Escape Dynamics and the Microwave Thermal Spaceplane

Iskierka · February 18, 2015

As I've pointed out repeatedly, Microwave Thermal Thrusters get more than twice the Specific Impulse (ISP) of Hydro/LOX. And, they're much lighter than chemical engines (for 2-3x the TWR, due mostly to lower mass rather than higher Thrust). So it's more like you can't afford *NOT* to carry a Microwave Thermal Thruster- in that it will MASSIVELY reduce your mass-requirements (pun intended).

You are actually forgetting a major factor here, and that is that while hydrolox may be only 450 seconds or so, resulting in a lot of mass, nearly all of that mass is concentrated in very dense liquid oxygen, which only requires minute tanks. While 800 seconds is a big increase over this, with pure thermal addition, no chemical reaction, you're carrying only hydrogen, which is much, much, much less dense, and 800 seconds actually does not reduce the amount of hydrogen required - this type of vehicle would not, under any circumstances, be small, as it would require colossal LH2 tanks to carry the fuel proposed. This vehicle would have to be considerably larger than Skylon for less payload, so the sketches offered on their website are totally unrealistic to begin with. Remote transmission can't compete with the sheer energy density of stored LH2/LOx.

This vehicle would also run into further problems with acceleration, which you've attempted to dismiss without actually considering properly. Firstly, acceleration is the MOST power-demanding phase of flight, as there is the least airflow available, and the aircraft needs to accelerate enough to get lift in a very limited space. Tangential to this, you're suggesting that these engines could somehow be lighter - but to get any notable thrust below mach 1, it would require turbines, and the highest TWR typically achievable by turbine engines, even those that do not need massive heat exchanger technology, is around 12. SABRE engines can peak at around 14, after intense development by the current world leaders in heat exchanger technology, and using the huge energy density of LH2 fuel to provide the actual propulsive energy - a MTT could not even come close to competing with this in any world, as it lacks all of the thermodynamic features that allow SABRE to target such a performance level.

Once airborne, this vehicle would not be capable of attaining particularly high velocity without the use of fuel. Very rapidly the airflow would exceed the heat exchanger's ability to add heat to it, resulting in near-zero thrust as early as mach 2-3, and no part of the vehicle would be capable of active cooling against mach heating, as no fuel is being used which could be jettisoned before reaching vapour temperatures. Once switching over to fuel use, far earlier than desirable for a multi-cycle airbreathing engine, it would have to target the most efficient ascent profile, which is "up as much as possible", meaning the wings have provided very little benefit over the course of this vehicle's ascent. Spending any longer in a shallow ascent is simply wasteful of fuel, and would require the vehicle to be designed ever larger and heavier.

In conclusion, the reason no-one has flown a beamed power vehicle of any kind, microwave or not, is because beamed power cannot compete with the immense energy available simply by using a conventional fuel. You say you can get 1 MW microwave transmitters at a market price of $2 million per unit - how many of these do you need? Consider that a conventional fuel costs almost nothing, and a simple fuel rocket engine is a few tens of millions, at most. A Delta-IV main engine produces 11 GW useful propulsive energy at liftoff for that kind of price - so even if you could get great power efficiency on the order of 50%, meaning 22 GW from source to target, that would be around 45 billion USD to set the system up? You'd likely need comparable power for comparable payload, as you're suggesting 2.5x the exhaust velocity, resulting in a less-than-proportional mass reduction, and more-than-proportional power increase.

Beamed power is an interesting concept - but it takes crazy dreams to think it can come close to the energy a vehicle carrying conventional fuel brings with it, and even if that means carrying more mass, it's much less mass than the additional energy can deal with.

Northstar1989 · February 18, 2015

You are actually forgetting a major factor here, and that is that while hydrolox may be only 450 seconds or so, resulting in a lot of mass, nearly all of that mass is concentrated in very dense liquid oxygen, which only requires minute tanks. While 800 seconds is a big increase over this, with pure thermal addition, no chemical reaction, you're carrying only hydrogen, which is much, much, much less dense, and 800 seconds actually does not reduce the amount of hydrogen required - this type of vehicle would not, under any circumstances, be small, as it would require colossal LH2 tanks to carry the fuel proposed. This vehicle would have to be considerably larger than Skylon for less payload, so the sketches offered on their website are totally unrealistic to begin with. Remote transmission can't compete with the sheer energy density of stored LH2/LOx.

The sketches are just artist-renditions, and are indeed wildly-inaccurate (not surprising, really- most artists who draw about space don't even have the faintest clue about rocket-science...) And, you are also correct in that the fuel-density would be abysmal. But you are missing two thing here:

(1) Unlike with Skylon, you don't have to carry any fuel for the initial ascent. You can achieve this purely with Thermal Turbojets.... (which use the atmosphere for 100% of their reaction-mass)

(2) The vehicles is much lighter, so requires much less propellant to reach orbit to begin with, and less Thrust to stay airborne. A Microwave Thermal Thruster weighs much less than a SABRE engine, and your fuel to reach orbit weighs much less as well...

(3) You're not limited purely to Liquid Hydrogen. You can use Liquid Methane for an initial "kicker" of higher Thrust, and save the higher-ISP Liquid Hydrogen for later in the flight.

(4) Both Hydrogen and Methane will combust with atmospheric Oxygen. So it's possible to have an intermediate propulsion-mode between a purely-atmospheric Thermal Turbojet, and a Thermal Rocket driven purely by internal propellant, where you essentially have a thermally-augmented SABRE engine (that is, you pump additional heat into the exhaust stream to increase Thrust and Exhaust Velocity- allowing operation either with a greater bypass-ratio than a SABRE for the same Exhaust Velocity and thus high-speed performance...) The added complexity and mass may not be worth it for the very limited window of time where it outperforms both a Thermal Turbojet and a Thermal Rocket, however...

This vehicle would also run into further problems with acceleration, which you've attempted to dismiss without actually considering properly.

I've considered it. No need to attack me, but go on...

Firstly, acceleration is the MOST power-demanding phase of flight, as there is the least airflow available, and the aircraft needs to accelerate enough to get lift in a very limited space.

Correction. Takeoff is the most THRUST-demanding phase of flight. At the altitude of the runway, there is a LOT of airflow available to potentially increase your working-mass. If your air intakes are large enough compared to your engines, you can easily achieve very high ratios of working mass to input-power. Which means very high Thrust relative to your input-power at low speeds. You will need lots of Intake Area at high-altitudes anyways in order to keep enough airflow to your Thermal Turbojets (although you will also need advanced precoolers for the very high compression-ratios, which is why SABRE's new precooler designs could be IMMENSELY helpful for a Hydrogen-powered Microwave Thermal Spaceplane...) Your takeoff velocity will also be VERY low due to the incredibly low wing-loading involved in a Hydrogen Thermal Spaceplane (due to the low propellant density) with any kind of decent ratio of wing area to fuselage size. A simple rocket-sled or electromagnetic "kicker" will be enough to give your plane a burst of speed on the runway if all that still doesn't cut it.

Tangential to this, you're suggesting that these engines could somehow be lighter - but to get any notable thrust below mach 1, it would require turbines, and the highest TWR typically achievable by turbine engines, even those that do not need massive heat exchanger technology, is around 12.

Are you talking high-bypass or low-bypass turbofans? Because there's a BIG difference...

Also, I said that a Microwave Thermal Rocket is lighter than a chemical rocket, not that a Microwave Thermal Turbojet is lighter than a conventional jet-engine. See below...

SABRE engines can peak at around 14, after intense development by the current world leaders in heat exchanger technology, and using the huge energy density of LH2 fuel to provide the actual propulsive energy - a MTT could not even come close to competing with this in any world, as it lacks all of the thermodynamic features that allow SABRE to target such a performance level.

The lack of *WHAT* thermodynamic features?

I assume you're referring to having large stores of onboard Liquid Hydrogen enabling SABRE to use that as a heat-sink for the precoolers. But as we've already discussed, a Microwave Thermal Spaceplane would ALSO have that same available heat-sink, in the form of large supplies of onboard Liquid Hydrogen (in fact, as you already pointed out, it would have MORE Liquid Hydrogen onboard, as 100% of its rocket-propellant mass would be LH2, whereas a Hydrolox Rocket carries more than 8/9ths of its propellant mass as Liquid Oxygen...)

The Microwave Thermal Receivers aren't particularly heavy either- in fact, compared to a combustion chamber they're quite light. So, I would expect a Microwave Thermal Turbojet to be able to achieve a TWR of *AT LEAST* 15-20. Which isn't particularly impressive compared to a Microwave Thermal Rocket, which can achieve a TWR of over 200, but a Microwave Thermal Turbojet can produce *MUCH* more Thrust/MW than a Microwave Thermal Rocket...

Once airborne, this vehicle would not be capable of attaining particularly high velocity without the use of fuel. Very rapidly the airflow would exceed the heat exchanger's ability to add heat to it, resulting in near-zero thrust as early as mach 2-3, and no part of the vehicle would be capable of active cooling against mach heating, as no fuel is being used which could be jettisoned before reaching vapour temperatures.

First of all, I think you're mixing up hypersonic and supersonic propulsion- at supersonic velocities compression-heating of the intake air is not NEARLY as large of an issue as you think, and it's still possible to pump significant amounts of heat into the working-mass (the Heat Exchanger operates at a temperature of over 2000 K- I'd like to see an intake airstream that heats *NEARLY* that hot below Mach 3).

As for hypersonic propulsion (let's define this as starting at about Mach 4) you have a HUGE heat-sink in the form of Liquid Hydrogen available, and a gigantic surface-area you can pump coolant from the precooler to in order to allow it to give back its heat to the atmosphere thanks to the very low fuel-density (not all of the airframe heats up- there are parts of a hypersonic plane that are shielded from the compression-shockwave and actually remain quite cool at hypersonic velocities...)

Of course, the LH2 can still only absorb so much heat before boiling off, so here is where you might want to start thinking about switching over to a fuel-consuming mode: possibly initially to a sort of microwave-augmented SABRE design (where the Microwave Thermal Receivers do their best to add heat to the exhaust gasses from a LH2/Atmosphere combustion reaction, allowing for use of a much higher bypass-ratio at the same Exhaust Velocity if enough intake air is available...), although the numbers for engine-mass and the requisite amount of airflow necessary to operate something like this might not be favorable, and then to Microwave Thermal Rocket mode.

With a "hybrid" Thermal Turbojet design you would shut down airflow to some off the turbojets, and instead switch them over to passing internal fuel (Liquid Hydrogen) over the same Microwave Thermal Receiver. You would want multiple Hybrid Turbojets at this point, though, as the rockets would allow you to continue operating some of them in airbreathing-mode to higher altitude-speed by creating a ram-effect driving more airflow into the intakes...

Once switching over to fuel use, far earlier than desirable for a multi-cycle airbreathing engine, it would have to target the most efficient ascent profile, which is "up as much as possible", meaning the wings have provided very little benefit over the course of this vehicle's ascent. Spending any longer in a shallow ascent is simply wasteful of fuel, and would require the vehicle to be designed ever larger and heavier.

Here is the crux of the matter. I think you misunderstand the purpose of the Thermal Turbojets and the wings. They exist *NOT* to provide a significant portion of orbital velocity (indeed, you could probably only make it up to about Mach 4-5 at best using Thermal Turbojets and SABRE-style precoolers), BUT TO LIFT THE ROCKETS UP ABOVE THE THICKEST PART OF THE ATMOSPHERE TO REDUCE ATMOSPHERIC-COMPRESSION OF THE EXHAUST-STREAM.

Due to the *VERY* low wing-loading of a Hydrogen-fueled design (due to the incredibly low fuel-density), you would expect to be able to reach significant altitudes even at only Mach 4-5 (or Mach 2-3 for that matter) using Thermal Turbojets alone. At that point, you would switch over to rocket-propulsion at a MUCH lower ambient pressure than if you tried to operate a Microwave Thermal Rocket right off the launchpad. Which means higher Specific Impulse for the same Exhaust Velocity and Mass Flow Rate, which means more Thrust, and more Thrust/MW. And, as the cost-limiting factor on a Microwave Thermal Spaceplane or Rocket is the cost of the Microwave Transmitters, ANYTHING you can do to improve the Thrust/MW is going to be useful.

Let's say your sea-level ISP is only 360 seconds using a Microwave Thermal Rocket. But you vacuum ISP is 850 seconds! If you fly up above the thickest part of the atmosphere using Thermal Turbojets, you can more than DOUBLE your Thrust/MW!

Additionally, even as a rocket-propelled plane the whole way up (which is what Escape Dynamics is actually working on- not a design using Thermal Turbojets), the Microwave Thermal Spaceplane will still have superior cost-effectiveness to a conventional rocket propelled by Microwave Thermal Thrusters. This is as the wings allow you to ascend with a TWR much less than 1. By the time you go ballistic, and actually *NEED* a higher TWR, you should be at such a low atmospheric pressure and have consumed so much of your original fuel-mass that you should EASILY be able to achieve the necessary TWR with the same amount of beamed-power that only gave you a TWR of maybe 0.2 or 0.3 when you first activated the Microwave Thermal Rockets...

In conclusion, the reason no-one has flown a beamed power vehicle of any kind, microwave or not, is because beamed power cannot compete with the immense energy available simply by using a conventional fuel.

There is no basis for that conclusion in your arguments, and I've poked your arguments full of holes anyways...

You say you can get 1 MW microwave transmitters at a market price of $2 million per unit - how many of these do you need?

How many MW of mcirowaves do you need? Now THAT'S an interesting question...

At sea-level you can get more than 1 kN of Thrust per MW using Microwave Thermal Rockets with Liquid Hydrogen (the Timberwind Nuclear Thermal Rocket designs got about this at *MUCH* higher Exhaust Velocities and Temperatures, and thus inferior Thrust/MW). Using pure Liquid Methane, you can get more than 2.8 kN of Thrust per MW- but at much lower Specific Impulse (around 300-360 seconds, so in the same range as Kero/LOX rockets, but at lower fuel-density). And, of course, your Thrust/MW is *much* better at lower atmospheric pressures...

Using LOX-injection for an afterburning-effect, you can get higher Thrust than the Molecular Mass of the exhaust gasses would otherwise dictate for a purely Thermal design (Water will only net a bit over 3 kN/MW at an ISP of about 270-290 seconds, for instance- but LH2/LOX will produce significantly more Thrust for the same Mass Flow Rate and MW of Thermal Power, thus improving both Thrust/MW *AND* Specific Impulse...) The beauty of a Microwave Thermal Rocket is that it's fuel-agnostic, meaning you can easily switch from N2 (an even denser propellant with an even higher Thrust/MW but lower ISP) to Methane to LH2 fuel-modes in-flight if you want with no additional equipment. This allows you to start off with denser, lower-ISP propellants, and switch to less dense but higher-ISP propellants as your velocity increases... (thus more closely matching Exhaust Velocity to your plane/rocket's velocity for maximum energy-efficiency...)

Consider that a conventional fuel costs almost nothing, and a simple fuel rocket engine is a few tens of millions, at most. A Delta-IV main engine produces 11 GW useful propulsive energy at liftoff for that kind of price - so even if you could get great power efficiency on the order of 50%, meaning 22 GW from source to target, that would be around 45 billion USD to set the system up? You'd likely need comparable power for comparable payload, as you're suggesting 2.5x the exhaust velocity, resulting in a less-than-proportional mass reduction, and more-than-proportional power increase.

Your math is off. Do you even understand the Rocket Equation? Decreasing your Specific Impulse EXPONENTIALLY increases your total fuel-requirements, and thus your required Thrust at every point in your flight.

A Delta IV (in its standard configuration) has 3,140 kN RS-68A engine for its launch stage, and weighs over 250 metric tons on the launchpad, for a liftoff TWR of roughly 1.28 with a payload-capacity of 9.42 metric tons to LEO.

A comparable Microwave Thermal Rocket, on the other hand, would only require a roughly 62.8 ton rocket (official estimates of payload-fraction with one are about 15% using LH2 the whole way, due to the MUCH higher ISP) and thus only about 790 kN of Thrust on the launchpad for a slightly greater liftoff TWR.

That means about $1.58 billion in Microwave Transmitters at current costs- which means it would only take about 17 launches to amortize the cost of the transmitter-array down to a cost of about $94.2 million/launch (this is actually a bit lower than the cost of a Delta IV launch). Of course, there are also R&D costs, the cost of the Microwave Thermal Rocket (the actual rocket itself is exceedingly cheap compared to a chemical rocket- as Microwave Thermal Thrusters are VERY cheap to manufacture), and the costs of propellant and electricity to worry about.

So, not cheap by any means. There have only been about 28 Delta IV launches in the 12 years since it was developed, which is about the length of time the Microwave Transmitters would last before some of the units in the array started needing replacement... But it *IS* cheaper than chemical rocketry, even if only by a hair.

Of course, none of this accounts for the following:

(1) Microwave Transmitter costs have declined DRAMATICALLY over the past 50 years, *ESPECIALLY* the last 20. If they continue to come down (something Escape Dynamics is working very hard on, and has already had some success with in its experimental units) then the economics should only become MORE favorable.

(2) Spaceplanes DON'T NEED as high a TWR as rockets to get to orbit. A 62.8 ton spaceplane doesn't need 790 kN of thrust on the runway- it can EASILY get off the ground with just 200 kN of Thrust. That means a third the cost in Microwave Transmitters just on that basis alone- and that's BEFORE you account of the fact that a Thermal Turbojet could probably achieve Thrust/MW ratings many-fold higher than Microwave Thermal Rockets... (if you can merely get *twice* the Thrust/MW with TTJ's, and climb to an altitude where your Microwave Thermal Rocket ISP using LH2 is 720 seconds instead of the 360 or seconds you get at sea-level, then you can reach orbit with less than half the power-requirements again, for only 1/6th the Microwave Transmitter costs of a rocket with the same mass on the launchpad)

(3) If you include a Methane or LH2/LOX "kicker" when you first activate your rockets (REMEMBER, Microwave Thermal Rockets are fuel-agnostic, you only need additional equipment if you decide to use LOX afterburners), you can get by with EVEN LESS beamed-power, whether using a rocket or spaceplane. Your vessel mass (and Thrust requirements) will go up, but you get 2.82 (to be precise, the square-root of the ratio of the Molecular Mass of CH4 to H2) times the Thrust/MW. As your vessel mass decreases due to fuel-consumption, and your Thrust requirements go down, you then switch over to LH2 propulsion- allowing you to get by with slightly less beamed power and a significantly more compact plane/rocket (due to the higher fuel-density of Liquid Methane or LH2/LOX than pure LH2...)

Beamed power is an interesting concept - but it takes crazy dreams to think it can come close to the energy a vehicle carrying conventional fuel brings with it, and even if that means carrying more mass, it's much less mass than the additional energy can deal with.

No, and no again. The numbers say otherwise. The pace of human progress (and future progress that is expected in Microwave Thermal Rockets to improve the performance and costs even further) say otherwise as well.

Regards,

Northstar

Edited February 18, 2015 by Northstar1989

Northstar1989 · February 19, 2015

I wanted to explain the Methane-Kicker thing a little better, so here goes my best shot...

Methane gets 2.82 times the Thrust/MW in vacuum due to the following equations:

E = 1/2 m v²

Thrust = Mass Flow Rate * Exhaust Velocity - Exit Area * Background Pressure

The first of these equations describes the relationship between energy and mass/velocity, and can be used to *approximate* the relationship between the Molecular Mass of a propellant and the Exhaust Velocity (there are also components for the specific heat capacity of the propellant and the expansion-ratio of the engine nozzle in the equation for Exhaust Velocity, with the heavier fuels having lower specific heat capacity but a larger optimal expansion-ratio at any given ambient pressure, so let's just ignore those for now and say this equation is all we need...)

The second equation describes the relationship between Mass Flow Rate, Vacuum Specific Impulse (which is equal to the Exhaust Velocity divided by "g", i.e. 9.8 m/s^2), ambient atmospheric pressure, and the size of the rocket nozzle (the "Exit Area"). You'll notice that for a given-sized rocket nozzle, the losses to atmospheric-compression are constant for a given ambient pressure- so the difference between Vacuum ISP and Sea-Level ISP are based on the relationship between Vacuum Thrust (MAss Flow Rate * Exhaust Velocity) and nozzle size (Exit Area).

Anyways, if you assume that you heat two different propellants to the same temperature, then you can set the Energy/molecule of propellant equal for two different propellants, and you get the following:

m₁ * v₁² = m₂ * v₂²

If you do out the math, then the ratio of Exhaust Velocities becomes approximately the following for two propellants at the same temperature:

v₁/v₂ = SqRt (m₂/m₁)

and the ratio of Vacuum Thrust:

Thrust₂/Thrust₁ = m₂v₂/m₁v₁ = SqRt (m₂/m₁)

*THIS* is where the number than Methane produce 2.82 (the square-root of 8) times the Thrust/MW of H2 comes from (where molecular masses are valued at 16 and 2, respectively), as well as that Methane only has (1/2.82) times the Vacuum ISP.

Now, back to the question-at-hand: how does any of this apply to using a "Methane Kicker" in Thermal Rocketry?

Well, let's say you have enough Thrust to keep a 20 ton rocket or spaceplane ascending at the desired rate with LH₂ as the propellant in the atmosphere.

Now, what happens if you replace the FIRST 5 TONS of LH₂ with an equivalent volume of Liquid Methane? (which is approximately 8 times denser/liter)

You now have a rocket that weighs 55 tons, however the Thrust-Weight Ratio is actually GREATER thanks to getting 2.82 times the Thrust from Methane (the rocket fuel only weighs 2.75 times as much, and the dry mass is the same).

Of this mass, 40 tons is Liquid Methane, and 15 tons in LH₂.

What's more, this same rocket actually has slightly more Delta-V for the same sized rocket as well: because 40 tons of Methane produces more Delta-V than 5 tons on LH2 (the TWR starts off just as high as with the 5 tons of LH₂, and you have the same burn-time, but you lose mass much more rapidly, leading to a higher TWR later in the Methane-powered portion of the flight: as much as 2.82 times as high *just before* you run out of Methane).

Now whoa, whoa Northstar, you might say: "why not replace the entire 20 tons of LH₂ with Liquid Methane then?"

Well, you *COULD*, but then you'd get the following...

Rocket Mass: 160 tons

Relative Thrust: 2.82

Relative TWR: 0.3525

So, if your rocket would have the a much lower TWR, and most likely would not have enough Thrust to make it off the launchpad or continue to hold altitude as a plane... Your time-to-orbit would dramatically increase at the very least (ALTHOUGH your Ballistic Coefficient would DRAMATICALLY improve- and you would lose much less Delta-V to atmospheric drag with your rocket), and most likely you would end up needing a lot more Delta-V (and a bigger rocket) to get there as a result...

There are a couple more considerations that favor Methane, however:

(1) Methane produces *MORE* than 2.82 times the Thrust of Hydrogen in-atmosphere. This is because, due to the much higher Vacuum Thrust of Methane (2.82 times as high, remember) but having the same (or a similar) sized rocket-nozzle (the optimal nozzle-size *optimizes* to a larger size, but you get 2.82 times the Vacuum Thrust with the same nozzle- I can explain that later if you guys want) you end up with a much better ratio of Vacuum Thrust to Exit Area * Ambient Pressure, and a smaller difference between Vacuum and Sea-Level ISP as a result... (another way to explain this is that the much higher Mass Flow Rate leads to a higher Exhaust Pressure...) In other words, the Hydrogen-propelled rocket/plane loses relatively more of its Thrust to atmospheric-compression, and the ratio of Thrust with Methane vs. Hydrogen grows as a result.

(2) A rocket with Methane replacing some (or all) of its LH₂ will have a higher Ballistic Coefficient than a rocket with just LH₂, and will lose LESS Delta-V to atmospheric drag during its ascent as a result (however a spaceplane with Methane will have a higher wing-loading unless you increase the size of the wings...)

Anyways, using a Methane-kicker allows you to get more Delta-V for the same fuel volume, and (with a rocket) decreases your Delta-V lost to atmospheric-drag (better Ballistic Coefficient) and gravity (shorter time-to-orbit due to slightly higher TWR and better Ballistic Coefficient). The result is a better-performing rocket (or spaceplane) with the same amount of Microwave Beamed-Power.

Regards,

Northstar

Edited February 19, 2015 by Northstar1989

Northstar1989 · February 19, 2015

Some additional advantages to a Methane-kicker:

(3) Methane has much higher boiling-point (-161.5 Celsius) than H₂ (-252.9 Celsius), making it much easier to store long-term. You can get away with less insulation on the fuel tanks of a launch vehicle containing Methane (reducing mass), and it becomes more feasible to store reserve fuel from the uppermost stage of the rocket or from a spaceplane in orbital fuel depots long-term for use beyond LEO (to boost satellites to Geostationary orbits, for instance).

(4) Because Methane Thermal Rockets have a higher Exhaust Pressure than Hydrogen you can utilize a larger engine nozzle in-atmosphere. The ideal rocket nozzle size expands the exhaust stream until it equals ambient pressure, and with a higher pressure to work with before the nozzle, the ideal nozzle size becomes larger. The larger your rocket nozzle becomes, the more you increase your Exhaust Velocity with the nozzle, and the higher your Specific Impulse and Thrust both become. Thus, while a Thermal Rocket using Methane as propellant might produce 3 times the Thrust at sea-level as a Thermal Rocket using Hydrogen as propellant with the same size nozzle, a Thermal Rocket using Methane at optimal expansion-ratio will generate significantly MORE Thrust at a higher Specific Impulse for the same Mass Flow Rate due to the higher Expansion-Ratio of the optimal nozzle...

Of course, since a Thermal Rocket is fuel-agnostic, there is nothing stopping you from using MORE than two types of propellant for even BETTER performance. For instance, you could do something like this:

CO₂ --> N₂ --> CH₄ --> H₂

As you'll notice, each progressive change in fuel-mode has a lower Molecular Mass that the one before it. This is actually a very clever way of effectively staging the rocket's propulsion (even though there's nothing stopping you from switching fuels like this is a spaceplane or SSTO, and no actual separation of stages may be involved), as each fuel-component will be heavier than the one after it...

Here are the Thrust values of each propellant (note that the Vacuum ISP is in all cases the inverse of the relative Thrust, i.e. Methane has SqRt(8) time the Thrust and 1/SqRt(8) times the ISP compared to H₂...) starting with the values for Hydrogen as a benchmark:

H₂

Thrust: approx. 1 kN/MW

Vacuum ISP: 850 - 1000 seconds

CH₄

Thrust: approx. 2.82 kN/MW

Vacuum ISP: 300 - 354 seconds (*before* accounting for larger nozzle compared to Hydrogen- actual performances of 320-360 seconds are expected)

N₂

Thrust: approx. 3.74 kN/MW

Vacuum ISP: 227 - 267 seconds (larger values expected after accounting for larger nozzle)

CO₂

Thrust: approx 4.69 kN/MW

Vacuum ISP: 181 - 213 seconds (larger values expected after accounting for larger nozzle)

EDIT: EEK! I made a major calculation-error and forgot to divide the Molecular Mass by 2 (the Molecular Mass of diatomic Hydrogen) before calculating the Thrust and ISP of N2 and CO2- resulting in excessively high #'s for Thrust and low #'s for ISP for these two propellants. Now fixed.

Note that although the optimal nozzle size increases with increasing Molecular Mass of the propellant, the optimal nozzle size *also* increases with ALTITUDE- meaning that if you start off with heavier propellants at lower altitude, and work your way to lighter propellants as you ascend, you can keep the same nozzle the whole way through the ascent- as the decreasing ambient pressure will make up for the decreasing Exhaust Pressure...

In every case, replacing a lighter fuel with a heavier fuel makes sense approximately UP UNTIL THE POINT where you end up with the same TWR for the heavier propellant as if you had just stuck with the lighter propellant. Thus, replacing H₂ with CH₄ makes sense up until your rocket has 2.82 times the mass as if you had stuck with just Hydrogen, replacing CH₄ with N₂ makes sense up until you would have 1.87 (5.29/2.82) times the total rocket mass as if you had stuck with just Methane, etc.

At the point where your initial TWR for the denser propellant is equal to your TWR for the lighter propellant, your Ballistic Coefficient will be higher (and thus atmospheric-drag losses lower), your rocket will require less insulation (although ideally you would include the reduced insulation requirements in calculating the new TWR and deciding how much lighter propellant to replace with the heavier one), and your time-to-orbit (and thus gravity-losses) will be shorter (as your TWR will increase more rapidly due to your lower Specific Impulse and thus proportionally higher Mass Flow Rate with the heavier fuel, and less of your velocity will be eaten up by atmospheric-drag due to your higher Ballistic Coefficient).

Thus, when using heavier propellants for the initial stages of your ascent, your Delta-V requirements to orbit will be reduced, and you will have more Delta-V available with the same sized rocket (allowing your overall rocket to be smaller) despite the lower Specific Impulse (your payload-fraction will decline, but your payload-capacity will increase...)

There are a few drawbacks that need to be taken into account, though:

With a spaceplane, your wing-load will increase when swapping in heavier fuels (and you will need larger wings to bring it back down- this must also be taken into account when calculating the new TWR using the heavier propellant). However this has the side-benefit that you will have a lower wing-load when you get to the next (lighter) fuel-mode in line, which will make it EASIER to continue to gain altitude/speed...

With a rocket, your structural mass will increase due to the higher weight of the heavier propellants, as each stage must support the mass of the stages above it (and once again, this must be taken into account when calculating how much lighter propellant is ideal to replace with a denser propellant...) However, this has the side-benefit of making stages safer/easier to recover after jettisoning them- as you will no longer have to contend with the weight of the full fuel tanks or the stages above, and your structure will be stronger relative to the outside forces exerted upon it... (a lower stage that was designed to support 400 tons of rocket above it will be much stronger than one that was only designed to support 100 tons of rocket in stages higher up...)

Regards,

Northstar

Edited February 19, 2015 by Northstar1989

Iskierka · February 19, 2015

(2) The vehicles is much lighter, so requires much less propellant to reach orbit to begin with, and less Thrust to stay airborne. A Microwave Thermal Thruster weighs much less than a SABRE engine, and your fuel to reach orbit weighs much less as well...
(4) Both Hydrogen and Methane will combust with atmospheric Oxygen. So it's possible to have an intermediate propulsion-mode between a purely-atmospheric Thermal Turbojet, and a Thermal Rocket driven purely by internal propellant, where you essentially have a thermally-augmented SABRE engine (that is, you pump additional heat into the exhaust stream to increase Thrust and Exhaust Velocity- allowing operation either with a greater bypass-ratio than a SABRE for the same Exhaust Velocity and thus high-speed performance...) The added complexity and mass may not be worth it for the very limited window of time where it outperforms both a Thermal Turbojet and a Thermal Rocket, however...

The vehicle is only lighter if it is smaller, which a pure LH2 design would not be. A CH4 mode may be an improvement that gets around size issues. Potentially a rocket-only microwave power engine would be lighter, though I personally still find it dubious that it could outperform pressure-fed liquid motors, but any air-breathing thruster has to add in at minimum the vast majority of the mass from any other air-breathing engine, and cannot be claimed to be "much" less.

Are you talking high-bypass or low-bypass turbofans? Because there's a BIG difference...

There actually isn't as big a difference as you think, at least for equal number of spools and comparable purpose. For low efficiency, high-thrust engines, two-spool high-bypass are typically 8-10, low-bypass 10-12 or so. Single-spools of both classes are the ones that achieve 12-14 ranges, though are rarely used due to the lower specific thrust. Triple-spools that get the required specific thrust only appear in high-bypass engines, which is why you can see a noticeable difference of high BPR engines sitting around 6. The highest ever achieved was a single-spool turbojet designed to completely sacrifice all efficiency and endurance considerations, as its design purpose was a short-duration lift engine for VTOL aircraft, achieving TWR of 18.75. Do not expect to easily match that performance, though, with anything that has a design consideration other than "lots of thrust at low speed." The RB.162 was a true exception - nothing else exceeds around 14.5, because nothing else is so highly specialised towards high TWR.

Correction. Takeoff is the most THRUST-demanding phase of flight. At the altitude of the runway, there is a LOT of airflow available to potentially increase your working-mass. If your air intakes are large enough compared to your engines, you can easily achieve very high ratios of working mass to input-power. Which means very high Thrust relative to your input-power at low speeds. You will need lots of Intake Area at high-altitudes anyways in order to keep enough airflow to your Thermal Turbojets (although you will also need advanced precoolers for the very high compression-ratios, which is why SABRE's new precooler designs could be IMMENSELY helpful for a Hydrogen-powered Microwave Thermal Spaceplane...) Your takeoff velocity will also be VERY low due to the incredibly low wing-loading involved in a Hydrogen Thermal Spaceplane (due to the low propellant density) with any kind of decent ratio of wing area to fuselage size. A simple rocket-sled or electromagnetic "kicker" will be enough to give your plane a burst of speed on the runway if all that still doesn't cut it.

It is -also- the most power demanding, as you do not have the airflow available to generate thrust with minimal power - your ideal power/thrust comes once airborne, and the craft's own velocity is gathering air for you. Having larger and larger intakes is going to dramatically increase your mass, and no, you actually do not need large intakes at all at speed and altitude. Every single high-speed supersonic aircraft to date has actually had intakes designed to reduce in area at high velocity - Concorde and the SR-71 had spill doors that made more air available at a standstill, then were used to jettison excess at speed. Skylon's intakes are specifically designed to simply close off most of its area at speed. One of the biggest problems with designing high-speed aircraft to be efficient at low speed is that having the required area at low speed means huge intake drag at high speed.

I assume you're referring to having large stores of onboard Liquid Hydrogen enabling SABRE to use that as a heat-sink for the precoolers. But as we've already discussed, a Microwave Thermal Spaceplane would ALSO have that same available heat-sink, in the form of large supplies of onboard Liquid Hydrogen (in fact, as you already pointed out, it would have MORE Liquid Hydrogen onboard, as 100% of its rocket-propellant mass would be LH2, whereas a Hydrolox Rocket carries more than 8/9ths of its propellant mass as Liquid Oxygen...)

The key mis-assumption here is that the internal fuel can be used as a heat-sink. It can't - Skylon is not designed to store any fuel that has been used as coolant, 100% of it must be consumed, as even a few dozen tons of LH2 cannot absorb that much heat without vaporising. And to actually provide sufficient cooling, MORE LH2 than the core engine can consume is required - hence the SABRE's surrounding array of ramjets, as it's genuinely not possible to simply re-store any of the fuel, so it is used in whatever useful manner it can be. If you're going to try use the LH2 as coolant, then you're already going to use all of it as fuel, and once you're getting all the propulsion you can out of fuel energy, the microwave power is useless as it cannot heat the exhaust any further (material thermal limits; the fuel could already do better than it does if not for these) and is simply excess weight on a lifting ascent stage.

The Microwave Thermal Receivers aren't particularly heavy either- in fact, compared to a combustion chamber they're quite light. So, I would expect a Microwave Thermal Turbojet to be able to achieve a TWR of *AT LEAST* 15-20. Which isn't particularly impressive compared to a Microwave Thermal Rocket, which can achieve a TWR of over 200, but a Microwave Thermal Turbojet can produce *MUCH* more Thrust/MW than a Microwave Thermal Rocket...

This might be a valid point if combustion chambers were actually a significant portion of the mass of an air-breathing engine; they're not, and just moments ago you were suggesting having BOTH, so you cannot claim any significant air-breathing TWR advantage.

First of all, I think you're mixing up hypersonic and supersonic propulsion- at supersonic velocities compression-heating of the intake air is not NEARLY as large of an issue as you think, and it's still possible to pump significant amounts of heat into the working-mass (the Heat Exchanger operates at a temperature of over 2000 K- I'd like to see an intake airstream that heats *NEARLY* that hot below Mach 3).
As for hypersonic propulsion (let's define this as starting at about Mach 4) you have a HUGE heat-sink in the form of Liquid Hydrogen available, and a gigantic surface-area you can pump coolant from the precooler to in order to allow it to give back its heat to the atmosphere thanks to the very low fuel-density (not all of the airframe heats up- there are parts of a hypersonic plane that are shielded from the compression-shockwave and actually remain quite cool at hypersonic velocities...)

Compression-heating on the external parts of a vehicle is not so significant, but stagnation temperature rises very unfavourably to the point of being problematic before even Mach 3 - there's a reason most of the J-58's high speed thrust did not come from its actual turbine section, and specific thrust is quite bad from ramjets. The intake airstream does not have to be that hot, because the compressor does plenty of additional heating anyway, very rapidly bringing intake air to unworkable temperatures.

And again, you do not have any heat-sink on-board. You have a consumable coolant, IF you use it as fuel, and if you do, then the microwave part of this propulsion stage becomes pointless as it cannot add any useful quantity of energy. If this stage is to be microwave-powered, it's very limited in how fast it can go.

Here is the crux of the matter. I think you misunderstand the purpose of the Thermal Turbojets and the wings. They exist *NOT* to provide a significant portion of orbital velocity (indeed, you could probably only make it up to about Mach 4-5 at best using Thermal Turbojets and SABRE-style precoolers), BUT TO LIFT THE ROCKETS UP ABOVE THE THICKEST PART OF THE ATMOSPHERE TO REDUCE ATMOSPHERIC-COMPRESSION OF THE EXHAUST-STREAM.

Which, interestingly, you can achieve much more cheaply, and just as effectively, with a couple of quite small solid boosters or similar high-thrust first stage kick. Skylon actually intends for its air-breathing stage to be useful, bringing the dV to orbit down from 9800+ to around 6700, which allows a mass ratio of less than 5 for its rocket stage, most of which is in a very dense oxidizer, keeping the vehicle (relatively) small. If you just want to get past dense air, which is all a microwave lower stage could reasonably do, it would be far more practical to just kick it up there with cheap boosters. Hell, use LH2 liquid fuel flyback boosters, so your same fuel manufacturing systems can supply the boosters, which are also totally and rapidly reusable - would still be far more effective than a microwave lifting stage.

A comparable Microwave Thermal Rocket, on the other hand, would only require a roughly 62.8 ton rocket (official estimates of payload-fraction with one are about 15% using LH2 the whole way, due to the MUCH higher ISP) and thus only about 790 kN of Thrust on the launchpad for a slightly greater liftoff TWR.
That means about $1.58 billion in Microwave Transmitters at current costs- which means it would only take about 17 launches to amortize the cost of the transmitter-array down to a cost of about $94.2 million/launch (this is actually a bit lower than the cost of a Delta IV launch). Of course, there are also R&D costs, the cost of the Microwave Thermal Rocket (the actual rocket itself is exceedingly cheap compared to a chemical rocket- as Microwave Thermal Thrusters are VERY cheap to manufacture), and the costs of propellant and electricity to worry about.
So, not cheap by any means. There have only been about 28 Delta IV launches in the 12 years since it was developed, which is about the length of time the Microwave Transmitters would last before some of the units in the array started needing replacement... But it *IS* cheaper than chemical rocketry, even if only by a hair.

I'm sorry, but I can't trust your maths now, as you've clearly demonstrated not understanding the relationship between thrust, Isp, and power. P = Ve*T, so for 790 kN at even just 360 seconds, that's 2.7 GW - so even if your system was magical and could get such high efficiency from transmission to exhaust velocity that we just rounded up to 3 GW, that's $6 bn in transmitter costs. More likely $12 bn with a more reasonable 50%, and even that seems like a high efficiency - and would mean not being able to increase Isp later without a massive thrust hit, unless you set it up for over $25 bn of transmitters to allow thrust to remain high when Isp hits the claimed 800+ seconds.

I assume the $2mn/MW figure is accurate, as you did not correct it, so I don't know what mistake you made to somehow conclude $1.58 bn, as I can't see any obvious pathway to that number, but it's very clearly wrong, and would take hundreds of launches to amortise costs - assuming maintenance on the array is 0. So, no, it is not cheaper, and would require immense cost drops before anyone would consider having a meeting to consider this technology.

(2) Spaceplanes DON'T NEED as high a TWR as rockets to get to orbit. A 62.8 ton spaceplane doesn't need 790 kN of thrust on the runway- it can EASILY get off the ground with just 200 kN of Thrust. That means a third the cost in Microwave Transmitters just on that basis alone- and that's BEFORE you account of the fact that a Thermal Turbojet could probably achieve Thrust/MW ratings many-fold higher than Microwave Thermal Rockets... (if you can merely get *twice* the Thrust/MW with TTJ's, and climb to an altitude where your Microwave Thermal Rocket ISP using LH2 is 720 seconds instead of the 360 or seconds you get at sea-level, then you can reach orbit with less than half the power-requirements again, for only 1/6th the Microwave Transmitter costs of a rocket with the same mass on the launchpad)

They don't need a TWR above 1, no, but every study of HTOL launch vehicles so far has concluded a TWR of around 0.5 minimum on the runway, and when you have a multi-GW maser aimed at your plane, you're probably not going to want to spend lots of time in that state. No matter how hopeful you are of targeting technology, even a little bit off-target is going to cause many problems for what else it hits, and the heating chamber will not be 100% efficient.

Edited February 19, 2015 by Iskierka

Northstar1989 · February 21, 2015

The vehicle is only lighter if it is smaller, which a pure LH2 design would not be. A CH4 mode may be an improvement that gets around size issues.

ABSOLUTELY NOT. You couldn't POSSIBLY be more wrong. A spaceplane, just like a rocket, is designed for VERY HIGH fuel-fraction. The VAST majority of the weight is in fuel. And fuel tanks are much lighter than the fuel they hold inside them. If you used LH2, you would end up with a MUCH lighter spaceplane... If you don't understand that, you don't understand the Rocket Equation, and there's little point in us even talking further...

Potentially a rocket-only microwave power engine would be lighter, though I personally still find it dubious that it could outperform pressure-fed liquid motors, but any air-breathing thruster has to add in at minimum the vast majority of the mass from any other air-breathing engine, and cannot be claimed to be "much" less.

No, it *IS* much less. Because you get a *VASTLY* superior Thrust for the same engine-size, require no fuel for the atmospheric portion of the flight, and thus can get away with a significantly smaller engine.

See my current attempts to implement realistic total Thrust values for the KSP-Interstellar Thermal Turbojets, where, by applying realistic equations about the relationship between thermal energy and exhaust velocity, I find that at an Exhaust Velocity of about 1333.78 m/s (which is INCREDIBLY slow for any type of Thermal Rocket, but very fast for a jet engine) the Thrust/MW reaches over 22 kN/MW at ideal operating conditions, and 7.24 kn/MW at sea-level.

I don't know how this compares in terms of kN/MW to a chemical jet engine, but I *DO* know this: it is perfectly reasonable for a Thermal Rocket at that Exhaust Velocity, and it matches well to predictions for Thermal Turbojets.

Considering a 1/4 scale version of the Timberwind 75 Nuclear Thermal Rocket could achieve 187.5 MW of ThermalPower (and the Timberwind 75, with 163.7376% more cross-sectional area and 4x the total size, could produce a solid 750 MW), you can easily end up with OVER 1200 kN of Thrust at sea-level with just 187.5 MW of ThermalPower using a 3000 K heat-source.

However, although Microwave Thermal Turbojets would likely operate at SIGNIFICANTLY lower power-levels on take-off, their heat-source would also be much cooler- around 2000 K instead of 3000 K. Why does this matter? It means you get EVEN HIGHER Thrust/MW, at the expense of Exhaust Velocity.

A 2000K heat-source could easily mean only 2/3rd the Exhaust Velocity (I'd have to run some calculations to determine the actual number), and thus more than 1/3rd better Thrust/MW. Which means that sea-level Thrust/MW of more than 9.6 kN/MW, and optimal Thrust/MW of more than 28 kN/MW could *EASILY* be achievable.

Why all this talk of Thrust/MW? Because, with a tiny, lightweight Microwave Receiver (an appropriate-sized receiver would not weigh more than 200 kg AT MOST) that is having only a paltry 18 MW of Microwave Power beamed at it (EASILY achievable, considering a Microwave-Powered launch vehicle would operate in the range of AT LEAST 100 MW for the main transmitter array), you could achieve a Thrust of over 172 kN at sea-level, which is better than the 151 kN maximum Thrust of the J58 engines that powered the SR-71 Blackbird.

And while an SR-71 carried 38.9 metric tons of fuel just for its relatively short bouts of hypersonic flight, a pair of 18 MW Thermal Turbojets could achieve better Thrust than this at SEA-LEVEL (and more than 3 times this Thrust at optimal flight-conditions) for no more than the weight of the engines (which would be LIGHTER than the J58 engines- which weighed 2700 kg each) and intake system (compressor weight is part of the 2700 kg figure for the J58 engines).

NOW are you starting to get the picture of why the performance allowed by Microwave Thermal Turbojets is so great?

There actually isn't as big a difference as you think, at least for equal number of spools and comparable purpose. For low efficiency, high-thrust engines, two-spool high-bypass are typically 8-10, low-bypass 10-12 or so.

I'd call that a HUGE difference. That's a more than 20% difference in TWR!

Single-spools of both classes are the ones that achieve 12-14 ranges, though are rarely used due to the lower specific thrust. Triple-spools that get the required specific thrust only appear in high-bypass engines, which is why you can see a noticeable difference of high BPR engines sitting around 6. The highest ever achieved was a single-spool turbojet designed to completely sacrifice all efficiency and endurance considerations, as its design purpose was a short-duration lift engine for VTOL aircraft, achieving TWR of 18.75. Do not expect to easily match that performance, though, with anything that has a design consideration other than "lots of thrust at low speed." The RB.162 was a true exception - nothing else exceeds around 14.5, because nothing else is so highly specialised towards high TWR.

One of the things you have to realize about Thermal Turbojets is that they don't suffer the same trade-offs as chemical jet engines.

A chemical jet engine, for instance, has to slow its exhaust down to certain speed and pressure conditions in order to effect combustion. Not so with a Thermal Turbojet. Because there is no combustion reaction that occurs, the engine design is MUCH more tolerant of higher speeds and lower pressures after the compressor. Which means one of THE MOST DIFFICULT aspects of engineering a high-performance jet engine (the Compressor) becomes a FRACTION of the difficulty when engineering a Thermal Turbojet...

The other extremely difficult thing to engineer about a chemical jet engines is having a stable mixture-ratio of fuel to air. Too much of either, and the combustion reaction dies out. Not so with a Thermal Turbojet. Because there is NO combustion reaction, there is ABSOLUTELY NO CONCERN about staying within this envelope. Which means,, once again, that a Thermal Turbojet can be designed within a MUCH wider range of tolerances as to internal operating conditions...

It is -also- the most power demanding, as you do not have the airflow available to generate thrust with minimal power - your ideal power/thrust comes once airborne, and the craft's own velocity is gathering air for you.

How many times do I have to say this. A Microwave Thermal Spaceplane does NOT have a difficult time getting airborne.

Not only are its Thermal Turbojet engines MUCH better-performing than chemical jet engines- both in terms of TWR and maximum Thrust (with nuclear-reactor like power-levels, it's possible to get THOUSANDS of kN of Thrust with sufficient available airflow), which means if ANYTHING should be able to get airborne, it's the Microwave Thermal Spaceplane; the aircraft also requires MUCH less fuel-mass than any existing chemical spaceplane design (such as Skylon), which means a lower wing-loading, which means an easier take-off...

Having larger and larger intakes is going to dramatically increase your mass, and no, you actually do not need large intakes at all at speed and altitude.

No, you *DO* need large intakes with a Thermal Turbojet. Because the Thrust potential is so enormous (we're talking THOUSANDS of kN of Thrust once you turn on the high-powered transmitter array down-range and have a couple hundred MW of ThermalPower available...) you need an enormous airflow in order to produce it. And, with an ideal theoretical spaceplane design, you want as high a ratio on Intake Area to total drag as possible- the only reason chemical spaceplane designs shrug off so much potential airflow is because they aren't able to USE all that extra airflow, unlike a Thermal Turbojet (with a chemical jet engine, adding too much air will kill the combustion reaction- NOT an issue AT ALL with a Thermal Turbojet...)

Every single high-speed supersonic aircraft to date has actually had intakes designed to reduce in area at high velocity - Concorde and the SR-71 had spill doors that made more air available at a standstill, then were used to jettison excess at speed. Skylon's intakes are specifically designed to simply close off most of its area at speed. One of the biggest problems with designing high-speed aircraft to be efficient at low speed is that having the required area at low speed means huge intake drag at high speed.

These jet designs had to close off some of their intake area because:

(1) Their compressors weren't capable of handling such a large airflow (NOT an issue with a Microwave Thermal Turbojet aircraft, where a *FAR* lower compression-ratio is necessary, and you have a lot more mass to devote to bigger compressors thanks to your massive fuel-savings).

(2) Too much airflow would have snuffled out their combustion reactions. Once again, NOT an issue with Thermal Turbojets.

The optimal design criteria change *GREATLY* with a Thermal Turbojet vs. a chemical jet engine. Instead of trying to keep internal conditions within a narrow range of airflows, pressures, and speeds so as to create a sustained combustion reaction, with a Thermal Turbojet you're simply trying to shove as much airflow through as quickly as possible, so as to give the Heat Exchanges the chance to dissipate very high power-levels.

And remember what I said before about possibly having a SABRE-like fuel mode with a Thermal Turbojet aircraft at high speed/altitude? Forget it. It was a stupid idea- the more I learn about the greatest difficulties of high-performance jet engine design, the more I realize that the advantages of not having to worry about a combustion reaction GREATLY outweigh any extra Thermal Power you might be able to squeeze out of the engine this way...

The key mis-assumption here is that the internal fuel can be used as a heat-sink. It can't - Skylon is not designed to store any fuel that has been used as coolant, 100% of it must be consumed, as even a few dozen tons of LH2 cannot absorb that much heat without vaporising. And to actually provide sufficient cooling, MORE LH2 than the core engine can consume is required - hence the SABRE's surrounding array of ramjets, as it's genuinely not possible to simply re-store any of the fuel, so it is used in whatever useful manner it can be. If you're going to try use the LH2 as coolant, then you're already going to use all of it as fuel, and once you're getting all the propulsion you can out of fuel energy, the microwave power is useless as it cannot heat the exhaust any further (material thermal limits; the fuel could already do better than it does if not for these) and is simply excess weight on a lifting ascent stage.

You're right about not being able to re-use the LH2, and why the SABRE uses it in a series of surrounding Ramjets. I forgot about that before (yes, I DID previously know about that). But, there is absolutely no reason that a Microwave Thermal Spaceplane can't do the same exact thing- using LH2 as coolant, and then burn it in a series of surrounding ramjets entirely separate from the main (Thermal Turbojet) engine assembly. Thanks for reminding me of the SABRE ramjets, because otherwise I wouldn't have considered that very simply solution to a very obnoxious problem (pre-cooling the air intakes at hypersonic speeds).

This might be a valid point if combustion chambers were actually a significant portion of the mass of an air-breathing engine; they're not, and just moments ago you were suggesting having BOTH, so you cannot claim any significant air-breathing TWR advantage.

Don't mis-construe my ideas. You know, it's getting rather obnoxious that you keep doing that? I don't like it, and would ask that you stop- I'm not doing it to you.

The point of my statement was VERY clear- you said Microwave Thermal Receivers are HEAVY, and I said they're not- in fact, they're lighter than combustion chambers. Just because I said they're lighter than combustion chambers doesn't mean I think combustion chambers are heavy- only that Microwave Thermal Receivers AREN'T- and that they're lighter than the comparable component they replace.

As for the superior TWR, I didn't explain the math/reasoning behind that very well. I've done a MUCH better job of explaining it here. It's NOT that Thermal Turbojets are necessarily LIGHTER than chemical jet engines for their size/volume (they are also quite heavy), it's that they produce MUCH MORE THRUST for their size/volume- which means you get a higher TWR, and are COMPARATIVELY very light for their performance. Did I do a better job of explaining it here?

Do you get it now? Because a Thermal Turbojet is capable of INCREDIBLE Thrust performance (not having to worry about maintaining a combustion reaction, it can work with VERY high airflows, and thus attains the advantages of getting an ENORMOUS working-mass to accelerate) it doesn't have to be as large as a chemical jet to provide the same Thrust. Which means it won't be as heavy- simply because it's smaller. A 1 meter jet engine is going to weigh a LOT less, generally speaking, and a 2 meter jet engine (just to use some made-up diameters...)

Compression-heating on the external parts of a vehicle is not so significant, but stagnation temperature rises very unfavourably to the point of being problematic before even Mach 3 - there's a reason most of the J-58's high speed thrust did not come from its actual turbine section, and specific thrust is quite bad from ramjets. The intake airstream does not have to be that hot, because the compressor does plenty of additional heating anyway, very rapidly bringing intake air to unworkable temperatures.

This is why the fact that a Thermal Turbojet doesn't require high compression-factors (or in fact much compression-factor at all) works GREATLY to its advantage... Less load on the compressor means less heating, which means more favorable stagnation temperatures up into very high speeds...

As I said before, though, the point of a Thermal Turbojet is *NOT* to gain a significant fraction of orbital velocity with air-breathing engines. It *MIGHT* make it up to Mach 4 or 5 in air-breathing mode, depending on exactly how much work you want to put into making the design work well at higher speeds. After that, you're going to want to switch over to Thermal Rocket propulsion. The beauty of the wide range of Thermal Turbojet internal operating conditions that you can design to is that it should be comparably EASY to switch into a rocket-mode using the same engine, unlike the difficulty of engineering something like SABRE. However even if you CAN'T do that, a new Thermal Receiver and rocket-nozzle doesn't weigh that much anyways (maybe a few hundred kg at most)- the VAST majority of the weight in a Microwave Thermal Rocket is in the turbopump- which is an extremely heavy component you're NOT going to need on the Thermal Turbojet anyways... (and one of the reasons a Thermal Turbojet *IS* somewhat lighter than a chemical jet engine of the same size, even if I agree, you ARE going to need a lot of the other heavy components like compressors...)

And again, you do not have any heat-sink on-board. You have a consumable coolant,

Which is therefore a heat-sink. You're talking yourself in circles again. :sticktongue:

IF you use it as fuel, and if you do, then the microwave part of this propulsion stage becomes pointless as it cannot add any useful quantity of energy.

You're joking right? No seriously, tell me you're joking. A Microwave Thermal Receiver heats up to 2000 Kelvin or more. Liquid Hydrogen used as coolant is NOT going to get *NEARLY* that hot before you eject it through a Thermal Rocket (if at the point in your flight where you're running rockets and jets side-by-side), or earlier in the flight through a secondary ramjet (like with SABRE).

If this stage is to be microwave-powered, it's very limited in how fast it can go.

No, it's NOT. Microwaves don't just power Thermal Turbojets- they power ROCKETS too- as I've REPEATEDLY made a point of saying. Now if you're talking about the Thermal Turbojets, it's true that they won't operate into very high speeds- they're going to start losing Thrust fast after Mach 3 (although their Thrust before this point will be SO HIGH that they can lose quite a LOT of Thrust before it starts to become a problem), and are going to probably require concurrent use of rockets after about Mach 4. By Mach 5, the air-breathing components are nothing more than deadweight- but by this point you could have climbed to a VERY HIGH altitude thanks to the low wing-loading, and extreme endurance of the craft when under Thermal Turbojet power, and will be able to acquire a MUCH higher Thrust/MW and ISP from the Thermal Rockets than if you had fired them off at sea-level...

Which, interestingly, you can achieve much more cheaply, and just as effectively, with a couple of quite small solid boosters or similar high-thrust first stage kick.

You talk about how much deadweight a Thermal Turbojet is going to pose (2-4 metric tons per turbojet if their weight is even remotely similar to that of the J58- which weighed 2700 kg- and in fact due to the lack of a turbopump and the need for a less powerful compressor for their airflow rate, they could be a good bit lighter...) and then you propose a SOLID ROCKET BOOSTER? :huh: I just don't know what to do with you... ;.;

The point of a spaceplane isn't just the airbreathing engines. In fact, with a Microwave Thermal Spaceplane, that's not even the primary concern. Whereas with a chemical spaceplane, the jets are the primary advantage, and the wings are just a secondary benefit; with a Microwave Thermal Spaceplane, the WINGS are the primary advantage, and the Thermal Turbojets are just a secondary benefit...

Why? Because Lift/Drag.

Anything, and I mean *ANYTHING* you can do to reduce the amount of Microwave Beamed Power your launch vehicle is going to need to stay in the air, and ultimately climb, is going to pay off MASSIVELY in the long run. This is because the BIGGEST, BADDEST, MOST EXPENSIVE component of a Microwave Thermal Spaceplane isn't in the plane at all- it is in the Microwave Transmitters on the ground.

The simple fact that you can not only continue climbing, but actually handily GAIN speed and altitude with a TWR of say, 0.8 or 0.9 on a Microwave Thermal Spaceplane under rocket-propulsion is a *HUGE* advantage. This works at high speeds/altitudes too: even if your Lift/Drag is only 1.5 or 1.2 due to very high speed, it's still saving you an *ENORMOUS* cost in Microwave Transmitters in the long run...

In fact, the optimal wing-loading on a Microwave Thermal Spaceplane probably is much less than on a chemical spaceplane from a cost-perspective, because you're not trying to minimize fuel-consumption or reduce the size or cost of the plane itself- you're trying to get as heavy a payload to orbit as possible for as little Microwave Beamed Power as possible. Almost NOTHING, and I repeat NOTHING else really matters in the final cost-analysis of a Microwave Thermal Spaceplane other than the number of ground-based Microwave Transmitters you need to build to make it fly and allow it to reach orbit...

Skylon actually intends for its air-breathing stage to be useful, bringing the dV to orbit down from 9800+ to around 6700, which allows a mass ratio of less than 5 for its rocket stage, most of which is in a very dense oxidizer, keeping the vehicle (relatively) small.

A Microwave Thermal Spaceplane doesn't have to worry about Delta-V to *NEARLY* the same degree as a chemical spaceplane does. If you look at the Rocket Equation, a mass-ratio of just 3.4 will give you MORE than 10 km/s of Delta-V with a Hydrogen-propelled Microwave Thermal Spaceplane. And, as I pointed out before, you can use "stages" of heavier to lighter fuel-modes to help bring down the total vessel-size if you want, although that will lead to a significantly higher required mass-fraction (but a net DECREASE in the size and dry-mass of the spaceplane...)

If you just want to get past dense air, which is all a microwave lower stage could reasonably do,

Either you're blissfully unaware of the subject you're talking about, or you're just *trying* to be a troll at this point. :mad:

How many times have I stated the EXCEPTIONAL performance of a Microwave Thermal ROCKET. The Thermal Turbojet-propelled stage may exist only to get the rocket-propelled stage up in the air (although a similar thing could be said in all honesty of ANY spaceplane design), but the rocket-propelled stage DRASTICALLY out-performs *ANY* known chemical rocket, both in terms of Thrust-Weight Ratio and Specific Impulse. There can be no if's, and's, or but's about that if you've any sense on how to read numbers whatsoever...

it would be far more practical to just kick it up there with cheap boosters.

Such as? I once spent WEEKS trying to find a cheap booster that could kick a spaceplane designed for actual horizontal flight (rather than something like a Shuttle) to gain speed and altitude under rocket-propulsion to high-altitude, where the spaceplane could then safely detach without plummeting to its doom due to its lack of significant horizontal velocity to not quickly stall out. There is NOT SUCH THING as a cheap booster that could get the spaceplane up to high altitude. The BEST, and most cost-effective solution BY FAR is just to fly the spaceplane up there with Thermal Turbojets before switching over to rockets (which is where the game's really at when it comes to Microwave Thermal Spaceplanes. STOP. FOCUSING. ON THE THERMAL TURBOJETS!)

Hell, use LH2 liquid fuel flyback boosters, so your same fuel manufacturing systems can supply the boosters, which are also totally and rapidly reusable - would still be far more effective than a microwave lifting stage.

No, it wouldn't. A Thermal Turbojet requires no fuel. It can operate up to AT LEAST Mach 3 and 24,000 meters altitude with the kind of wing-loads of a hydrogen-propelled Microwave Thermal Spaceplane without problem. And it *ONLY* weighs a few tons for several THOUSAND kN of Thrust at a couple hundred Megawatts of beamed-power. That is going to be MUCH more cost-effective than any fly-back booster I know of...

I'm sorry, but I can't trust your maths now, as you've clearly demonstrated not understanding the relationship between thrust, Isp, and power. P = Ve*T,

Don't insult me. Especially when you're clearly wrong.

I'm struggling not to call you all sorts of foul names, but your behavior is INCREDIBLE arrogant. ThermalPower does *NOT* always equal Exhaust Velocity * Thrust. It should be *painfully* obvious to anyone with the most BASIC understanding of physics that as E = 1/2 * m * v^2 if you increase your Exhaust Velocity by a factor of 2, and keep Mass Flow Rate constant, you will increase your Thermal Power Requirements by 4, for instance.

The RELEVANT equations are:

Energy = 1/2 mass * velocity^2

and

Thrust = Mass Flow Rate * Exhaust Velocity

From these two simple equations, you can calculate 90% of what you need to know when it comes to Microwave Thermal Rocketry. There are other equations too, but I'm going to leave those aside for now, because you clearly don't wish to try and understand the actual numbers or math behind any of this, and I've gone on for quite long enough to make my eyes burn just staring at this computer screen...

so for 790 kN at even just 360 seconds, that's 2.7 GW - so even if your system was magical and could get such high efficiency from transmission to exhaust velocity that we just rounded up to 3 GW, that's $6 bn in transmitter costs.

Your numbers are wrong. The $1.58 billion figure (approximately 790 MW in Microwave Beamed Power) stands.

That figure comes from nowhere. As I *REPEATEDLY* pointed out, the Timberwind Nuclear Thermal Rocket designs were supposed to be able to achieve just short of 1 kN/MW of ThermalPower at a Specific Impulse of 1000 seconds and a core temperature of 3000 Kelvin. At a heat exchanger temperature of a bit over 2000 Kelvin and a Specific Impulse of 850 seconds, a Microwave Thermal Rocket gets *MUCH* higher Thrust/MW than that... But I assume the rest of the improvement over 1 kn/MW is lost to transmission inefficiency and such- is where the approximation of 1 kN.MW comes from...

More likely $12 bn with a more reasonable 50%, and even that seems like a high efficiency - and would mean not being able to increase Isp later without a massive thrust hit, unless you set it up for over $25 bn of transmitters to allow thrust to remain high when Isp hits the claimed 800+ seconds.

Your numbers are just ridiculous. They are COMPLETELY out of line with known real-world designs. They are out of line with the mathematical equations. There are even out of line with common sense. And you do *NOTHING* to back them- you just state them as a matter of fact.

I have math to back me. I have real designs to back me. What do you have?

I assume the $2mn/MW figure is accurate, as you did not correct it, so I don't know what mistake you made to somehow conclude $1.58 bn, as I can't see any obvious pathway to that number, but it's very clearly wrong, and would take hundreds of launches to amortise costs - assuming maintenance on the array is 0. So, no, it is not cheaper, and would require immense cost drops before anyone would consider having a meeting to consider this technology.

The MATH backs $1.58 billion. The numbers are $2 million/ MW, and 1 kN/ MW. That comes out to 790 MW of power, which means $1.58 billion in transmitters. How much simpler can I make this for you? You question the figure of 1 kN/MW, but you provide absolutely no reasoning to back the assertion that it is wrong, whereas I have MANY real-world examples of Nuclear Thermal Rocket designs that back my figure (including NERVA, the Russian analog, Timberwind, the US Air Force SNTP Program, and the joint US/Russian Bimodal Nuclear Thermal Propulsion Program- ALL of which can be used to support a figure of approximately 1 kN/MW at 1000 seconds of Specific Impulse using Hydrogen...)

For instance, Project Timberwind designs, with a much higher exhaust temperature of 3000 Kelvin, and an ISP of 1000 seconds (both these numbers are important, because they help provide an idea how much Thrust is coming from the nozzle's Expansion Ratio- which was not very much in the case of Timberwind, thus leaving SIGNIFICANT room for improvement with a Microwave Thermal Rocket using a larger nozzle, if you lifted it up above the thickest part of the atmosphere with a Thermal Turbojet...) were expected to achieve a Thrust/MW of approximately 1 kN/MW.

If you can't accept the hard physical facts of reality, there's no point in talking further.

Support your assertions with math, and *ACTUALLY KNOW* your equations.

They don't need a TWR above 1, no, but every study of HTOL launch vehicles so far has concluded a TWR of around 0.5 minimum on the runway, and when you have a multi-GW maser aimed at your plane, you're probably not going to want to spend lots of time in that state. No matter how hopeful you are of targeting technology, even a little bit off-target is going to cause many problems for what else it hits, and the heating chamber will not be 100% efficient.

Multi-GW? Since *WHEN* did 790 MW for a Thermal Rocket (which only gets 1 kN/MW) to get a TWR of 1.28 turn into multiple GIGAWATTS for a Thermal Turbojet (which gets a Thrust/MW of *at least* 7.2 kN/MW, and likely about 9.6 kN/MW of the runway) for a spaceplane (of undetermined mass) to get a TWR of 0.5 on the runway?

Oh, that's right. You inflated your figure for the rocket to 3 GW (which is COMPLETELY inaccurate). That's STILL not even a single GW on the runway- if you divide 3 GW by 7.2, and then again by 2.56 (1.28/0.5) you are only going to require about 163 MW- and once again that figure of 3 GW is *WILDLY* inaccurate, and the CORRECT figure is only 790 MW for the rocket (and only 43 MW for a 62.8 ton spaceplane to get off the runway- although at THAT power-level, it won't be reaching orbit with a considerable-sized cargo anytime soon...)

Regards,

Northstar

Edited February 21, 2015 by Northstar1989

shynung · February 21, 2015

ABSOLUTELY NOT. You couldn't POSSIBLY be more wrong. A spaceplane, just like a rocket, is designed for VERY HIGH fuel-fraction. The VAST majority of the weight is in fuel. And fuel tanks are much lighter than the fuel they hold inside them. If you used LH2, you would end up with a MUCH lighter spaceplane... If you don't understand that, you don't understand the Rocket Equation, and there's little point in us even talking further...

1 cubic meter of LH2 only weighs 70 kg. A typical fuel oil of the same volume weighs almost a ton in comparison. Any vehicle carrying any significant amount of LH2 will have very large tanks, which, along with all the insulation to keep the hydrogen liquid, will invariably be heavy in comparison to a tank carrying the same mass of kerosene. Tankage weight and drag penalties (due to its size) will be significant, despite the propellant being lightweight.

Remember the Space Shuttle External Tank? Around 3/4 of the volume inside is liquid hydrogen, yet it only represents about 16% of the total propellant mass, the rest being taken up by liquid oxygen.

Edited February 22, 2015 by shynung
wrong number

Iskierka · February 21, 2015

ABSOLUTELY NOT. You couldn't POSSIBLY be more wrong. A spaceplane, just like a rocket, is designed for VERY HIGH fuel-fraction. The VAST majority of the weight is in fuel. And fuel tanks are much lighter than the fuel they hold inside them. If you used LH2, you would end up with a MUCH lighter spaceplane... If you don't understand that, you don't understand the Rocket Equation, and there's little point in us even talking further...

No, you actually wouldn't, not because of the rocket equation, but because of engineering. Yes, most rocket tanks are light, about 1% of the mass of internal fuel. LH2 tanks are not - they are about 10% of the fuel's internal mass. Your vehicle is bigger and uses a fuel that requires a far worse structure - you can assume around equal mass, maybe marginally less, not considerably less.

No, it *IS* much less. Because you get a *VASTLY* superior Thrust for the same engine-size, require no fuel for the atmospheric portion of the flight, and thus can get away with a significantly smaller engine.
And while an SR-71 carried 38.9 metric tons of fuel just for its relatively short bouts of hypersonic flight, a pair of 18 MW Thermal Turbojets could achieve better Thrust than this at SEA-LEVEL (and more than 3 times this Thrust at optimal flight-conditions) for no more than the weight of the engines (which would be LIGHTER than the J58 engines- which weighed 2700 kg each) and intake system (compressor weight is part of the 2700 kg figure for the J58 engines).
NOW are you starting to get the picture of why the performance allowed by Microwave Thermal Turbojets is so great?

No, because you have stated nothing to explain why mass would be reduced. Power is irrelevant except as a potential factor to increase mass due to a receiver component needing to be large enough to function - I'm giving benefit of the doubt that this is not the case, so waiting for all the other components that would be shared to be explained away.

One of the things you have to realize about Thermal Turbojets is that they don't suffer the same trade-offs as chemical jet engines.
A chemical jet engine, for instance, has to slow its exhaust down to certain speed and pressure conditions in order to effect combustion. Not so with a Thermal Turbojet. Because there is no combustion reaction that occurs, the engine design is MUCH more tolerant of higher speeds and lower pressures after the compressor. Which means one of THE MOST DIFFICULT aspects of engineering a high-performance jet engine (the Compressor) becomes a FRACTION of the difficulty when engineering a Thermal Turbojet...

Uh, no, chemical jets don't slow their combustion down much at all - velocity in the combustion chamber is above Mach 0.5. The combustor also does not care about pressure - it simply burns things, so as long as you send it through at a sensible velocity for mixing, the air can be any pressure. Pressure is required to turn the compressor, and to give the engine any propulsive capability whatsoever - both things that a microwave design would require, so you can't claim this is not a concern. Additionally, engineering difficulty is not proportional to required mass addition - one can be easier but no lighter.

How many times do I have to say this. A Microwave Thermal Spaceplane does NOT have a difficult time getting airborne.
Not only are its Thermal Turbojet engines MUCH better-performing than chemical jet engines- both in terms of TWR and maximum Thrust (with nuclear-reactor like power-levels, it's possible to get THOUSANDS of kN of Thrust with sufficient available airflow), which means if ANYTHING should be able to get airborne, it's the Microwave Thermal Spaceplane; the aircraft also requires MUCH less fuel-mass than any existing chemical spaceplane design (such as Skylon), which means a lower wing-loading, which means an easier take-off...

TWR unproven, thrust is dependent on airflow, which is, again, very limited when static. I'm sure it can get airborne - but then, so can a brick with black powder detonated beneath it. The challenge is getting airborne effectively.

No, you *DO* need large intakes with a Thermal Turbojet. Because the Thrust potential is so enormous (we're talking THOUSANDS of kN of Thrust once you turn on the high-powered transmitter array down-range and have a couple hundred MW of ThermalPower available...) you need an enormous airflow in order to produce it. And, with an ideal theoretical spaceplane design, you want as high a ratio on Intake Area to total drag as possible- the only reason chemical spaceplane designs shrug off so much potential airflow is because they aren't able to USE all that extra airflow, unlike a Thermal Turbojet (with a chemical jet engine, adding too much air will kill the combustion reaction- NOT an issue AT ALL with a Thermal Turbojet...)
These jet designs had to close off some of their intake area because:
(1) Their compressors weren't capable of handling such a large airflow (NOT an issue with a Microwave Thermal Turbojet aircraft, where a *FAR* lower compression-ratio is necessary, and you have a lot more mass to devote to bigger compressors thanks to your massive fuel-savings).
(2) Too much airflow would have snuffled out their combustion reactions. Once again, NOT an issue with Thermal Turbojets.
The optimal design criteria change *GREATLY* with a Thermal Turbojet vs. a chemical jet engine. Instead of trying to keep internal conditions within a narrow range of airflows, pressures, and speeds so as to create a sustained combustion reaction, with a Thermal Turbojet you're simply trying to shove as much airflow through as quickly as possible, so as to give the Heat Exchanges the chance to dissipate very high power-levels.

This is not an issue of sheer size or such - fundamentally, you can only stuff so much air down a certain area, and as you accelerate, for a certain intake area, more air is trying to come in. This is about how a chosen size performs at different speeds - you can either design for high thrust at static by having a large intake that can provide enough air, then have to dump most of it at speed as it simply does not fit into the engine, regardless of compression ratio, or you can have tiny intakes that provide very little thrust at static, but allow the engine to operate properly at speed. Or you can take the solution actually demonstrated, have a limited ability to open the intakes in both directions, and get slightly better static thrust at a small sacrifice of intake drag at speed. It's only barely even physics - it's geometry of "you can only push the air through at this maximum v(compressor), for this area, and the conditions get less favourable as v(aircraft) increases".

Don't mis-construe my ideas. You know, it's getting rather obnoxious that you keep doing that? I don't like it, and would ask that you stop- I'm not doing it to you.
The point of my statement was VERY clear- you said Microwave Thermal Receivers are HEAVY, and I said they're not- in fact, they're lighter than combustion chambers. Just because I said they're lighter than combustion chambers doesn't mean I think combustion chambers are heavy- only that Microwave Thermal Receivers AREN'T- and that they're lighter than the comparable component they replace.
As for the superior TWR, I didn't explain the math/reasoning behind that very well. I've done a MUCH better job of explaining it here. It's NOT that Thermal Turbojets are necessarily LIGHTER than chemical jet engines for their size/volume (they are also quite heavy), it's that they produce MUCH MORE THRUST for their size/volume- which means you get a higher TWR, and are COMPARATIVELY very light for their performance. Did I do a better job of explaining it here?
Do you get it now? Because a Thermal Turbojet is capable of INCREDIBLE Thrust performance (not having to worry about maintaining a combustion reaction, it can work with VERY high airflows, and thus attains the advantages of getting an ENORMOUS working-mass to accelerate) it doesn't have to be as large as a chemical jet to provide the same Thrust. Which means it won't be as heavy- simply because it's smaller. A 1 meter jet engine is going to weigh a LOT less, generally speaking, and a 2 meter jet engine (just to use some made-up diameters...)

I didn't mis-construe anything - you plainly suggested hybrid combustion/thermal heating, and only just reverted on that idea. Speaking of, while I personally question whether they would be so light, I have not been stating that the receivers would be heavy - I have been pointing out all the OTHER masses that are being dismissed. You claimed TWR of 15-20 reasonable, when even the best jets that are not hyper-specialised peak at 14, and all of these are low-velocity engines. You also won't get much better thrust for the same size - power is irrelevant, as you've confirmed heating temperatures of around 2000K - equal to what hydrocarbon jet engines achieve, with no considerable other advantages, as you do have the same mass-flow limits. For any comparable design purpose, a microwave thermal jet will not notably out-perform a gas turbine jet, as it experiences all the same performance-limiting design considerations, even if there are potentially less design considerations that require significant effort to get the correct solution.

However even if you CAN'T do that, a new Thermal Receiver and rocket-nozzle doesn't weigh that much anyways (maybe a few hundred kg at most)- the VAST majority of the weight in a Microwave Thermal Rocket is in the turbopump- which is an extremely heavy component you're NOT going to need on the Thermal Turbojet anyways... (and one of the reasons a Thermal Turbojet *IS* somewhat lighter than a chemical jet engine of the same size, even if I agree, you ARE going to need a lot of the other heavy components like compressors...)

And that turbopump is going to be huge, having to deal with LH2 for all of the propellant. From the wiki article nuclear thermal rockets, we find that a J-2 thrust spec engine would need turbopumps almost as large as the F-1's. As a side note, the fuel pump is not a heavy component in a jet engine, nor is it a turbopump - where rockets are dealing with hundreds to thousands of litres per second, even the most powerful turbofans peak at less than 5.

You're joking right? No seriously, tell me you're joking. A Microwave Thermal Receiver heats up to 2000 Kelvin or more. Liquid Hydrogen used as coolant is NOT going to get *NEARLY* that hot before you eject it through a Thermal Rocket (if at the point in your flight where you're running rockets and jets side-by-side), or earlier in the flight through a secondary ramjet (like with SABRE).
No, it's NOT. Microwaves don't just power Thermal Turbojets- they power ROCKETS too- as I've REPEATEDLY made a point of saying. Now if you're talking about the Thermal Turbojets, it's true that they won't operate into very high speeds- they're going to start losing Thrust fast after Mach 3 (although their Thrust before this point will be SO HIGH that they can lose quite a LOT of Thrust before it starts to become a problem), and are going to probably require concurrent use of rockets after about Mach 4.

It's going to get hotter than that if used as a fuel, which is what was being discussed. In that consideration, yes, as stated, the microwave section is useless. And I also specifically stated "this stage"; ie, the airbreathing phase of flight.

You talk about how much deadweight a Thermal Turbojet is going to pose (2-4 metric tons per turbojet if their weight is even remotely similar to that of the J58- which weighed 2700 kg- and in fact due to the lack of a turbopump and the need for a less powerful compressor for their airflow rate, they could be a good bit lighter...) and then you propose a SOLID ROCKET BOOSTER? I just don't know what to do with you...

An SRB is phenomenal thrust at minimal expense, and dropped when expended and forgotten about. That's 0 dead weight. You were saying?

The simple fact that you can not only continue climbing, but actually handily GAIN speed and altitude with a TWR of say, 0.8 or 0.9 on a Microwave Thermal Spaceplane under rocket-propulsion is a *HUGE* advantage. This works at high speeds/altitudes too: even if your Lift/Drag is only 1.5 or 1.2 due to very high speed, it's still saving you an *ENORMOUS* cost in Microwave Transmitters in the long run...
In fact, the optimal wing-loading on a Microwave Thermal Spaceplane probably is much less than on a chemical spaceplane from a cost-perspective, because you're not trying to minimize fuel-consumption or reduce the size or cost of the plane itself- you're trying to get as heavy a payload to orbit as possible for as little Microwave Beamed Power as possible. Almost NOTHING, and I repeat NOTHING else really matters in the final cost-analysis of a Microwave Thermal Spaceplane other than the number of ground-based Microwave Transmitters you need to build to make it fly and allow it to reach orbit...

The moment you're in rocket mode, you want to be getting the hell to orbit. Any excess fuel expenditure from that point is a waste - and while it does reduce fuel flow, remaining in lifting flight massively increases time, and thus total fuel requirement. Yes, it reduces peak power - but maybe that indicates a problem with this propulsion system. You also need to ascend as fast as possible, as if you try continue climbing flight, you're going to throw yourself straight into excessive heating phases rivalling re-entry, and enduring much longer.

Also, your wing-loading is likely going to cause aerodynamic issues. I've avoided pointing it out, as there's plenty of problems with just getting the propulsion, but at your sheer size you're going to get terrible wetted aspect ratio, so you're really going to struggle for L/D. And it's L/D you're wanting, not simply low wing-loading. I'm sure -a- vehicle is constructable, but it's only ever going to be fairly mediocre aerodynamic performance, at best.

How many times have I stated the EXCEPTIONAL performance of a Microwave Thermal ROCKET. The Thermal Turbojet-propelled stage may exist only to get the rocket-propelled stage up in the air (although a similar thing could be said in all honesty of ANY spaceplane design), but the rocket-propelled stage DRASTICALLY out-performs *ANY* known chemical rocket, both in terms of Thrust-Weight Ratio and Specific Impulse. There can be no if's, and's, or but's about that if you've any sense on how to read numbers whatsoever...

I've yet to see solid numbers that make sense. You haven't explained your colossal but featherweight turbo- or other fuel pump, and we'll address things like power shortly ...

Such as? I once spent WEEKS trying to find a cheap booster that could kick a spaceplane designed for actual horizontal flight (rather than something like a Shuttle) to gain speed and altitude under rocket-propulsion to high-altitude, where the spaceplane could then safely detach without plummeting to its doom due to its lack of significant horizontal velocity to not quickly stall out. There is NOT SUCH THING as a cheap booster that could get the spaceplane up to high altitude. The BEST, and most cost-effective solution BY FAR is just to fly the spaceplane up there with Thermal Turbojets before switching over to rockets (which is where the game's really at when it comes to Microwave Thermal Spaceplanes. STOP. FOCUSING. ON THE THERMAL TURBOJETS!)

Why would you enter lifting flight? Get a big booster, of which there are plenty of examples (SS-SRB, Ariane 5, etc), kick it up to so stupidly high and fast it can't fall down quickly enough to be an issue, then just cruise to orbit on rocket power, as slow as you like. Why keep focussing on the plane part?

I'm struggling not to call you all sorts of foul names, but your behavior is INCREDIBLE arrogant. ThermalPower does *NOT* always equal Exhaust Velocity * Thrust. It should be *painfully* obvious to anyone with the most BASIC understanding of physics that as E = 1/2 * m * v^2 if you increase your Exhaust Velocity by a factor of 2, and keep Mass Flow Rate constant, you will increase your Thermal Power Requirements by 4, for instance.
The RELEVANT equations are:
Energy = 1/2 mass * velocity^2
and
Thrust = Mass Flow Rate * Exhaust Velocity
From these two simple equations, you can calculate 90% of what you need to know when it comes to Microwave Thermal Rocketry. There are other equations too, but I'm going to leave those aside for now, because you clearly don't wish to try and understand the actual numbers or math behind any of this, and I've gone on for quite long enough to make my eyes burn just staring at this computer screen...

Okay, let's do so. E = 1/2 m * v^2, so;

P = 1/2 m(dot) * v * v

T = m(dot) * v

Hang on, I see a substitution here ...

P = 1/2 T * v

Yes, I'll grant I forgot the 1/2 factor - cut all my previous claims in two. That's still far above your estimates, and your own equations have just proven that my equation, with a constant, -is- correct, and thermal power, at absolute minimum, is half what I claimed.

Your numbers are wrong. The $1.58 billion figure (approximately 790 MW in Microwave Beamed Power) stands.
That figure comes from nowhere. As I *REPEATEDLY* pointed out, the Timberwind Nuclear Thermal Rocket designs were supposed to be able to achieve just short of 1 kN/MW of ThermalPower at a Specific Impulse of 1000 seconds and a core temperature of 3000 Kelvin. At a heat exchanger temperature of a bit over 2000 Kelvin and a Specific Impulse of 850 seconds, a Microwave Thermal Rocket gets *MUCH* higher Thrust/MW than that... But I assume the rest of the improvement over 1 kn/MW is lost to transmission inefficiency and such- is where the approximation of 1 kN.MW comes from...
Your numbers are just ridiculous. They are COMPLETELY out of line with known real-world designs. They are out of line with the mathematical equations. There are even out of line with common sense. And you do *NOTHING* to back them- you just state them as a matter of fact.
I have math to back me. I have real designs to back me. What do you have?

No, you actually don't. Please, point out where the power of Timberwind is claimed? I can find no references to such on its wiki article, or on most of the external links and references - the rest I did not check as they didn't look like ones that would address it. Similarly, such information is also missing from NERVA's page and references. I -did- eventually find a reference to what T/P is possible - on the previously mentioned nuclear thermal rocket article, in a different section, one of the NERVA models gets it performance confirmed. It produced 334 kN - at 1100 MW. So, about a third of your claim. It's also notable that nowhere is it stated that this thrust was achieved at maximum Isp, which I'll address in a moment.

If we return to the section we were originally looking at in the NTR article, we find that it actually performs calculations of T/P for various engines. Interestingly, in all cases, the same P = T*Ve/2 equation is used. It concludes that at 850 seconds, you would be capable of achieving less than 0.25 kN/MW, at perfect efficiency. This is the fundamental limit by ratio of velocity to energy. Even at the J-2's measly 414 seconds, the performance ratio is less than 0.5 kN/MW. We can see that the NERVA's 334kN/1100MW is between these two - so we must conclude that it was operating at reduced Ve, to gain the thrust necessary for whatever phase of flight demanded it.

In fact, the equation can be rearranged, and we find that to achieve the claimed 1 kN/MW, your exhaust velocity would have to be 2 km/s - or less, if inefficiencies exist. This is amusingly worse than almost every chemical fuel bar cold-gas - though with a dense and cheap working fluid, it might make an effective launch vehicle by allowing exceptionally tiny sizes, and making the power requirements almost reasonable. I guess then you have to balance whether you want to power the crazy thrust for such an awful mass-ratio, or the crazy Isp to actually keep the vehicle light.

You didn't actually use the maths, and it turns out it was against you. You didn't cite how the real-world examples possibly supported your assertions, and now that I've actually done research on them, it turns out, they support the opposite. You will, fundamentally, require very large powers for the rocket phase. And an air-breathing phase is dubious - further supported by the fact that the only people researching microwave power are ignoring the possibility.

The MATH backs $1.58 billion. The numbers are $2 million/ MW, and 1 kN/ MW. That comes out to 790 MW of power, which means $1.58 billion in transmitters. How much simpler can I make this for you? You question the figure of 1 kN/MW, but you provide absolutely no reasoning to back the assertion that it is wrong, whereas I have MANY real-world examples of Nuclear Thermal Rocket designs that back my figure (including NERVA, the Russian analog, Timberwind, the US Air Force SNTP Program, and the joint US/Russian Bimodal Nuclear Thermal Propulsion Program- ALL of which can be used to support a figure of approximately 1 kN/MW at 1000 seconds of Specific Impulse using Hydrogen...)
If you can't accept the hard physical facts of reality, there's no point in talking further.
Support your assertions with math, and *ACTUALLY KNOW* your equations.

Yes, I'm sorry I forgot the 1/2 factor. I made a hundred or so launches required to amortise costs into a couple hundred. Sorry. But the 1 kN/MW figure has just been totally debunked, with none of your examples actually supporting it, hopefully to your satisfaction, and so many of the previous cost estimates stand - and yours remains physically impossible, unless you want to completely rework your design into a tiny, hyperdense vehicle with only 2 km/s Ve.

Multi-GW? Since *WHEN* did 790 MW for a Thermal Rocket (which only gets 1 kN/MW) to get a TWR of 1.28 turn into multiple GIGAWATTS for a Thermal Turbojet (which gets a Thrust/MW of *at least* 7.2 kN/MW, and likely about 9.6 kN/MW of the runway) for a spaceplane (of undetermined mass) to get a TWR of 0.5 on the runway?
Oh, that's right. You inflated your figure for the rocket to 3 GW (which is COMPLETELY inaccurate). That's STILL not even a single GW on the runway- if you divide 3 GW by 7.2, and then again by 2.56 (1.28/0.5) you are only going to require about 163 MW- and once again that figure of 3 GW is *WILDLY* inaccurate, and the CORRECT figure is only 790 MW for the rocket (and only 43 MW for a 62.8 ton spaceplane to get off the runway- although at THAT power-level, it won't be reaching orbit with a considerable-sized cargo anytime soon...)
Regards,
Northstar

As above, multi-GW stands at any actually high Isp. 3 GW is entirely valid at 50% efficiency, which is what rockets themselves typically get just from thermal->kinetic conversion by expansion near sea level. 75+%is seen at very high expansions in very low pressure, but that still leaves you happily around 2 GW - though I'll grant you could probably reduce thrust to maybe 1.5 GW peak power by these conditions. Assuming no transmission loss, of course. Thermal turbojet numbers are highly dubious, as many of your calculations seem to be based on ratios to rocket performance, which has been debunked, and also jet design difficulties, even thermal, mean we should consider ignoring them and focus on the rocket phase - which still doesn't stand up.

EDIT for the post I missed:

1 cubic meter of LH2 only weighs 7 kg. A typical fuel oil of the same volume weighs almost a ton in comparison. Any vehicle carrying any significant amount of LH2 will have very large tanks, which, along with all the insulation to keep the hydrogen liquid, will invariably be heavy in comparison to a tank carrying the same mass of kerosene. Tankage weight and drag penalties (due to its size) will be significant, despite the propellant being lightweight.
Remember the Space Shuttle External Tank? Around 3/4 of the volume inside is liquid hydrogen, yet it only represents about 16% of the total propellant mass, the rest being taken up by liquid oxygen.

LH2 is actually 70 kg/m^3, but the point stands, yes. As mentioned at the start of my post - if you do research on the actual tanks, LOx and RP-1, hypergolic fuels, etc, that are all closer to 1 t/m^3, all have ideal tankage ratios around 1%, but hydrogen is always around 10% at best, due to the volume, extreme insulation required, and funky material requirements due to hydrogen's reactivity - everything else can just use the best material by weight/cost, hydrogen needs the best that it won't eat from the inside.

Edited February 21, 2015 by Iskierka

Northstar1989 · February 21, 2015

OK, first of all, I do honestly believe you're just trying to be a troll. If not by ignoring my arguments and wildly-misrepresenting the actual facts, then by posting long walls of text that say nothing and *intentionally* trying to get my pulse up.

As such, I'm going to answer your post, but I would ask that you refrain from posting anything further longer than about two paragraphs at a time on this thread (I will try to keep my response length the same), so that our discussion does not hijack this thread. IF you can't abide by that, I would respectfully ask that you leave my thread (I started this thread) and never come back.

No, you actually wouldn't, not because of the rocket equation, but because of engineering. Yes, most rocket tanks are light, about 1% of the mass of internal fuel. LH2 tanks are not - they are about 10% of the fuel's internal mass. Your vehicle is bigger and uses a fuel that requires a far worse structure - you can assume around equal mass, maybe marginally less, not considerably less.

Then go with Hydro/LOX, and have a LOX afterburner LANTR-style.

Your numbers for LH2 fuel tank mass fractions are wildly inaccurate, though. Consider the Atlas rockets or Centaur upper stage (can't provide a link tight now, as my internet is being wacky). They had total mass fractions in excess of 95%, with LH2/LOX propulsion the whole way. How would that be possible if the vast majority of their fuel tanks had a mass-fraction of 10%?

No, because you have stated nothing to explain why mass would be reduced.

The need for a lighter compressor. The lack of a need for a fuel-pump? The replacement of a combustion chamber with (relatively much lighter) Microwave Thermal Receivers- even if the combustion chamber is only a small proportion of mass to begin with? I've already listed several ways you save on mass, and you've ignored every one of them- which reaffirms my suspicion you're not looking to have a reasonable discussion.

Power is irrelevant except as a potential factor to increase mass due to a receiver component needing to be large enough to function - I'm giving benefit of the doubt that this is not the case, so waiting for all the other components that would be shared to be explained away.

No, Power is the most important thing, period. You know that. The more power you can get for a given mass, the higher your TWR and the less engine mass you need in the first place. There are two ways to get very high TWR- one is to have very low mass, and the other is to have very high Thrust. The Thermal Turbojet takes mainly the latter approach, even if it is also*slightly* lighter due to the lack of need for a fuel pump or combustion chamber.

Uh, no, chemical jets don't slow their combustion down much at all - velocity in the combustion chamber is above Mach 0.5.

When you're traveling at Mach 3, that's a HUGE slow-down. The data says differently- most jets combust at around Mach 0.4-0.5, which is *not* "above Mach 0.5". A Thermal Turbojet operates well as supersonic flow conditions internally, by contrast. Note I said "turbojet"- you keep equating it with a Turbofan, which it is not. There are no fan blades involved in the design.

The combustor also does not care about pressure - it simply burns things, so as long as you send it through at a sensible velocity for mixing, the air can be any pressure.

Countless engineering studies say differently. The combustion process is sensitive to temperature. It is sensitive to very low pressures. And, it is sensitive to the mixing-ratio between air and fuel, which is the primary reason you have to limit airflow at high speeds. It's not that you CAN'T fit the entire airflow through at high speed/altitude flight (the Venturi Effect guarantees that you can- the airspeed inside the engine can always increase to increase mass flow rate), it's that doing so would move the internal conditions outside the range where combustion can efficiently occur (in particular, the fuel-pump doesn't magically become capable of pumping more fuel just because you have a higher airflow, and you need a certain fuel:air ratio for the combustion process to be stable. Thermal Turbojets avoid this issue by not needing a fuel-pump!)

Pressure is required to turn the compressor, and to give the engine any propulsive capability whatsoever - both things that a microwave design would require, so you can't claim this is not a concern.

Not, it's just *less* of a concern- because a Thermal Turbojet can operate at much higher internal pressures and speeds. Half our problem is that you're comparing Turbojet performance (which works better at high speeds and altitudes) to Turbofan performance (which works poorly at high speeds and altitudes). They are *not* the same thing.

Additionally, engineering difficulty is not proportional to required mass addition - one can be easier but no lighter.

This statement is completely unclear as to what it is referring to.

TWR unproven, thrust is dependent on airflow, which is, again, very limited when static. I'm sure it can get airborne - but then, so can a brick with black powder detonated beneath it. The challenge is getting airborne effectively.

It doesn't need to be effective- it just needs to work. Because Thermal Turbojets are *turbojets* and not *turbofans* they work better at higher speeds and pressures. You only need enough Thrust to get off the runway... It's not an issue, so stop making it out to be one.

This is not an issue of sheer size or such - fundamentally, you can only stuff so much air down a certain area, and as you accelerate, for a certain intake area, more air is trying to come in.

Your understanding of fluid dynamics is, at best, drastically messed up. The airflow can do any of several things as it compresses- it can move faster, it can get hotter, or it can increase in pressure. Normally, it does a combination of all 3 of these things. The increase in temperature, in particular, is a major concern, and the usual limiting-factor on airflow to a chemical jet is the temperature reached by the final blade of the compressor (where it gets hottest).

There is absolutely no concern about being able to fit the full airflow of the intake through the engine- there is a concern about being able to do it without the compressor blades melting. By relying more on ram-compression effects, and having a less powerful compressor to begin with, it's possible to get MUCH higher airflows at high speed/altitude, at the expense of sea-level performance... (which I've already admitted, is comparably low for a Thermal Turbojet compared to its performance at speed/altitude- but still not a significant problem for getting off the runway. Because TWR improves as you ascend, and you're optimizing for minimal power-usage instead of fuel-consumption, you don't need a liftoff TWR of 0.5 on the runway for an optimal ascent- a liftoff TWR of 0.2 is fine if the TWR climbs as your plane does...)

This is about how a chosen size performs at different speeds - you can either design for high thrust at static by having a large intake that can provide enough air, then have to dump most of it at speed as it simply does not fit into the engine,

Once again, the airflow *does* fit into the engine. It's just a matter of doing it without melting the engine. For which, pre-coolers help (and you can dump the vaporized Hydrogen overboard via a small ramjet like with SABRE). It's better to have a larger intake at high altitude if you can effectively utilize all that airflow- which a cehmical jet can't do, but a TTJ can do.

regardless of compression ratio, or you can have tiny intakes that provide very little thrust at static, but allow the engine to operate properly at speed. Or you can take the solution actually demonstrated, have a limited ability to open the intakes in both directions, and get slightly better static thrust at a small sacrifice of intake drag at speed. It's only barely even physics - it's geometry of "you can only push the air through at this maximum v(compressor), for this area, and the conditions get less favourable as v(aircraft) increases".

Your understanding of fluid dynamics, and the reason for having to limit airflow at high speeds is faulty. I won't discuss this more, because you won't listen to me when I tell you why and how you're wrong, and your arguments are scientifically inaccurate.

I didn't mis-construe anything - you plainly suggested hybrid combustion/thermal heating, and only just reverted on that idea.

But that wasn't a part of the discussion, or what I said you were mis-construing, was it?

Speaking of, while I personally question whether they would be so light,

Because they're made of lightweight semiconductors, rather than metal. This is also the reason they only operate at 2000 Kelivin when the beamed-pwoer could easily heat them hotter than this...

I have not been stating that the receivers would be heavy

You did. Do I have to go back and quote you?

- I have been pointing out all the OTHER masses that are being dismissed. You claimed TWR of 15-20 reasonable, when even the best jets that are not hyper-specialised peak at 14, and all of these are low-velocity engines.[/quote[
Once again, you do it by having higher Thrust, not lower mass.
You also won't get much better thrust for the same size - power is irrelevant, as you've confirmed heating temperatures of around 2000K - equal to what hydrocarbon jet engines achieve,
The difference is how much airflow you are heating at that temperature. Jet engines are limited in the airflow they can process, if nothing else, by the size of their fuel-pump (as you've already acknowledged, the J58 fuel-pump was HUGE). A 2000 K thermal receiver could have 10x the airflow or more of a combustion chamber without issue.
with no considerable other advantages, as you do have the same mass-flow limits.
You don't have the same mass-flow limits, as I've repeatedly stated.
For any comparable design purpose, a microwave thermal jet will not notably out-perform a gas turbine jet, as it experiences all the same performance-limiting design considerations, even if there are potentially less design considerations that require significant effort to get the correct solution.
The limiting considerations are *not* the same, and I have to get going irl.
Other parts of your post will be competently addressed later. I don't have time to do them justice right now.
Regards,
Northstar

Edited February 22, 2015 by Northstar1989

KSK · February 21, 2015

Your numbers for LH2 fuel tank mass fractions are wildly inaccurate, though. Consider the Atlas rockets or Centaur upper stage (can't provide a link tight now, as my internet is being wacky). They had total mass fractions in excess of 95%, with LH2/LOX propulsion the whole way. How would that be possible if the vast majority of their fuel tanks had a mass-fraction of 10%?

Atlas V - three RP1/LOX burning Common Core Boosters with an LH2/LOX (Centaur) upper stage. Previous generations of Atlas have also used RP1/LOX. Not LH2/LOX the whole way.

Northstar1989 · February 22, 2015

Atlas V - three RP1/LOX burning Common Core Boosters with an LH2/LOX (Centaur) upper stage. Previous generations of Atlas have also used RP1/LOX. Not LH2/LOX the whole way.

I was thinking of the SM-65 Atlas ICBM rocket, not the Atlas V or earlier Atlas launch-family rockets. Sorry for any confusion this caused...

It turns out the Centaur balloon tanks were quite a bit less impressive that I remembered, though- Centaur only managed 90% ratios (therein lies the 1:10 mass ratio Iskierka mentioned before), although it must be pointed out this was the mass-ratio of the *entire* upper stage...

I do have to apologize though. My Thrust/MW figures were *VERY* off for Thermal Turbojets. Not because I didn't know how to calculate Thrust/MW correctly (it's, in fact, something I've become very skilled at lately) but because of a stupid math error- I accidentally forgot to divide standard gravity ("g", 9.80665 m/s^2) out at one point in my equations. The result was a Thrust/MW value that was 9.80665 times too high... :blush:

The *CORRECT* Thrust/MW for a Thermal Turbojet:

Sea-Level: 0.738 kN/MW

Ideal Speed/Altitude: 2.306 kN/MW

Now THAT seems a lot more reasonable, doesn't it?

Regards,

Northstar

shynung · February 22, 2015

LH2 is actually 70 kg/m^3, but the point stands, yes. As mentioned at the start of my post - if you do research on the actual tanks, LOx and RP-1, hypergolic fuels, etc, that are all closer to 1 t/m^3, all have ideal tankage ratios around 1%, but hydrogen is always around 10% at best, due to the volume, extreme insulation required, and funky material requirements due to hydrogen's reactivity - everything else can just use the best material by weight/cost, hydrogen needs the best that it won't eat from the inside.

Ah, missed. Thanks.

Northstar1989 · February 22, 2015

And that turbopump is going to be huge, having to deal with LH2 for all of the propellant. From the wiki article nuclear thermal rockets, we find that a J-2 thrust spec engine would need turbopumps almost as large as the F-1's. As a side note, the fuel pump is not a heavy component in a jet engine, nor is it a turbopump - where rockets are dealing with hundreds to thousands of litres per second, even the most powerful turbofans peak at less than 5.

Indeed the turbopump WILL have to be huge. So, we finally agree on something?

I never said Thermal Turbojets require a fuel pump, though- because they don't. However I did imply that they save significantly on mass by not having one. I'd like to see your data on just how "light" a high-thrust jet engine fuel-pump is, though...

Between not having a fuel pump, not having a combustion chamber, having a much lighter compressor for the same Thrust, and not having any fan blades, you end up with a much better TWR at high-speed/high-altitude flight. Thermal Turbojets are turbojets, not turbofans- they work best at high speeds an altitudes. Compared to a chemical turbofan, they are (at least slightly) lighter, they can operate faster/higher (the reduced compression-ratio means cooler airflow inside the engine for a given speed, and turbojets work better at high speeds and altitudes than turbofans anyways...) and they require absolutely no fuel...

The last of which is a very big deal- because the SR-71 had 5.4 tons of weight in engines, but many times that weight in fuel it had to carry onboard. Not needing to carry any fuel means you get a lighter wingload (the higher ISP of the rocket mode means you get a lighter wingload still compared to a chemical spaceplane). The lighter wing-load can be either kept, or exchanged for smaller (and more highly-swept, lower-drag) wings. Either one will benefit your ascent to orbit.

As for the low density of Hydrogen, I don't think you get, *it really doesn't matter*. Even if you can only achieve a 1:10 mass ratio for Hydrogen fuel-tanks, you can only achieve a 1:50 ratio for Oxygen or Kerosene fuel-tanks (your figure of 1:100 mass-ratio was *highly* unrealistic).

The vast majority of your plane's mass will be in fuel, wings, engines, and payload either way- not in the fuel tanks. Even if you need 5x more fuel-tank mass, that doesn't significantly cut into the total mass-savings you get from using Hydrogen instead of Hydro/LOX or Kero/LOX... And, as I stated repeatedly, you don't want Hydrogen the whole way up- you want, at the very least, a Methane "kicker" to decrease the total size of your spaceplane. Methane is actually denser that Hydro/LOX (due to the very low density of the LH₂-component of the mix).

Methane also works as a potential coolant for pre-coolers (it is still cold enough to make a useful heat-sink), but it is denser than Hydrogen, and has a greater temperature-difference between its melting point and boiling point, meaning you can store it *just above* its melting-point, and heat it all the way until it starts to vaporize, for a potential reduction in the total propellant-mass consumed by the pre-coolers (you can also still use the Methane in ramjets, just like Hydrogen...)

It's going to get hotter than that if used as a fuel, which is what was being discussed. In that consideration, yes, as stated, the microwave section is useless. And I also specifically stated "this stage"; ie, the airbreathing phase of flight.

We're not still talking about microwave thrust-augmentation of a SABRE-like stage, so please let go of that. If we *WERE*, though, you've got it wrong way around- you don't combust the Hydrogen and *then* heat the exhaust gasses, you pre-heat air (with Microwaves) and *then* combust it with Hydroge- to achieve even more thrust production than with either a SABRE or a Microwave Thermal Turbojet alone. I'm still not sure it would make sense from a mass-standpoint though, as you would essentially be building the entire mass of a SABRE engine onto the rear of a Thermal Turbojet for a relatively minor increase in Thrust- which is why I said it was ultimately probably a silly idea...

An SRB is phenomenal thrust at minimal expense, and dropped when expended and forgotten about. That's 0 dead weight. You were saying?

With an plane, the SRB isn't going to get you any higher/faster than a Microwave Thermal Turbojet if you take off horizontally, and is going to create HUGE safety-issues. If you're talking about a rocket on top of an SRB, then you're talking about an entirely different beast altogether- and *not* one that is superior to a plane (because, Lift/Drag works *greatly* to your advantage, even when under rocket-propulsion...)

The moment you're in rocket mode, you want to be getting the hell to orbit.

No, you don't. You're actually starting to sound a bit like one of those KSP players who launches "straight up then over"... An *efficient* ascent focuses on horizontal velocity first and foremost- vertical ascent is only a secondary-priority to make sure you're staying in thin enough atmosphere that you don't end up traveling faster than your terminal velocity for a given speed/altitude (which is, in theory, the optimal speed to ascend at), or reach a dynamic pressure that becomes unmanageable...

An *efficient* spaceplane ascent spends as much time inside the upper atmosphere as possible, relying on wings to hold it up while its entire engine power goes into building horizontal velocity. This holds true in Kerbal Space Program (with FAR installed), and it holds true in real life. You will get to orbit *faster* this way than if you push yourself into a suborbital trajectory, and end up having to use *Thrust* to hold your spaceplane up...

Any excess fuel expenditure from that point is a waste - and while it does reduce fuel flow, remaining in lifting flight massively increases time, and thus total fuel requirement.

Nope. The fuel-requirement is *less* this way. I can't be bothered to try further proving that to you- if you don't understand that, you don't understand how spaceplanes work at all in the first place...

Yes, it reduces peak power - but maybe that indicates a problem with this propulsion system. You also need to ascend as fast as possible, as if you try continue climbing flight, you're going to throw yourself straight into excessive heating phases rivalling re-entry, and enduring much longer.

You *are* correct in that the atmospheric heating is a significant problem. But not *nearly* so much as you think. Because lift is proportional to velocity^2, you can fly at extremely high altitudes where atmospheric heating poses a *much* smaller issue than during re-entry (the only parts of your vessel that will have a really hard time with it at this altitudes are the leading-edges of your wings and the nose- both of which require some sort of active-cooling or use of thermal tiles for a spaceplane ascent that spends an extended period of time in-atmosphere under rocket propulsion...)

Also, your wing-loading is likely going to cause aerodynamic issues. I've avoided pointing it out, as there's plenty of problems with just getting the propulsion, but at your sheer size you're going to get terrible wetted aspect ratio, so you're really going to struggle for L/D. And it's L/D you're wanting, not simply low wing-loading. I'm sure -a- vehicle is constructable, but it's only ever going to be fairly mediocre aerodynamic performance, at best.

I think you have some mis-conceptions about how supersonic/hypersonic flight really works. The wave drag of the wing is much, much, much more important than induced drag at supersonic (and even more so, hypersonic) speeds.

Therefore, you want wings with a very high sweep, even if this comes at the expense of aspect ratio, in order to decrease the length of the shock wave at supersonic speeds... How many supersonic aircraft design have you EVER seen with high aspect-ratio wings? The Concorde is an example of a good wing shape for supersonic flight. The ASH 31 is an example of a *terrible* wing-shape for supersonic flight...

When you're optimizing for low wave-drag, having a *long* fuselage with a low density is ideal, because it allows you to build in short, highly-swept wings. This is *precisely* what using Hydrogen as your rocket propellant provides. It also so happens that use of Thermal Turbojets and Hydrogen-propelled Thermal Rockets will bring down your wing-loading, because the *only* fuel you have to carry is Hydrogen for the rocket portion of your ascent. And if your spaceplane is 80% fuel by mass, then the Specific Impulse of that fuel is *much* more important than the density of that fuel to reducing total fuel-weight...

Of course, you probably still want some Methane onboard- which will increase your wing-loading a bit, but is still efficient enough to be useful for its superior thrust and fuel-density...

I've yet to see solid numbers that make sense. You haven't explained your colossal but featherweight turbo- or other fuel pump, and we'll address things like power shortly ...

I never said that the turbopump was light. In fact, I *explicitly states* that it was heavy, when I said it was "by far the heaviest part of the Thermal Rocket". It's a cost in mass you have to bear for the higher ISP of Hydrogen...

Why would you enter lifting flight? Get a big booster, of which there are plenty of examples (SS-SRB, Ariane 5, etc), kick it up to so stupidly high and fast it can't fall down quickly enough to be an issue, then just cruise to orbit on rocket power, as slow as you like. Why keep focussing on the plane part?

Because an Ariane 5, or an SS-SRB, or any of the other boosters you're talking about, they *AREN'T* cheap. They're quite expensive. If they were in fact cheap, there would be no reason to discuss needing to use Microwave Thermal Rocketry in the first place, because we would just load our 5000 Mars colonists up on our colossal expendable boosters and send them off, no problem. But the fact that in reality it costs $10,000/kg to get anything to orbit bears witness to the fact that a better way is needed.

A plane is useful because of Lift/Drag. Because of a spaceplane ascent's advantages- spend a long period of time building up horizontal velocity at high altitude before finally going ballistic and kicking your apoapsis outside the atmosphere. It's better to use Lift than to use Thrust to hold yourself up- especially when you have to pay dearly for every kN of Thrust with a Microwave Thermal Spaceplane/Rocket (the advantage is very high ISP, and thus the ability to construct spaceplanes or very high payload-fraction rockets- *not* a low cost/kN of Microwave Thermal propulsion...)

Okay, let's do so. E = 1/2 m * v^2, so;
P = 1/2 m(dot) * v * v
T = m(dot) * v
Hang on, I see a substitution here ...
P = 1/2 T * v
Yes, I'll grant I forgot the 1/2 factor - cut all my previous claims in two.

That was why I said you were wrong.

That's still far above your estimates, and your own equations have just proven that my equation, with a constant, -is- correct, and thermal power, at absolute minimum, is half what I claimed.

All you proved is that you finally learned the proper math to calculate things. That doesn't mean any of the numbers you gave before were correct... You haven't even repeated the calculations for your old #'s, nevertheless addressed the fact that your starting-point was wrong...

No, you actually don't. Please, point out where the power of Timberwind is claimed? I can find no references to such on its wiki article, or on most of the external links and references - the rest I did not check as they didn't look like ones that would address it. Similarly, such information is also missing from NERVA's page and references. I -did- eventually find a reference to what T/P is possible - on the previously mentioned nuclear thermal rocket article, in a different section, one of the NERVA models gets it performance confirmed. It produced 334 kN - at 1100 MW. So, about a third of your claim. It's also notable that nowhere is it stated that this thrust was achieved at maximum Isp, which I'll address in a moment.

You are indeed correct, my figure for the Thrust/MW of the Timberwind was incorrect. I assumed that the "75" in "Timberwind 75" (which would have produced 735.5 kN of Thrust at 1000 seconds) was 750 MW. Going back, it looks like the actual ThermalPower was closer to 3750 MW (a bit less)- corrected, and using the NERVA numbers as a reference (0.3 kN/MW at 2600 K), the Thrust/MW only comes out to around 0.2 kN/MW at 3000K. Or about 0.5 kN/MW at a little over 2000 K...

So, your power-estimates had a 2x factor too high built into them, by your own admission. Mine has a 0.5x factor built into them- and were just as far off, but in the wrong direction...

Using the new #'s, the cost of a Microwave Thermal Rocket that uses LH2 the whole way up and achievers 0.5 kN/MW is over $3 billion- which is more than the lifetime cost of the DeltaIV program. I'll admit when I'm wrong- a pure Hydrogen Microwave Thermal Rocket doesn't make economic sense *yet* (the cost of Microwave Transmitters continues to come down). However, if you use heavier fuels like Methane and Hydro/LOX (in a LOX-augmented Thermal Rocket) in the lower stages, you can still get a rocket that is close to cost-competitive for the same payload-capacity...

All this assumes that it is optimal to try and match the payload-capacity of a DeltaIV rocket, however. The numbers say otherwise. It would be *MUCH* better to launch a larger # of smaller payloads. For instance, if you need a satellite in Geosynchronous orbit, to launch the satellite in one launch and a disposable transfer-vehicle in another (which is inferior to a reusable infrastructure of tugs and fuel-depots, but doesn't have the same engineering challenges...) And, because the same Microwave Beamed Power can be used *beyond* Low Earth Orbit as to get there (but NOW for plasma-thrusters: don't make me get started on what a 50 MW plasma-thruster powered by Microwave Beamed-Power is capable of...), you should actually be able to get by with launching less fuel-mass to orbit in the first place...

If we return to the section we were originally looking at in the NTR article, we find that it actually performs calculations of T/P for various engines. Interestingly, in all cases, the same P = T*Ve/2 equation is used. It concludes that at 850 seconds, you would be capable of achieving less than 0.25 kN/MW, at perfect efficiency.

That's nonsense. That's Wikipedia users being stupid. We've already seen that NERVA accomplished better than that- 0.3 kN/Mw at 850 seconds. Those numbers have been *proven*- so anything Wikipedia users say about the maximum theoretical performance being lower than something that was actually tested and demonstrated to work better than that is just pure nonsense...

And with a lower exhaust-temperature, you would get better Thrust/MW (the 850-1000 seconds of ISP with a Microwave Thermal Rocket that all the sources quote apparently comes from a larger expansion-ratio or a more efficient nozzle-design... Or perhaps "hotter than 2000 K" really meant closer to 2600 K- in which case you're right back to the *proven* 0.3 kN/MW of the NERVA)

This is the fundamental limit by ratio of velocity to energy. Even at the J-2's measly 414 seconds, the performance ratio is less than 0.5 kN/MW.

Nope. It's 0.30 kN/MW with NERVA at 850 seconds. That numbers is a hard *fact*. So the worst it could possibly be at 414 seconds is around 0.6 kN/MW...

We can see that the NERVA's 334kN/1100MW is between these two - so we must conclude that it was operating at reduced Ve, to gain the thrust necessary for whatever phase of flight demanded it.

The 334 kN was indeed at vacuum ISP. The engine didn't even have the CAPACITY to reduce its ISP to improve its Thrust- so what you're saying is just pure nonsense... No offense.

In fact, the equation can be rearranged, and we find that to achieve the claimed 1 kN/MW, your exhaust velocity would have to be 2 km/s - or less, if inefficiencies exist. This is amusingly worse than almost every chemical fuel bar cold-gas - though with a dense and cheap working fluid, it might make an effective launch vehicle by allowing exceptionally tiny sizes, and making the power requirements almost reasonable. I guess then you have to balance whether you want to power the crazy thrust for such an awful mass-ratio, or the crazy Isp to actually keep the vehicle light.

Look, 0.3 kN/MW (the *proven* figure of NERVA at 850 seconds ISP) isn't all that bad. Even if it's not possible to exceed that at an ISP of 850 seconds, you haven't made any kind of a point that a Microwave Thermal Spaceplane is unfeasible. The *whole point* of using a spaceplane is to get by with less beamed-power. And wings allow you to do that quite well. And, you've said nothing that defeats the proven ISP performance of a Thermal Rocket either- so when it comes down to it, in the end Microwave Thermal Spaceplanes *do* make sense...

Because you don't *need* the payload-capacity of a Delta IV when you can bring down the cost/kg to half as much by launching half-size payloads (halving the payload/launch doubles the # of launches, and cuts down to half as much the amortized cost/launch of the Microwave Transmitters which ultimately drive up the cost...) And you don't *need* a TWR of 1.28 on the launchpad when you can get by with a TWR of 0.5 on the runway. And you don't need a complex plan to separately recover upper and lower-stages like Space-X has in mind (although so far, they're only close to recovering the launch-stages) when you can just fly the whole spaceplane back to the runway after every flight...

And, if all this STILL isn't convincing that you can build an SSTO Microwave Thermal Spaceplane, you can always build a suborbital spaceplane that releases a Microwave Thermal Rocket above the atmosphere. Because if you *really did* want to get "the hell to orbit" and out of the atmosphere as quickly as possible after switching to rocket-propulsion (although as I've already pointed out, that's *not* how a spaceplane would reach orbit, and that's *not* the most efficient ascent. You *want* to stay in the atmosphere as long as possible- or at least until Lift/Drag = 1...) then there's no point in having wings once you're above the atmosphere- and deploying an upper-stage rockets gives you all the mass-fraction benefits of staging...

You didn't actually use the maths, and it turns out it was against you.

No. I turns out we were bother off by a factor of 2- which is equally wrong...

You didn't cite how the real-world examples possibly supported your assertions, and now that I've actually done research on them, it turns out, they support the opposite.

I've extensively researched real-world examples. But I made some mistakes in my research, and you now corrected them. You can rest assured that the KSP-Interstellar Extension Config I've been helping to develop will now have a realistic Thurst/MW of around 0.2-0.3 kN/MW, like the real-world NERVA...

You will, fundamentally, require very large powers for the rocket phase.

I never said anything to the contrary. To get 9.24 tons to LEO, it turns out you need around 1.6 GW of beamed-power. I wouldn't call that small. But, at least half the mass of a satellite in LEO destined for GEO is fuel and engines- so if you simply launch a transfer-stage and the satellite in two separate launches, you only need 800 MW...

The *minimum* mass to LEO to get a small 1-man capsule up there is around 1 ton (the weight of a Mercury capsule without a Launch Escape System...) So if you built a spaceplane that could carry 1 ton to LEO per launch, you could even launch manned missions with only about 100 MW of beamed-power (that's using your own figure of a TWR of 0.5 on the runway being all you need to get to orbit...) Which only amounts to $200 million in Microwave Transmitters, and maybe $400 million in R&D costs...

Amortize that over, say, 20 launches a year to the ISS (most of those crew-rotations and consumables, with 1 man riding up on each manned launch...) over 10 years, and you get a cost of *only* $3,000/kg to LEO, with a marginal cost of virtually nothing if you want to launch more to orbit each year than just ISS consumables and crew (because you've *already paid for* the Microwave Transmitters and R&D, additional launches each year cost virtually nothing but fuel and power...)

And an air-breathing phase is dubious - further supported by the fact that the only people researching microwave power are ignoring the possibility.

There's nothing dubious about an air-breathing phase. The only thing we should be debating is *just how good* it would be, not whether it would work in the first place. We worked on Nuclear Thermal Turbojets back in the 1960's, and concluded there were no major obstacles to their use, although John F Kennedy shut the program down for political reasons and because he was afraid it might lead to a nuclear war...

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

Yes, I'm sorry I forgot the 1/2 factor. I made a hundred or so launches required to amortise costs into a couple hundred.

See my numbers above. 200 launches over 10 years with a 1 metric ton to LEO spaceplane equates to a cost/kg to LEO of maybe only $3000/kg (assuming $400 million in Research and Development costs- which is quite extravagant considering how close the technology already is to realization...)

Sorry. But the 1 kN/MW figure has just been totally debunked, with none of your examples actually supporting it, hopefully to your satisfaction, and so many of the previous cost estimates stand - and yours remains physically impossible, unless you want to completely rework your design into a tiny, hyperdense vehicle with only 2 km/s Ve.

You're right. 1 kN/MW was inaccurate with Hydrogen propulsion, as the figure for NERVA was only 0.3 kN/MW at 850 seconds (I should point out now that it's *ENTIRELY* realistic with a denser fuel like Ammonia, which would get more than 5x the Thrust/MW, at 63% the Specific Impulse, thanks to its breakdown into Nitrogen and Hydrogen at high temperatures *greatly* aiding Thrust production...)

That doesn't mean the concept is unfeasible- only that it's even more important to amortize the cost of the Microwave Transmitters (as more than 3x the transmission-capacity will be required of what I originally expected) over as many launches as possible. Which is why companies like Escape Dynamics are looking at building a low payload-capacity vehicle that launches larger payloads (like a large GEO satellite, or a 3-man crew for the ISS) over multiple smaller launches... (such as by a separate launch for the satellite and a transfer-stage, or 1 crew member at a time...)

As above, multi-GW stands at any actually high Isp.

No. Multi-GW is only necessary for large payloads. And if you are using a spaceplane, you only need about half the Thrust on the ground (and can use Thermal Turbojets to get up above the thickest part of the atmosphere- Mach 2 and 24 km is perfectly reasonable for them...) Build a spaceplane that launches 1-2 tons to LEO per mission, and you can amortize the cost of the Microwave Transmitters over many, many launches- and get a *very* low cost/kg to LEO...

3 GW is entirely valid at 50% efficiency, which is what rockets themselves typically get just from thermal->kinetic conversion by expansion near sea level. 75+%is seen at very high expansions in very low pressure, but that still leaves you happily around 2 GW - though I'll grant you could probably reduce thrust to maybe 1.5 GW peak power by these conditions.

So, 0.3 kN/Mw in vacuum, 0.2 kN/MW at sea-level, when using Hydrogen. But you don't use Hydrogen at sea-level, that would be silly. (because Hydrogen is highly vulnerably to ISP-losses from atmospheric-compression at sea-level)

You either use Thermal Turbojets or a heavier fuel such as Ammonia (which *does* get you 1 kN/MW at sea-level, at around 500 seconds ISP, and is *much* less vulnerable to atmospheric-compression due to the higher Exhaust Pressure you get this way- don't try to run this number through E = 1/2 m * v^2, because it doesn't compute. You're getting additional energy from the breakdown of Ammonia into Hydrogen and Nitrogen...) With one of those, you save the Hydrogen for high speed+altitude, where it's most effective... (theoretically, you get the best energy-efficiency from matching exhaust-velocity to vehicle velocity: so it makes more sense to use Hydrogen when you're already traveling fast...)

Assuming no transmission loss, of course. Thermal turbojet numbers are highly dubious, as many of your calculations seem to be based on ratios to rocket performance, which has been debunked, and also jet design difficulties, even thermal, mean we should consider ignoring them and focus on the rocket phase - which still doesn't stand up.

Building an air-breathing stage isn't an impossibility. Will you stop talking about one like it is? If jet engines were impossible, we wouldn't fly passenger airlines around the world *all the time*. The fact is, jet engines are perfectly doable, and Thermal Turbojets are just an extension of the same concept that is actually *easier* to design and build due to the lack of having to maintain a fragile combustion reaction. Since you still don't believe me when I tell you *combustion is sensitive in a jet engine*, please see the following figure:

This graph effectively summarizes the main reason why designing a Thermal Turbojet is so much easier than designing a normal jet engine (which is done all the time)- we don't have to deal with any of those combustion-limits... (and can operate at a much higher airflow- notice how the combustion reaction becomes more fragile at higher airflows? *That* is why we have to limit airflow at high speed on our chemical jet engines...)

LH2 is actually 70 kg/m^3, but the point stands, yes. As mentioned at the start of my post - if you do research on the actual tanks, LOx and RP-1, hypergolic fuels, etc, that are all closer to 1 t/m^3, all have ideal tankage ratios around 1%, but hydrogen is always around 10% at best, due to the volume, extreme insulation required, and funky material requirements due to hydrogen's reactivity - everything else can just use the best material by weight/cost, hydrogen needs the best that it won't eat from the inside.

A 1:100 mass-ration for LOX or RP-1 seems *highly* unrealistic to me. Try 1:50 for a more realistic number. Whereas a 1:10 mass-ratio is the actual mass-ratio of the Centaur upper stage- which utilized Hydro/LOX propulsion... (and, due to the low density of Hydrogen, the vast majority of that tankage was for LH₂...)

Regards,

Northstar

Edited February 22, 2015 by Northstar1989

DuoDex · February 22, 2015

While this thread does have high-level science discussion in it, it has devolved into acrimonious debate and personal attacks.

So, thread closed.

Edited February 27, 2015 by DuoDex

Escape Dynamics and the Microwave Thermal Spaceplane

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