Bi-propellant Liquid Fuel Engine Pressures Before Thrust Chamber

[Ozen]

TL;DR Summary

What does the pressure gradient look like at turbopump exit and before entering the thrust chamber?

Been reading Rocket Propulsion Elements 9th Edition and got approval from my university to design a bi-propellant liquid fuel rocket engine for my senior design project, and I've been understanding everything so far but I haven't quite found an answer to how the pressure works throughout the plumbing before the thrust chamber.

The turbopump raises the pressure from 2-4 bar inlet (from the fuel/oxidizer tank) but what are the constraints on the exit pressure? I know the turbopump needs designed to prevent cavitation and have sufficient pump head, but is there a specific requirement typically for the outlet pressure, or is it solely dependent on the pump head? Also, I am a bit confused on pump head for these in rockets. The turbopump is below the fuel/oxidizer tanks, why would the head matter when it naturally will want to suck out the propellant? Wouldn't the outlet pressure be a more useful measure for the pump?

After the turbopump, for the fuel, it will then be used in regenerative cooling and film cooling. Surely the pressure here would be critical to ensuring enough coolant is applied/fuel lost?

The rest of the fuel will go into the corresponding injector assembly, which usually has a large opening where the pressure will drop. And there will be pressure changes throughout the injector, for this instance a coaxial swirl injector.

Then it enters the thrust chamber and ignites, raising the pressure to around 7 MPa. I understand the initial condition and end conditions, but the intermediary surely has an effect on overall performance, right? Are there any books that go into the criteria of these components? I have on my backlog to read next Liquid Rocket Thrust Chambers: Aspects of Modeling, Analysis, and Design; Liquid Rocket Engine Turbopump Inducers; Modern Engineering For Design of Liquid-Propellant Rocket Engines; & Arakaki High Speed Centrifugal Pump Design for Rocket Engine 2017.

PS: I am well aware of the safety concerns and hazards and the overall complexity of designing a liquid fuel rocket engine.
PPS: The engine configuration I described is one using an electric turbopump instead of a gas generator cycle and turbine; this reduces complexity and engine cost at the expense of additional weight.

 

[jrmichler]

what are the constraints on the exit pressure?

The pump exit pressure has one constraint - it must be high enough to deliver the liquid. The liquid is delivered into the combustion chamber, which is at a pressure. There will be a pressure drop at the nozzle, and pressure drop in every line and fitting between the pump and the nozzle. You start at the combustion chamber, follow the fuel path back to the pump, and add up every pressure drop on that path. That gives you the pressure at the pump discharge.

Similarly, the fuel and oxidizer tanks each have a pressure. The lines, fittings, and valves in the flow path from the tanks to the pump inlets all have pressure drops. The tank pressures minus the pressure drops give you the pressure at the pump inlets.

You also need to calculate the net positive suction head (NPSH) at the pump inlets.

Then you are ready to start designing the pump. A good general reference on pumps is Centrifugal and Axial Flow Pumps, 2nd Edition, by Stepanoff: https://www.amazon.com/dp/0894647237/?tag=pfamazon01-20. There are other pump books that may be better, this is the one that I have. This figure, from that book (page 76), shows how the relationship between flow rate, RPM, and pump head affects the design of the pump.

Also, be sure to calculate the power required to drive the pump.

A fun project, but you will need to carefully constrain your efforts in order to finish in time. Enjoy!

 

[Benjies]

After the turbopump, for the fuel, it will then be used in regenerative cooling and film cooling. Surely the pressure here would be critical to ensuring enough coolant is applied/fuel lost?

One thing I was taught in my rocket propulsion course is that oftentimes the regen cooling channels are designed such that the liquid that enters the cooling channels increases in temperature to a point just before phase-changing to a supercritical fluid. This means that mass flow can be fixed through the regen cooling channels.

If you study an engine at an assigned thrust value- say, nominal thrust seen on a test stand- then this thrust occurs at a fixed chamber pressure. Assuming you are referring to an engine that is regen cooled, rather than dump cooled, this means that this chamber pressure only occurs at a specific mass flow mix between the regen cooling, and the direct injection to the combustion chamber (aka, you know the mass flow of fuel that bypasses cooling).

I understand your post has a multitude of questions to address, but specifically to your question of coolant mass flow, I hope this sheds some light on the mass flow that is often "assigned" to the cooling channels: Just enough to allow the coolant flow to not phase-change to a supercritical fluid. This minimizes coolant pressure losses and maximizes engine chamber pressure, thus maximizing thrust. If your question was more directed to the NPSH, or pressure rise and work required by the pump, then in a similar manner, for a given thrust value, this design parameter minimizes the power requirements for your pump.

Of course, as you've stated, there are many variables that inform the power requirements of your pump- I've only outlined how engineers optimize one design characteristic to minimize load on the pumps, here.