China's Space Program Thread II

[by78]

A model of the Red Dragon One engine from the private company Tianhui Aerospace. LOX/Kerosene, 1000kn thrust, specific impulse 303s, adjustable thrust 50% to 105%.

 

[Helius]

Dude, it is absolutely no way that any solid engine can have 383.9 Isp, except adding black magic into propellant.

Typical NEPE has a Isp around 260s at the sea level, and HTPB around 248s. China's best propellant is around 271-273s at sea level, by replacing Al with AlH3 into propellant. But it can't be mass produced as AlH3 has discrepancy between batches.

View attachment 100939

Hey you're preaching to the choir. What do think I'd been saying these past few pages? I didn't believe it'd be this high either.

Then again, 150t in 115s is a really fast burn. If I were to guess, either the propellant is highly reactive or the fuel mix is very thin, possibly semi-liquid, maybe both. It's the fuel velocity that really jacks up the Isp.

So I think you'd be better off getting someone else to defend it because I'm just as stumped by these results.

 

[Helius]

Hey you're preaching to the choir. What do think I'd been saying these past few pages? I didn't believe it'd be this high either.

Then again, 150t in 115s is a really fast burn. If I were to guess, either the propellant is highly reactive or the fuel mix is very thin, possibly semi-liquid, maybe both. It's the fuel velocity that really jacks up the Isp.

So I think you'd be better off getting someone else to defend it because I'm just as stumped by these results.

Correction: It should be denser, because of the high thrust.

I need to sleep...

 

[Kalec]

Hey you're preaching to the choir. What do think I'd been saying these past few pages? I didn't believe it'd be this high either.

Then again, 150t in 115s is a really fast burn. If I were to guess, either the propellant is highly reactive or the fuel mix is very thin, possibly semi-liquid, maybe both. It's the fuel velocity that really jacks up the Isp.

So I think you'd be better off getting someone else to defend it because I'm just as stumped by these results.

Dude, you guys are debating pages and pages, even longer than CASC newsletter.

From my understand, the exhaust velocity is determined by the nozzle not the propellant, but the nozzle has a yield point otherwise propellant wouldn't be effectively burnt out. So I think the Isp is actually determined by what propellant you are using not the nozzle, at least in the first stage.

It really don't matter between HTPB, NEPE or GAP here. The Isp is determined by AP, HMX, RDX or CL-20, TKX-50, ADN. NEPE propellant has higher Isp because it allows manufacturer to mix more HMX into the propellant without sacrificing mechanical properties. And GAP is even better as it has more energy in bind agent than NEPE. The propellant has more Isp when you put more oxidizer or aluminum in it, so the cost of propellant is limited by how many expensive oxidizer you can afford.

 

[Quickie]

Dude, you guys are debating pages and pages, even longer than CASC newsletter.

From my understand, the exhaust velocity is determined by the nozzle not the propellant, but the nozzle has a yield point otherwise propellant wouldn't be effectively burnt out. So I think the Isp is actually determined by what propellant you are using not the nozzle, at least in the first stage.

It really don't matter between HTPB, NEPE or GAP here. The Isp is determined by AP, HMX, RDX or CL-20, TKX-50, ADN. NEPE propellant has higher Isp because it allows manufacturer to mix more HMX into the propellant without sacrificing mechanical properties. And GAP is even better as it has more energy in bind agent than NEPE. The propellant has more Isp when you put more oxidizer or aluminum in it, so the cost of propellant is limited by how many expensive oxidizer you can afford.

From my understand, the exhaust velocity is determined by the nozzle not the propellant, but the nozzle has a yield point otherwise propellant wouldn't be effectively burnt out. So I think the Isp is actually determined by what propellant you are using not the nozzle, at least in the first stage.

That is kind of like a rude response. Did you really read through all the posts?

We already stated that the Isp is way bit too high at the start of the discussion. I tried to explain it by saying the stated 500-ton force is the max thrust, not the average thrust which should be lower. This is then confirmed by Taxiya's post of a chart of the thrust profile of the space shuttle solid booster.

There is no way the Isp is that high. That is the Isp of a staged combustion LOX/Kerosene engine. At best a solid rocket should be in the high 200s.

Yeah that's what I thought at first. I had it pegged at the mid-270s intitially.

But then I tried to follow Quickie's math and it seems to check out -

View attachment 100937

m-dot = 150,000kg / 115sec = 1,304.35kg/s

View attachment 100938

Pretty impressive stuff if that's really what CSAC has managed to achieve.

As already explained, the stated 500-ton force is the max value of the whole test.

If we were to estimate the average thrust to be about 70% of the max (The Space Shuttle booster chart may be used to give a good estimate here), I think the calculated Isp would come out to be more reasonable.

 

[by78]

Long March 9 model at Zhuhai Airshow.

 

[by78]

80-ton re-usable LOX/Methane engine.

 

[by78]

 

[by78]

Enclosing the Mengtian module in payload fairing.

 

[Kalec]

CASC tested YF-130, the 500t staged kerolox engine today.

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And here comes 80t gas-generator metholox, YF-209.

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