FF for Jet Engines?

[zapkitty]

Aeronaut wrote: I thought ‘less than friendly neighbor’ meant drastically limiting shielding mass…

For the concept that I wandered off exploring the FF cores are the thrusters. But that involves passing the alpha beams through the shielding

The only hazard in that concept would be if you were directly behind the ship… but that would get very unhealthy very fast. Thus the cost of doing business with that design goes up equally fast.

And with the power available with FF we’re no longer at the stage of trying to reach orbit by any means possible.

The MHD research means that FF-powered supersonic jets can be done for far less MWe than I had thought… but there’s a ways to go from Mach 3 to orbit.

[Aeronaut]

zapkitty wrote:

I thought ‘less than friendly neighbor’ meant drastically limiting shielding mass…

For the concept that I wandered off exploring the FF cores are the thrusters. But that involves passing the alpha beams through the shielding

The only hazard in that concept would be if you were directly behind the ship… but that would get very unhealthy very fast. Thus the cost of doing business with that design goes up equally fast.

And with the power available with FF we’re no longer at the stage of trying to reach orbit by any means possible.

The MHD research means that FF-powered supersonic jets can be done for far less MWe than I had thought… but there’s a ways to go from Mach 3 to orbit.

Doesn’t the alpha (of helium ions) moving at near lightspeed already pass through the shielding? A central core flanked by the remaining cores would only need about a 1 foot spacer in the front to get all the beams to converge where you want to heat the propellant. In that case, existing “Beware Jet Blast” signage should suffice, unless you want to add an image of Donald Duck getting blown *ss over tea kettle.

What might happen if we approached this from ionizing atmospheric nitrogen for the MHD? That’s the bulk of a turbojet engine’s reaction mass. All the oxygen does is make the fuel combustible.

[Brian H]

Aeronaut wrote:
…
Doesn’t the alpha (of helium ions) moving at near lightspeed already pass through the shielding?
…

!!
Relativistic alpha particles? Seems unlikely. Where did you get that info?

[vansig]

exit beam alphas should be roughly from 600 to 2900 keV, which is *not* relativistic.

at the upper end of this, 2900 keV = 4.646e-13 joule, mass ~ 4u = 6.64e-27 kg
and kinetic energy, E_k = .5 m v² so

v = sqrt( 2 E_k / m) = sqrt ( 2 x 4.646e-13 / 6.64e-27 ) = .04 c

the bulk of them are probably closer to 600 keV = 9.613e-14 joule,
v = sqrt ( 2 x 9.613e-14 / 6.64e-27 ) = 0.018 c

*after* extracting 80% of their energy with the capture coil, they’ll still be traveling at 2400 km/s, which is 42x the relative momentum of any given particle in the compressed airstream at mach-24. but even a small number of collisions with any air particles will change that, in the same way as shooting a break at a pool table slows the cue ball.

by the way, for each 5 MW anode, i estimate ~ 10^17 alphas in a pulse, 1000 pulses/sec.

for a 400 MW engine, that’s 8 x 10^21 alphas exiting each second. (about 53 mg)

in other news, industrial physicist has a good article by Dean Andreadis, showing scramjet engine design
http://www.aip.org/tip/INPHFA/vol-10/iss-4/p24.pdf

[zapkitty]

Aeronaut wrote:
A central core flanked by the remaining cores would only need about a 1 foot spacer in the front to get all the beams to converge where you want to heat the propellant. In that case, existing “Beware Jet Blast” signage should suffice, unless you want to add an image of Donald Duck getting blown *ss over tea kettle.

The problem here is the radiation flux from the cores. Channels that would let the alphas out would perforce let other things out that aren’t nearly so benign.

Aeronaut wrote: What might happen if we approached this from ionizing atmospheric nitrogen for the MHD? That’s the bulk of a turbojet engine’s reaction mass. All the oxygen does is make the fuel combustible.

Not really an issue, I think. The MHDs can handle the atmospheric portion of the flight pretty much as they are described in the literature.

The issue is the transition from .1 atm at 65kft to .001 atm and beyond. My brilliant idea had the cores augment the MHD thrust by direct alpha heating the atmosphere as long as there was atmosphere to get in the way. I figured that’d be sufficient to get a suborbital trajectory at mach 18+… enough that the FF thrusters could take over alone and complete the ascent to orbit as my initial estimate had the craft at 34 tons… it was a tiny thing… but onboard propellant and the structure to hold it will change that…

… of course if you go old-school and just have a shadow shield the problem is easily solved but the operational costs will go up accordingly…

[vansig]

Quantitatively, what is the expected drag on a craft as it reaches orbital speed?

A 100 tonne craft, (roughly 3 m² area, and 16 m long), flying at ~36 km altitude (pressure=.01 bar) and mach-24 (8 km/s), needs 980 kN lift to remain climbing, but most of the lift is provided by centripetal motion, since 8 km/s is sufficient for orbital velocity.
( v²/R = 8² /(6371 + 36) = 0.0999 km/s² = 9.99 m/s² )

So only forward drag really matters, here. From the formulae at
http://www.grc.nasa.gov/WWW/K-12/airplane/drageq.html
http://www.grc.nasa.gov/WWW/K-12/airplane/dragco.html and
http://www.grc.nasa.gov/WWW/K-12/airplane/eqstat.html
as well as
http://www.lpi.usra.edu/meetings/lpsc2009/pdf/2059.pdf

We get that form drag dominates at high mach numbers. Assuming drag coefficient, Cd ~ 0.92; cross-sectional area, A ~ 3m², and density of air (at p=.01 bar and 244 K), rho = 0.0143 kg/m³,

then drag = Cd x Area x rho x v²/2
= .92 x 3 m² x 0.0143 kg/m³ x (8000 m/s)² /2 = 1263 kN

practical flight will have to exceed this by a good margin, for at least several minutes as these speeds and heights are obtained.
compare this needed performance to a gang of 3x GEM-40 boosters, (at 1000mm diameter and 492.9 kN each, 64 seconds burn time, only).

[zapkitty]

I’m confused… why would you try to do mach 24 in the stratosphere?

My idea was to start direct alpha augmentation along with the MHDs in the stratosphere in order to accelerate as much as feasible and throw the craft into an arc that would carry it into the thermosphere. By the time the MHDs quit and the FFs ran out of air to augment their thrust… well, by then drag wouldn’t be a problem (until the arc ended in reentry, that is.) I wasn’t figuring on the FF units fighting atmospheric drag by themselves.

[zapkitty]

zapkitty wrote: I’m confused… why would you try to do mach 24 in the stratosphere?

…. because that’s what I said in that wishlist I jotted off.

My apologies. I shouldn’t have been juggling metric and imperial. I put 100kft when I was thinking 100km.

[vansig]

I might not want to do mach 24 in the stratosphere, but it is a conversation point.

The engine needs to have enough air intake to give net forward thrust and lift, after overcoming drag, at any given altitude and velocity. Flying higher reduces drag by reducing density of air, but thrust also declines, requiring faster flight to feed the engine. There is a cross-over point from air-breathing to on-board propellant, as the min and max altitudes for hypothetical top speed converge.

At speeds much lower than mach-24, the craft cannot rely on centripetal motion, so engines and surfaces must provide for both drag and lift. it is interesting that there is a point where these are in the same magnitude. my hope is that the discussion will lead to a proper formula for a flight curve for both suborbital trans-continental transport and orbital insertion.

neglecting drag, perigee at 8 km/s and 36 km altitude would entail apogee near 400 km, and delta-v of ~70 m/s to complete the maneuver. finding the point along this curve where we must switch to on-board propellant will yield knowledge of how much propellant we need to carry.

[zapkitty]

vansig wrote: I might not want to do mach 24 in the stratosphere, but it is a conversation point.

I understand. But I also have to go back and see if any of my patched-together approximations have any connection to reality…

[vansig]

Mine, too. 🙂

One very significant approximation is the power required for hypersonic flight. I’m beginning to wonder whether my early guess of 400 MW is in the right order of magnitude.

My present best-case reasoning is leading to 800 MW, and worst-case 17.6 GW, so i need to debunk my assumptions.

Comparing again to solid rocket booster performance,
the GEM-40 burns HTPB 12%/ammonium perchlorate 68%/powdered aluminum 20%,
at a rate of 185.8 kg/s, thrust 499 kN, Isp 274 s, nozzle ratio of 16:1

Typical ramjet Isp is higher: 1400 s, i am told. Scramjets would be higher still, i assume.
I plugged in assumptions for two different configurations:

“Small mouth” assumes total mass = 100t, form drag 1250 kN at v=8 km/s, air intake 102 kg/s at p=0.01 bar;
and produces minimum needed exhaust velocity = 20236 m/s.

“Big mouth” assumes total mass = 100t, form drag 1250 kN at v=8 km/s, air intake 1020 kg/s at p=0.01 bar;
and produces minimum needed exhaust velocity = 9224 m/s.

If this exhaust velocity is obtained by non-thermal processes, then the entire mass is accelerated aft, and the kinetic energy required each second comes from the difference between the propellant’s intake and exhaust speeds, relative to the ship.

E = E_new – E_old = .5 m v_new² – .5 m v_old² = .5 m ( v_new² – v_old² ),
= .5 (102) (20236² – 8000²) = 17.62 GJ for small mouth,
or
.5 (1020) (9224² – 8000²) = 10.75 GJ for big mouth.

If this exhaust velocity is obtained by thermal processes, then temperature is given by
T = v_rms² M / ( 3R ), where M = mass of 1 mole of air particles in kg, and R = 8.314 (the gas constant)..
= 20236² (0.029) / (3 x 8.314) = 476120 K, for small mouth; or
T = 9224² (0.029) / (3 x 8.314) = 98925 K, for big mouth.

I’m not liking these numbers, at all. They make me hope i’ve overestimated something. (eg: drag? total mass?).

But, rockets can reduce drag by flying higher; and jets can reduce drag by flying slower, each with their own consequences

[vansig]

ok, these numbers are a little better.. they do not yet deal with heat management, but i’ve looked at other flight parameters, such as lift, drag, rate of climb, shock waves; at high mach numbers, air intake needs to be a large fraction of the forward facing area, (about half). at hypersonic speeds, leading surfaces will be attached to the shock wave, to stabilize flight and prevent air from becoming stagnant.

let me know what looks impossible here.

assume 100t mass; (47t empty);
lift/drag ratio 4.5; glide ratio 4.5 (about the same ratios as the space shuttle)
122m² underwing area

calculate, density-versus speed, given approx constant air inflow of 80 kg/s, 600 kN thrust, 15500 m/s exhaust velocity, and drag from 218 kN to 300 kN. the engine possibly cannot operate at below mach 2.4, but i have not proved that. particularly, thrust may possibly be doubled in dense atmosphere, with extra energy.. which would enable vertical take-off.

mach 2.4+ at 36 km altitude (density ~.01 kg/m³; drag ~218 kN);
mach 8 at 45 km altitude (density ~.003 kg/m³);
mach 24 at 54 km altitude (density ~.001; drag ~300 kN);
net forward propulsion to >100km

disregarding efficiency, *any* sort of propulsion needs to add a lot of kinetic energy to the incoming air
E_k = .5mv² = .5(80 kg)(15500)² = 7 to 10 GJ each second

if system efficiency were terrible, say 1%, then
decaborane consumption rate would be 14.5 g /s, and this makes
Isp = 600,000 kg m/s² / ( .0145 kg/s ) / 9.8 m/s² = 4.2e6 s

efficiency is probably not that terrible. like VASIMR, traditional pressure from heating becomes inefficient at the needed exhaust velocity, so the engine will rapidly ionize air and then accelerate it magnetically, before it thermalizes. I’ve not seen a good argument as to why VASIMR is supposed to be operated in vacuum only. However, Ad Astra does discuss formula for efficiency of their engine. http://www.adastrarocket.com/AIAA-2010-6772-196_small.pdf

finally, thrust bottoms out at ~2 kN in space proper, where alpha particles go directly to exhaust

Oh, and here’s a nice article, mentioning DPF and nuclear thermal rockets
http://quantumg.blogspot.com/2011/02/making-fusion-rockets-relevant.html

[QuantumDot]

well there are some technologies that can change a number of those figures.

the biggest new technology would be plasma aerodynamics would has in scale models reduced drag by 30 percent and reduced heating by 50 percent. the current major problem with it is the amount of power needed to operate it.

the heat shielding system i think is about 30 percent of the dry weight could be significantly reduced with a better reenter trajectory and with new technologies the mass could be significantly reduced even more with flexible aerogels, new ceramic blankets, phase change materials, the already described plasma aerodynamics, meta-materials, etc.

current hydraulic flight control systems take about 10 percent of the dry weight of the space shuttle, there are currently two real alternatives either ferromagnetic shape memory alloy or plasma based flow control.

then there is the basic structure and design which could be moved to a more blended wing body and the mass could be reduced with new technologies. no real facts here but 20 percent more room and 30 percent less mass seems possible.

a precooler system should be able to improve the performance the only question is the trade off between something like the system for skylon or a pulse tube cryogenic cooler.

[zapkitty]

Seems that a lot of preliminary work has actually already been done on all-electric propulsion. It’s the obvious next step and keeps showing up in the trade spaces because of the benefits that accrue even if the electric generators are chemically fueled. Funny how that sort thing happens…

Yes, Virginia, you can have a supersonic Prius 🙂

More seriously, the mass savings that are to be had with fully-superconducting engines and all-electric control systems are indeed considerable and several of the larger proposed designs would seem to be large enough to swap FF units in directly. The FFs substitute for the generators and fuel, they would directly power cryocoolers thus reducing the storage volume for needed for cryogens, and they would greatly reduce the frontal drag of intake area as the FF units are not airbreathing.

Going to try and see where all that might take us with FF-powered craft before going to the next step…

An example from the ICAS 2010 conference
http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20100036222_2010039460.pdf

[vansig]

here is an article detailing scaling parameters for EHD thrusters.

http://www.wbabin.net/physics/borg1.pdf

[Author]

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