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## RD-180 Engine Diagram

RD-180 flow diagram

This is a bit different from the NK-33 done previously, but it’s still a full flow oxidizer rich staged combustion lox-kerosene engine.

It has no gears and no flexible shaft coupling between the pumps like the NK-33, making it a real one axis engine – except that it has separate booster pumps at the engine inlet. The fuel booster pump is powered by the fuel tapoff after one main pump stage and the oxidizer booster pump by the turbine exhaust. The starting is also different, but I omitted the starter hardware from the already complicated diagram, as it’s connected to many places – the main chamber, the gas generator and the first main fuel pump inlet. Also various valves, controllers and the tank heat exchanger are left out. And naturally I left out the other chamber and nozzle as well.

Both the RD-180 and the NK-33 have the same amount of pump stages – 3 for fuel entering the gas generator and 2 for fuel entering the main chamber and 2 for the oxidizer (all enters the gas generator).

Perhaps it can be thought, simplified, that the boost pumps are only hydraulically and not axis coupled to the main shaft system, and hence both can be better optimized to their environment (like lower rpm for the boost pumps) and hence the system can reach higher pressures than the NK-33, where the two oxidizer pump stages are on the same shaft. Or then it’s the later materials or more advanced pump design, after all the engines have some ten to twenty years between them.

Source for the drawing and explanation is this patent no. 6226980. Also, lpre.de has awesome pictures of the hardware, including the shaft with all the pump stages included. I assume it’s machined from one solid piece. Also pictures of the pipe stack injector / mixer and more diagrams of the engine operations. I don’t know much Russian (having finished half a course back somewhen), but if you know most of the cyrillic alphabet (helps if you know math as it’s very close to Greek), it’s practically quite easy to read as there are so many loan words – gasogenerator should not be a mystery to anyone. 🙂

The workhorse Soyuz RD-107 and RD-108 engines are completely different as they use a hydrogen peroxide gas generator design – very old-fashioned – but the RD-0124 used on the more modern Russian upper stages is the third interesting kerosene staged combustion engine that might become even more actual if Orbital are going to use it as a second stage engine on their Taurus II (currently they are moving on with solids). The fourth staged combustion engine is the RD-120 that’s bigger than RD-0124 and is used on the Zenit second stage. And then there’s the often overlooked forefather of staged combustion, Proton’s RD-253 / RD-275, that uses hypergolic propellants. The RD-0120 hydrogen engine of Buran / Energia is interesting as a comparison to the similar SSME. So there’s still plenty of study subjects in the Russian engine families.

## The US Air Force Tries To Do Reusables

But is not a real RLV program. It’s just a narrow test for one technology. Hence I think naming it Reusable Booster System Pathfinder is misleading.

## Overspecification

They overspecify the problem by requiring a glide landing. Why is it superior to powered landing? At the moment, there’s no clear reason to believe it is! Both need to be developed further to understand their advantages and drawbacks. To my knowledge, there have been only six liquid rocket VTVL prototype manufacturers so far: McDonnell Douglas, JAXA (who was the contractor?), Armadillo Aerospace, Blue Origin, Masten Space Systems and Unreasonable Rocket. Only a few of those have flown to higher than a few hundred meters. The design and operations space is mostly totally unexplored.

Nevermind the large number of other alternatives to boostback. Jon Goff had a recent “lecture series” about these.

I understand that this is just one program, but this should not gain the status of the reusables approach of the air force – stuff like that easily happens.

## Master Design Fallacy

They also discard evolution and competition – instead just requiring a single masterfully designed prototype before something operational. Sure, this is much better than starting a multi-billion dollar program without a first lower cost prototype, but nevertheless, it sucks. Somebody brief them on newspace! Rand Simberg, Monte Davis, Jonathan Goff, Clark Lindsey, or one of the numerous people who get it. Or one of the prominent company leaders: John Carmack, Jeff Greason, David Masten.

## An Ideal Program

Just specify some boost delta vee points and let companies demonstrate progress towards that. A popup tailflame lander would perhaps give more vertical velocity while some good glider or even a booster that has engines for cruising back could boost far down range to give lots of horizontal velocity. There ain’t a clear winner – there might not even be and multiple approaches would have their uses.

## Lunar Lander Challenge 2009

Masten and Unreasonable are still flying for second place I think (I’m not 100% clear on the rules) today!

Spacetransportnews is the place to watch all this. (Or it has the links collected.)

It’s historical in a sense. These rockets will serve as the basis for reusable sounding rockets, possibly high altitude tourist vehicles and later orbital system lower or upper stages. When the operations are routine and landings safe, the cost per flight goes down orders of magnitude, compared to ordinary rockets.

A new era for rocketry is dawning.

## Optimum Rocket Cruise

With some caveats. 🙂 Let’s assume a rocket is launched, and accelerates to constant speed v_c. Then it stays cruising at this speed and at a constant altitude. Landing is disregarded.

### The cruise

We must modify the rocket equation slightly for the cruise: $\frac{-dm}{dt} v_{ex} = F = \frac{gm(t)}{L/D}$ dm/dt is mass flow, v_ex is effective exhaust velocity, F is the thrust, g is the gravitational acceleration 9.81 m/s^2, m(t) is the mass as function of time, L/D is the lift to drag ratio. If we use the $\Delta t = x/v_c$ for time, (x is the cruise distance) we can integrate it from start to final mass just like the rocket equation and get the cruise mass ratio: $R_{mc} = e^{\frac{xg}{v_c v_{ex}L/D}}$ Notice how with increasing cruise speed, the required mass ratio for cruise is lessened. This is because less time is spent in the air and thus the gravity losses are lessened.

### The acceleration

But we have to take into account the acceleration to cruise speed as well, which requires some mass ratio as well. $R_{ma} = e^{\frac{v_c}{v_{ex}}}$ We don’t take into account the distance traveled during acceleration, or lift, as the acceleration is a relatively short time and distance phenomenon with rockets that easily optimize to have high T/W.

### Total effect

Now, for the total required mass ratio, we multiply the two mass ratios. Then we search for the minimum total mass ratio by derivating it and searching for the zero point. We get: $v_c = \sqrt{\frac{xg}{L/D}}$the optimum cruise speed (smallest mass ratio) Notice how the exhaust velocity cancels out, the optimal speed doesn’t depend on it.

### More considerations

If I calculated right, for a 6000 km transatlantic rocket powered flight with a lift to drag of 7, the best cruise speed for minimum mass ratio is 3 km/s. If you go slower, you waste fuel by hanging in the air, if you go faster, you waste fuel by accelerating too much. I think that’s about Mach 9 at some altitude. This didn’t take into account the deceleration: faster cruise speed takes some advantage there! Even if you shut the engine, it glides further. In real life there are multiple issues:

• acceleration takes time and distance too
• engine T/W size has an effect as well
• there is varying mass during flight .which reduces lift needs with time o which in turn effects L/D as you go higher or reduce AoA .which also requires throttling
• And a million other things.

Also L/D 7 is probably much too good. Oh, and in the transatlantic case, mass ratio required with exhaust velocity 3 km/s would be 7.