Re: sci space policy targeted by disinformation experts?

On Mar 28, 7:45 am, Willie.Moo...@xxxxxxxxx wrote:
On Mar 27, 7:59 am, Ian Parker <ianpark...@xxxxxxxxx> wrote:



On 27 Mar, 02:52, Willie.Moo...@xxxxxxxxx wrote:

A lot of speculation - I recall reading really interesting stuff, that
just falls off the radar screen so to speak for no damned good
reason. Usually when something doesn't work for a sound technical
reason, you can find some arcane journal article explaining why. When
you cannot find that, there is a possibility - if the ideas are sound
otherwise, they've been taken black.

One way to check that out is to track the researchers. Are they
teaching and not doing a damned thing, or are they busy and have moved
from where they were to points West and stopped publishing?

Thats another inferential point to anyone who cares.

Energy is a problem with high speed flight. Aurora nominally burning
hydrogen in air in an external combustion scramjet - and a 10%
structural fraction - producing thrust by intercepting the shock
waves. You eject the fuel into the stream at the stream velocity -
right at the shock wave at the nose - so its stationary in the flow.
By the time it reaches the thrust structure at the rear of the
aircraft, its mixed with an oxidizer - you stablize that with an
expansion shock, and detonate it with a laser or spark or particle
beam - and the shockwave and thrust surface are shaped to interact to
produce thrust.

Mach 6 and drag coefficient gives you an estimate of power. The X-15
had a drag coefficient at hypersonic speeds of Cd = 0.095

Drag force is equal to

F = 1/2 rho V^2 * Cd * A

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

area looks to be in the 30 sq m range

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

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

Mach 6 is around 1,800 m/sec, and rho=0.01 kg/m3

So,

F = 1/2 * 0.01 * (3.24e+6) * 0.095 * 30
= 46,170 newtons
= 4,701 kgf
= 10,343 lbf

at around 50 km altitude

Force times distance is energy.
Force times speed is power

So, 46,170 newtons x 1,800 m/sec = 88.106 megawatts

Hydrogen when burned in air releases 143 megajoules per kg. Assuming
1/4 of this energy is usefully applied to the propulsion system,and
3/4 of the energy is wasted in various ways - means 35.75 megajoules
of propulsive energy is available per kg of hydrogen. This gives us a
burn rate of 2.46 kg/sec to maintain that thrust. With a 50% cycle
efficiency - fuel use is cut in half 1.23 kg/sec

This is the likely fuel consumption of hydrogen for the aircraft at
this speed - from first principles.

Going back to our models of Aurora - it likely has a 600 cubic meter
fuel volume. and hydrogen has a density of 70 kg per cubic meter
which obtains 42,000 kg fuel mass. Enough to power the aircraft for
4 hours and 45 minutes at Mach 6 cruise - at thelower efficiency, and
9 hours 30 minutes at the higher efficiency. Enough to fly 3/4 of the
circumference of the Earth at cruise at the lower efficiency, and 1.5x
around the world at the higher efficiency.

One can imagine a number of interesting missions for such an aircraft
if it exists.

.

We can only argue plausably in this. 0.095 is in fact quite good for
supersonic speed. It should be recalled that a typical subsonic
aircraft has an L/D of about 20. Flying wing configurations improve
this as the drag from the fusilage is eliminated.

It will be recalled that Concorde was about 7:1 and with most of its
take off weight fuel just about limped to Washington from Heathrow. We
don't know whether Aurora was 10.5:1 or some worse figure. Discussing
an L/D ratio for hypersonics is a little bit misleading as airflow is
so integrated with engine performance.

- Ian Parker- Hide quoted text -

- Show quoted text -

The figure for the X-15 is accurate, and that was at hypersonic
speed. Cd(min) for subsonic speed was for the X-15 was 0.0645.

I figured the Aurora if it exists would do at least as well since it
built on this base.

Blended body design probably helps too, as does an integrated
propulsor. That is, no fans or inlet ducts, just energizing the
boundary layer with high speed fuel injection and detonating the air/
fuel mix at an appropriate spot, but according the helmhold equation,
aspect ratio helps even more.

Here are the Cd(min) which all occur at subsonic speeds, for the
various re-entry bodies tested in the USA;

M2-F1 M=0.15 Cd(min)=0.0618
M2-F2 M=0.45 Cd(min)=0.0645
HL-10 M=0.60 Cd(min)=0.0496
X-24A M=0.50 Cd(min)=0.0400
X-24B M=0.50 Cd(min)=0.0252
X-15 M=0.65 Cd(min)=0.0645
Shuttle M=0.50 Cd(min)=0.0604

Of course, having re-configurable wings, also helps in subsonic
speeds. The X-24B is an amazing aircraft.

http://www.nasa.gov/centers/dryden/history/pastprojects/Lifting/X24/i...

I thought it would be used for a 'siamese twin' two-stage launcher.
That could have been done at far less cost than the Space Shuttle.

Take two x-24B airframes and mate them belly to belly. The same
airframe size. One use a high density fuel, like kerosene, with LOX,
the other use a low density higher performing fuel, like hydrogen,
with LOX. You get an almost perfect two stage to orbit RLV.

The Skunk Works took over the X24 after they proposed an X24-C with
scramjets and a top speed of Mach 8

http://en.wikipedia.org/wiki/Martin-Marietta_X-24

Crew: one pilot
Length: 37 ft 6 in (11.43 m)
Wingspan: 19 ft 0 in (5.79 m)
Height: 9 ft 7 in (2.92 m)
Wing area: 330 ft² (30.7 m²)
Empty weight: 8,500 lb (3,855 kg)
Loaded weight: 11,800 lb (5,350 kg)
Max takeoff weight: 13,800 lb (6,260 kg)
Powerplant: 1× Reaction Motors Upgraded XLR-11 four-chamber rocket
engine, 8,480 lbf (37.7 KN)

Maximum speed: 1,164 mph (1,873 km/h)
Range: 45 miles (72 km)
Service ceiling 74,130 ft (22.59 km)
Wing loading: 205 kg/m² ()
Thrust/weight: 0.71

Even though the top speed is quoted here as 1,164 mph, the X-23
lifting bodies were based on the Air Force X-23 manuverable re-entry
vehicle program.

http://en.wikipedia.org/wiki/X-23_PRIME

this also had applications in MIRV vehicles, since warheads could
manuver across large distances, increasing throw-weight of launchers
to targets, and increasing weapon flexibility.

The thrust to weight is a function of the rocket engine used.

Space travel applications was the 'siamese twin' program, which I read
about in the day, but can find absolutely no mention of even in the
literature. Which is amazing.

I mentioned if I were to build an Aurora type plane, I would have used
a booster rocket to get it to speed, and switch on my scram jet.
Well, if I were asked to build one, I'd turn to the X-24B (and C) and
use one in rocket mode, to lift the pair belly to belly in a vertical
take off, and then release the second stage at Mach 8 and turn on the
boundary layer scramjet.

That may be what the Air-Force/National Security Agency did, and may
be why the few reports of the Siamese Twin launcher disappeared from
the radar screen.

But, who knows?

Even though the structural fraction of the one man test vehicle was
0.616, a larger version built with the best available technology of
the day would likely get down to 0.125, and using the best available
technology today, would likely get down to 0.062. Rocket engines
have thrust to weight of 80:1 and MEMs based rockets as well as
boundary layer scramjets, are likely to have thrust to weight of 800:1
or more for very sound technical reasons.

So, here's an interesting Siamese Twin system based on the X-24B

Stage 1
Crew: one pilot
Length 175 ft (53.0 m)
Wingspan 88.2 ft (26.8 m)
Height 44.5 ft (13.6 m)
Wing area: 7,109 ft² (661.4 m²)
Empty weight: 850,000 lb (385,500 kg)
Loaded weight: 11,800,000 lb (5,350,000 kg)
Powerplant: 14× Rocketdyne F1 rocket engine, 1,500,000 lbf (6.7 MN)
or 12x Glushko RD-171 rocket engine, 1,776,665 lbf
(7.55 MN)
or 12x Rocketdyne F1A rocket engine, 1,792,260 lbf
(8.00 MN)

The F1 has 265 sec Isp at lift off and 310 sec Isp at altitude.
The RD-171 has 309 sec Isp at lift off and 337 sec Isp at altitude
The F1A has 270 sec Isp at lift off and 310 sec Isp at altitude.

Stage Propellant Volume is 155,722 cubic ft. (4,411 cubic m) in each
stage, based on the ellipsoidal shape of the major airframe volume.

Volume of an ellipsoid = pi/6 * (major diam x minor diam x height)

So, put in numbers based on the span, height and length compute the
volume and divide by two - for a first order approximation - I used
100 ft x 75 ft x 80 ft as the fuel volume within the airframe - and
divided by two so the 80 ft became 40 ft. .

LOX Kerosene has an optimum oxidizer fuel ratio of 2.56 to 1.00 The
density of LOX is 1.14g/cc and Kerosene is 0.806 g/cc - this gives an
average propellant density of 1.02g/cc. That's 1.02 metric tons per
cubic meter - which gives the weight.

So this is a consistent system.

Oxygen hydrogen has an optimum oxidizer fuel ratio of 6.00 to 1.00 and
hydrogen has a density of 0.07g/cc so the mix has a density of 0.28 g/
cc. So, the same 4,411 cubic meters has 1,235,080 kg or 2,717,176 lbs
added to a lower 650,000 lbs of vehicle weight, along with any
payload. The weight is lowered due to the lower masses of fuel to
handle - principally anti-slosh baffles.

I like the German Saenger ATCRE with
Thrust(vac):1,280.000 kN (287,750 lbf).
Thrust(sl): 1,068.400 kN (240,186 lbf).
Isp: 490 sec.
Isp (sea level): 409 sec.

You'd need 10 to 12 of these,

alternatively you could use the 10 SSME,
or 2 RS-XXX

Isp is lower in all these other engines.

The engines nozzles I imagine would be blended spaced along a spanning
rib with exhaust blended into the tail section, using the airframe as
a partial aerospike for the SET of engines. Varying mass flow
...

read more »

Again, with all the usual Willie.Moo infowar hype you can muster on
behalf of sustaining your corrupt and spendy government that's taking
us into WWIII just as fast as they can, and before we run ourselves
entirely our of fossil fuel alternatives.
. - Brad Guth
.

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