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#51 2017-09-29 10:01:51
- GW Johnson
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- From: McGregor, Texas USA
- Registered: 2011-12-04
- Posts: 5,800
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Re: Amateur solid-fueled rockets to *orbital* space?
I submitted 22/22 chapters and 3 of 4 appendices to AIAA the other day. Their timeline is around 6-8 weeks to review it, then if approved 6 months or so to publish. They only asked for 2-4 chapters for the reviews, but I had nearly all of it done. We'll see what happens.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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#52 2017-09-29 19:19:12
- SpaceNut
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- From: New Hampshire
- Registered: 2004-07-22
- Posts: 29,431
Re: Amateur solid-fueled rockets to *orbital* space?
Nice looking forward to the read....
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#53 2017-12-15 19:18:17
Re: Amateur solid-fueled rockets to *orbital* space?
...
I ran my spreadsheet version of thrust coefficient at a fixed Pc = 200 psia, and reset Pe iteratively until I got Ae/At = 750.9 at Pe = 0.0108 psia. The CF for Pe = Pamb at that expansion is 1.98047. CF vac is 2.0226. These were figured for specific heat ratio 1.20 and a 15 degree effective conical half angle to the nozzle, pretty "typical" values.
Thrust per unit throat area is Pc CF = 404.52 psi, using the vacuum CF.
Can't tell you what the thrust is, without some way to set throat area. Normally that is sized to get the thrust you want out of the stage, or else it comes from an existing design. Cannot tell you what Isp is until I know something about a 200 psia chamber c*.
If for an aluminized solid based on AP-HTPB the 200 psia c* is near 4800 ft/sec, then Isp = CF c* / gc ~ 302 sec. If instead it was 4900 fps, then Isp ~ 308 sec. If it is nearer 5000 fps, then Isp ~ 314 sec.
Chamber c* is a power function of Pc of the form c* = k Pc^m, where for solids m ~ .01, although the variation of m from propellant to propellant is significant. At 200 vs 1000 psia, for m = 0.01, we lose about 2% of our c*.
If c* were known, then you could figure the propellant flow through the throat per unit throat area: w/At = Pc gc / c*.
Hope that helps.
GW
Thanks for that. Some of the Star series solid motors get remarkably high mass ratios considering the entire casing has to contain the high pressures of a solid motor combustion. See for example the Star 48B:
Spacecraft Propulsion, p. 163
https://books.google.com/books?id=P5dBC … e&q&f=true
It has a approx. 20 to 1 mass ratio, using titanium casing at an approx. 600 psi operating pressure. Suppose we could get a ca. 200 psi operating pressure. Then with proportionally thinner casing walls we might be able to get a 60 to 1 mass ratio(!) Clearly there never has been before a rocket stage with this high a mass ratio. But the prospect is tantalizing.
If the vacuum isp is then also 305s with a 750 to 1 nozzle area ratio, the vacuum delta-v could be 3050ln(60) = 12,500 m/s, well above that needed for an SSTO. There would need to be some modifications made though. According to the specs on the star 48B, the expansion ratio on the nozzle is in the range of 55 to 1. Then to get a 750 to 1 nozzle area ratio while maintaining the lightweight, it would require new, design and/or materials for the nozzle. Perhaps, the altitude compensation aerospike would work. Some other possibilities for getting altitude compensation at lightweight are discussed on my blog.
Imagine also the delta-v possible if the operating pressure could be brought down to only 100 psi.
Bob Clark
Last edited by RGClark (2017-12-15 19:19:32)
Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):
“Anything worth doing is worth doing for a billion dollars.”
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#54 2017-12-18 12:12:34
Re: Amateur solid-fueled rockets to *orbital* space?
GW, that info on the Star 48B is pretty detailed, such as nozzle throat size. Would that be enough under the assumption of an added on nozzle extension that brings the nozzle area ratio to 750 to 1, to calculate the thrust?
Bob Clark
Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):
“Anything worth doing is worth doing for a billion dollars.”
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#55 2017-12-20 15:52:07
Re: Amateur solid-fueled rockets to *orbital* space?
The aerogel material used for the space shuttle underside tiles may provide a lightweight, high temperature material for creating large nozzles at lightweight. These space shuttle tiles only had a specific gravity of 0.144.
And a newly developed ceramic is half that weight:
Thermal Protection System.
Benefits.
Low density (0.07 g/cm3 or 4.4 lb/ft3)
High temperature capability (4000°F [2204°C])
Low thermal conductivity (<1 W/m·K at 3600°F)
Ability to combine with ceramic matrix composite or coated carbon/carbon structural shells to produce an integrated airframe/insulator thermal protection system
Imperviousness to chemical attack below 302°F (150°C)
http://www.ultramet.com/thermalprotectionsystem.html
GW, has also discussed a lightweight ceramic that may be even lighter still:
Reusable Ceramic Heat Shields - GW Johnson - 16th Mars Society Convention.
http://www.youtube.com/watch?v=3MXYY3jnNr0
Bob Clark
Old Space rule of acquisition (with a nod to Star Trek - the Next Generation):
“Anything worth doing is worth doing for a billion dollars.”
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