Handbook of Turbomachinery Second Edition Revised - Ventech!

normally be considered a very small space. This situation has intensified over
the years, since the general desire has been to place an engine of a given thrust
in a progressively smaller volume. One may also view this as continuing to
increase the thrust of a constant volume engine. In any event, turbomachine
designers are confronted with the problem of displacing ever-increasing
quantities of propellant against increasing pressures within a volume that is
decreasing. The specific problems inherent with these trends are covered in
more detail in subsequent sections of this chapter.
To illustrate on a semiquantitative basis the trends in machine design
over the last 40 years, we consider some particular examples.
In the early 1950s, the Army Ballistic Missile Agency (ABMA) called
for the development of a 200-mile-range battlefield missile. This missile
came to the be the Redstone. The vehicle was, in almost every respect, an
Americanized V-2, including the use of liquid oxygen (LOX) and ethyl
alcohol as propellants. The engine was required to generate 78,000 lb of
thrust at a chamber pressure of approximately 300 psia, or almost 50%
greater than the V-2. To support the thrust chamber operation, the
turbopump was required to move 150 lb/sec of ethyl alcohol fuel (1,190 gal/
min) along with 205 lb/sec of LOX (1,290 gal/min). The head rises required
of the fuel and oxidizer pumps were approximately 1,140 ft and 600 ft,
respectively. Both pumps were arranged on a common shaft, which was
driven at a rotational speed of 4,700–4,800 rpm. With efficiencies of both
pumps on the order of 70%, approximately 740 hp of shaft power was
required. The turbine configuration was a two-row velocity-compounded
unit, also known as a Curtis staged turbine. Arranged in the middle of the
turbopump shaft, the turbine’s pitch diameter was approximately 20 in. The
drive fluid employed was decomposition products of concentrated hydrogen
peroxide, again similar to its German counterpart. These products of
decomposition were ducted to the turbine inlet manifold at a total inlet
pressure of slightly less than 400 psia at a temperature of nearly 700 8F. The
turbine expanded the working fluid through a pressure ratio of 22, with the
exhaust products leaving the turbine at just slightly higher than sea-level
pressure. This pressure ratio is of the order of magnitude of the majority of
rocket engines that operate on the so-called gas generator cycle. These
conditions provide high work per pound of turbine drive fluid, although
extracting this energy efficiently is problematic, as evidenced by the unit’s
efficiency of just under 40%.
A rather useful (though not all-important) parameter by which rocket
turbomachinery performance is gauged is by the use of a ‘‘power density’’
term. This is simply the shaft power supplied by the turbopump per unit of
machine weight. In the case of the Redstone engine turbopump, this term
computes to slightly over 2 hp/lb.
Copyright © 2003 Marcel Dekker, Inc.