is going to be a major factor in the cost of the spaceplane. Could we get re-usability at 20% inerts, allowing the upper stage to be 10% of the GLOW (I see you have a fuel fraction of 58% for the first stage in one of your designs, and this would be 70%)?

GW Johnson · 2013-08-22 08:23:00

Hi Terraformer:

I'd be surprised if a single-stage vehicle with a 20% inert weight fraction was actually robust enough to be long-life reusable. The X-15 was very reusable at 40% inert, but did not take off by itself. Most high-speed (subsonic!) bombers are around 50-60% inerts with all-metal construction, and have been since WW2.

If you replace around half the structure with composites, you might achieve around 27% inert in what was 40% all-metal. You can't replace it all, because composites have such a low tolerance for heating, which happens in-spades on entry, and somewhat (or more) on ascent.

Work-in a "reasonable" payload fraction, like 5 to 10% to make the economics easier to achieve, and there is (with the 27% inerts) about 63 to 68% of the vehicle's gross weight available to be propellants. Those are mass ratios of 2.7 to 3.1, too low for SSTO with chemical rocket propulsion, even as a vertical launch rocket (for min gravity and drag losses). It's worse with the depressed trajectories HTO spaceplanes fly: very large gravity and drag losses.

The airbreather might help, but its higher Isp is not constant, and only available over a small portion of the trajectory. Further, you have to blend it with rocket thrust, in order to climb, especially in the thin air around 60,000 feet. There's not much frontal thrust density available from the airbreather above that altitude. That's why I doubt that ramjet might make SSTO spaceplanes possible, when you restrict the design to larger payload fractions for the economics. TSTO, maybe.

GW

Terraformer · 2013-08-22 15:16:39

Well, I was talking about it in the context of the first stage of a TSTO. How fast could we get, assuming LOX/Methane for the rockets and ramjets, with a mass ratio of 3? With Methlox only, that would allow a craft to reach about 4km/s. Adding in the airbreather, could we reach 5km/s, leaving our upper stage to reach orbit in only 4.5km/s - allowing reusability with a slight payload? Or, using disposable upper stages, cheaply lob 1-2 tonne payloads into orbit, enough for a small crew capsule (maybe even spaceplane?), satellites, or cubesat scale probes...

GW Johnson · 2013-08-23 08:17:30

You'll have to use the airbreather for that portion of the trajectory where speeds and densities are appropriate.

A good supersonic-inlet (subsonic-combustion) ramjet design, aimed at high speed capability, will give useful thrust and impulse between takeover at around 0.5-0.6 km/s velocities, and its top speed of about 1.5 to 1.8 km/s. Those are predicated on stratosphere speeds of sound (295 m/s). The net effect on a depressed HTO trajectory so that the airbreather supplies 100% of the thrust during its operation is an acceleration through 1.1 km/s out of the required 7.7-plus-gravity-and-drag-losses. Drag losses are simply enormous on depressed trajectories, however.

If you do it as ramjet assist on a vertical launch trajectory, you go for the low-speed design with pitot inlet and lower performance capabilities. The rocket leaves the sensible atmosphere at around 80,000 feet (25-ish km) at around Mach 2-ish (around 0.6 km/s), long before the first stage burns out at 3-ish km/s in vacuum. That's where you stage off the airbreather, which becomes useless as the air thins to nothing. That kind of airbreather ignites subsonically at about Mach 0.7-ish (0.2-ish km/s), and might supply at most 20% of the ascent thrust.

To burn at speeds resembling 4 km/s (Mach 13-ish), would require scramjet. Scramjet is simply not yet "ready-for-prime-time", despite the 2 (of 4) successes achieved by the X-51 test vehicle. Those flights demonstrated engine burns in minutes at Mach 5. ASALM-PTV the ordinary ramjet accidentally (!!!!) flew Mach 6 3 decades ago. It was only designed to cruise Mach 4, but we had a throttle runaway on one flight test. Subsonic combustion dump combustor, kerosene-like fuel, supersonic inlet and nozzle, technologies considered well-developed since WW2-to-late-1960's.

I'm not sure vertical launch ramjet assist is really worth the trouble, but that answer remains to be found. For HTO spaceplanes, ramjet looks pretty promising, if you can stage near Mach 5 to Mach 6, at about 60,000 feet (just under 20 km), and manage to pull up to about 40 degree path by adding in some rocket thrust somewhere, all on a short transient. Your first stage is then a fully-reusable rocket/ramjet-equipped airplane. No new technologies need be developed to build it, this is all "prime-time-ready" stuff.

The delta-vee required of the second stage is (unfortunately) near 5.9 km/s, so the propellant fraction is inherently pretty high. At reusable inert fractions, payload fraction is inherently low, which is bad for economics. However, there is always a niche for this capability, down in the low payload mass range, where the first stage aircraft is not too large.

GW

louis · 2013-08-23 18:02:58

Don't think I'd seen the videos featuring Musk at the National Press Club (scroll down, bottom right).

http://www.spacex.com/news/2013/03/31/r … -planetary

Some interesting background on his views on reusable rockets and how that all links up to Mars colonisation.

RGClark · 2013-08-26 19:59:24

Argues Orbital Sciences can get a smaller rocket still able to get high payload to orbit by using the old Atlas rockets because of their remarkably high mass ratio's:

The Coming SSTO's: Page 2.
http://exoscientist.blogspot.com/2013/0 … age-2.html

Bob Clark

Last edited by RGClark (2013-08-26 20:03:31)

New Mars Forums

Announcement

#201 2012-08-29 10:26:02

Re: Reusable Rockets to Orbit

#202 2012-10-16 10:23:15

Re: Reusable Rockets to Orbit

#203 2012-12-27 01:43:42

Re: Reusable Rockets to Orbit

#204 2012-12-27 09:27:18

Re: Reusable Rockets to Orbit

#205 2012-12-29 13:00:06

Re: Reusable Rockets to Orbit

#206 2012-12-30 07:56:57

Re: Reusable Rockets to Orbit

#207 2013-07-22 02:00:08

Re: Reusable Rockets to Orbit

#208 2013-07-22 09:37:39

Re: Reusable Rockets to Orbit

#209 2013-07-22 23:53:41

Re: Reusable Rockets to Orbit

#210 2013-07-23 10:35:18

Re: Reusable Rockets to Orbit

#211 2013-07-31 10:37:54

Re: Reusable Rockets to Orbit

#212 2013-08-05 09:31:58

Re: Reusable Rockets to Orbit

#213 2013-08-09 17:10:12

Re: Reusable Rockets to Orbit

#214 2013-08-09 22:28:11

Re: Reusable Rockets to Orbit

#215 2013-08-13 19:45:31

Re: Reusable Rockets to Orbit

#216 2013-08-14 02:11:23

Re: Reusable Rockets to Orbit

#217 2013-08-14 21:05:24

Re: Reusable Rockets to Orbit

#218 2013-08-14 21:19:27

Re: Reusable Rockets to Orbit

#219 2013-08-20 09:23:07

Re: Reusable Rockets to Orbit

#220 2013-08-21 15:12:51

Re: Reusable Rockets to Orbit

#221 2013-08-22 08:23:00

Re: Reusable Rockets to Orbit

#222 2013-08-22 15:16:39

Re: Reusable Rockets to Orbit

#223 2013-08-23 08:17:30

Re: Reusable Rockets to Orbit

#224 2013-08-23 18:02:58

Re: Reusable Rockets to Orbit

#225 2013-08-26 19:59:24

Re: Reusable Rockets to Orbit

Board footer