Is it feasible to use either solid or bipropellant liquid rocket engines to lift off from the surface of Venus?

"Reference datum" air pressure on Venus is around 9.3 MPa; even on the top of, say, Maat Mons, the pressure would be around 5 MPa. Some modern engines have chamber pressures above 20 MPa (e.g. SSME, RD-170), so it seems possible for a rocket engine to work.

The shuttle SRBs produce substantially less pressure, around 4.3 MPa; is it practical to get a solid rocket booster up above 10MPa?

Due to the very high ambient pressure, I'd expect the expansion nozzle to be short and specific impulse to be poor. Can the performance specifications of a liquid bipropellant rocket engine optimized to exhaust into pressures from 1 to 5 MPa, using similar turbomachinery to an existing engine, be estimated? Ideally I'd like to get an estimate of the specific impulse curve of such an engine from 10 MPa to 0.1 MPa ambient.

Feel free to handwave away the thermal issues involved in maintaining liquid propellant tankage while waiting to launch from the Searing Black Calm. The goal here is to be able to simulate a rocket ascent from the surface of Venus to stable orbit, in order to answer another question on the site.

I’m well aware this is not going to be the most practical way to ascend to orbit from Venus. I’m not asking if it’s optimal or practical. Please constrain discussion to rocket engines only.

RPA crashes promptly when I ask it to compute performance at 60 bar exit pressure.

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    $\begingroup$ I looked quickly at cutting off an SSME nozzle because I had those curves...turns out that the SSME is right about 10 MPa at the throat so a stub nozzle indeed. For exit plane pressure of 1.7 MPa the specific impulse would be around ... 100. $\endgroup$ Feb 22, 2020 at 2:46
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    $\begingroup$ The structural limit on a chamber is the gauge pressure: the amount above ambient. So a chamber on Venus can still be the same pressure above ambient as on Earth. $\endgroup$ Feb 22, 2020 at 5:08
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    $\begingroup$ @RussellBorogove Yep! From my experience at JPL, sometimes the question-askers, when they have a favorite approach in mind, pose the question in such a way as to rule out the obvious best approach. $\endgroup$ Feb 22, 2020 at 5:26
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    $\begingroup$ @BobJacobsen Yes. But given a propellant chemistry, the axial temperature profile through the chamber and the throat is a function of the absolute pressures, so increasing the chamber and throat pressures will increase the down-chanber and throat temperatures. That's why I mentioned the cooling system, not the structural design. $\endgroup$ Feb 22, 2020 at 5:32
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    $\begingroup$ I am thinking the only way this could be remotely practical is with an aerospike nozzle - I tried doubling the SSME Pc with a conventional nozzle and it was still terrible performance. I'm trying to gen up up some numbers for an "ASSME" - no spaces in the acronym please - but not sure where I will end up. $\endgroup$ Feb 22, 2020 at 17:26


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