Quote from: Nomic on 01/20/2017 10:09 amAlways liked the idea of dual mixture ratio hydrolox for any one crazy enough to try reusable SSTO. Start off running LOX rich, switch to fuel rich at the equivalent of staging. The J2 did this. Sacrificing Isp to improve payload seems counter intuitive but it's the concept of putting less KE into the remaining mass of the vehicle.
Always liked the idea of dual mixture ratio hydrolox for any one crazy enough to try reusable SSTO. Start off running LOX rich, switch to fuel rich at the equivalent of staging.
BTW Aerojet seemed to have been quite fond of this tactic in many of their design studies. I think they went as high as 12:1.
Better to use nitrogen in that case instead of oxygen. There's like 4x as much nitrogen, and it's nearly inert.
Quote from: john smith 19 on 01/20/2017 07:35 pmQuote from: Nomic on 01/20/2017 10:09 amAlways liked the idea of dual mixture ratio hydrolox for any one crazy enough to try reusable SSTO. Start off running LOX rich, switch to fuel rich at the equivalent of staging. The J2 did this. Sacrificing Isp to improve payload seems counter intuitive but it's the concept of putting less KE into the remaining mass of the vehicle.I just had another look at the mixture-ratio data in Apollo by the Numbers, and I'm a bit puzzled. First of all, for each burn of each stage, two O/F figures are given: one corresponding to the start of the burn and the other supposedly the change during the burn. But the latter figure, however, is generally larger than the former, which implies that O/F more than doubled during the burn. That seems very unlikely, so I assume the second number is actually the O/F at shut-down.With that assumption, it seems that the O/F gets leaner (higher) during most if not all burns. But don't we expect just the opposite, namely getting richer (more hydrogen) to increase specific impulse as the burn proceeds?It also seems to be the case that the second S-IVB burn is on average leaner than the first (I say "seems to be," because the average during the burn isn't necessarily the average of the intial and final values). Rocket-equation-wise, this too seems counterintuitive, though maybe it's a way of maximizing stage performance in the face of hydrogen boil-off.What gives?QuoteBTW Aerojet seemed to have been quite fond of this tactic in many of their design studies. I think they went as high as 12:1.That sounds interesting -- do you have any pointers as where to find these studies?BTW, attached is a Rocketdyne study of a variable-mixture-ratio lox-hydrogen engine.
For a reusable TSTO can the upper stage and the capsule's service module be merged?
\Unfortunately IRL those also count as "benefits." QuoteWell, sure, but more for NASA/government projects... I was asking more about the lack of SSTO projects among the "new"/commercial space companies.QuoteYou have the problem backwards. If you don't have the "Bank of Elon" behind you you have to convince financiers this is a good idea. They will do due diligence by plugging in that vehicle size into their aerospace cost model, where cost is roughly proportional to GTOW. But where does this assumption ("cost proportional to GTOW") come from? It seems not to be true, or at least to be avoidable.QuoteBeing able to match a TSTO ELV in payload fraction matters. No historical SSTO had managed this. But why is payload fraction the relevant metric? Cost per flight is more important.Historically this has been something of a circular argument. An SSTO with 1/3 to 1/4 of the payload of a TSTO ELV has to fly 3-4x more often.
Well, sure, but more for NASA/government projects... I was asking more about the lack of SSTO projects among the "new"/commercial space companies.QuoteYou have the problem backwards. If you don't have the "Bank of Elon" behind you you have to convince financiers this is a good idea. They will do due diligence by plugging in that vehicle size into their aerospace cost model, where cost is roughly proportional to GTOW. But where does this assumption ("cost proportional to GTOW") come from? It seems not to be true, or at least to be avoidable.QuoteBeing able to match a TSTO ELV in payload fraction matters. No historical SSTO had managed this. But why is payload fraction the relevant metric? Cost per flight is more important.
You have the problem backwards. If you don't have the "Bank of Elon" behind you you have to convince financiers this is a good idea. They will do due diligence by plugging in that vehicle size into their aerospace cost model, where cost is roughly proportional to GTOW.
Being able to match a TSTO ELV in payload fraction matters. No historical SSTO had managed this.
Variations of mix ratio are too small to be meaningful.
2nd stage reuse is always simpler than reuseable SSTO or shuttle architecture on the same conditions. TPS of large rockets is a great pain.
BTW Aerojet seemed to have been quite fond of this tactic in many of their design studies. I think they went as high as 12:1. It's another of those "low hanging fruit" you should be looking at if you're going to do SSTO. IRL the joker is you run oxidizer rich, which makes US developers very nervous, and you transition through the stoichometric (maximum temperature) range. Obviously if you can scan through that range quickly enough without triggering combustion instability that's not a problem but I don't think anyone's ever really done that.
Aerojets work on the OTV engine touched on using mixture rations up to 13:1, some of the problems discussed on page 108 onwards, and solutions including gold platting the MCC (page 115) and platinum baffles. Idea behind the high mixture ratios was to use Lunar regolith as source of ISRU LOX. Don't think anyone has ever run an engine at these sort of mixture ratios though.
That's the neat thing about Thrust Augmented Nozzles. The expansion ratio for the main engine can be huge, but because you're adding a bunch of propellant downstream when at low altitudes, you don't get flow separation.
Ummm.. the reentry system of Sea Dragon was the pressure vessel. Never flew, obviously, but I don't think anyone really argues there's much doubt it wouldn't work. Big fluffy things reenter intact even when you don't want them too. It's small things that need painful TPS.
One reason I think a SSME with TAN might be the prefect engine for a simple near term SSTO.
Right now the nozzle is a compromise for sea level and vacuum performance that uses a lot of tricks to prevent flow separation.
But a TAN nozzle could have an expansion ratio that is better optimized for both sea level and vacuum operation.
Quote from: john smith 19 on 01/21/2017 09:04 am\Unfortunately IRL those also count as "benefits." Well, sure, but more for NASA/government projects... I was asking more about the lack of SSTO projects among the "new"/commercial space companies.
\Unfortunately IRL those also count as "benefits."
QuoteYou have the problem backwards. If you don't have the "Bank of Elon" behind you you have to convince financiers this is a good idea. They will do due diligence by plugging in that vehicle size into their aerospace cost model, where cost is roughly proportional to GTOW. But where does this assumption ("cost proportional to GTOW") come from? It seems not to be true, or at least to be avoidable.
QuoteBeing able to match a TSTO ELV in payload fraction matters. No historical SSTO had managed this. But why is payload fraction the relevant metric? Cost per flight is more important.
Yeah but it ought to be able to fly >10x more often - early on. Once the market expands and the technology matures >>100x. 1 flight per vehicle every day or two doesn't really seem that unreasonable... IF you computerize the thing to the max, so there's no human inspections between flights unless the vehicle's computers flag a problem.
High flight rate and low cost per flight will be dependent on low labor costs and therefore heavily automated systems. Also probably not using existing ground infrastructure.