FIRST STEPS TOWARD SPACE - Smithsonian Institution Libraries

236 SMITHSONIAN ANNALS OF FLIGHT
In contrast to the comments from the Austrian Department
of Defense, the test results prove the outstanding feasibility
of the propellant combination. Future engine developemnt
tests will be run alternately with gaseous and with liquid
oxygen, the latter ones only as complementary tests.
In spite of fuel-rich combustion and correspondingly
low combustion gas temperatures, the fuel
coolant temperatures in the cooling passages of SR-
11 and SR-12 went up to 450° C. To find the causes
of this temperature rise, SR-13 was equipped with
separate cooling passages for chamber and nozzle.
The combustion chamber was made of 6/8 mm (id/
od) copper tubing wound to shape for water and
later for lox as coolant and the thrust chamber of
2/4 mm (id/od) copper tubing was wound to
shape for fuel as coolant. The coolant velocities
ranged from 10 to 15 m/sec. After eight tests with
SR-13, Sanger wrote in his log book on 18 September
1934:
The current situation is as follows:
The combustion chamber made of carefully wound copper
tubing, faultlessly connected at both ends and brazed tight
on the outside with bronze wire, withstands all loads with
both water and fuel as coolants.
However, the same thermal design does not work at the
throat. Thrust chambers, whether cooled by fuel or water
and whether made of copper or steel, are burning through
near the inlet and in the throat area. Fuel-cooled copper
nozzles behave best and water-cooled steel nozzles worst. But
it seems that burn-through can be avoided by smooth surfaces
inside the nozzle. Obviously, the rough surfaces in the throat
area greatly increase the combustion gas-to-wall heat flux up
to 1.7 PS/cm 2 as measured under oxygen-rich combustion.
Convection heat transfer seems to be important. . . . The
wall thickness, especially that of copper tubing, is less important
for the required heat flow rates across the wall. Of
decisive importance is the ratio of combustion-gas heat flow
to wall and wall-to-coolant heat flow, as determined by the
boundary layers on each side.
The hot-side heat transfer is determined by (1) radiation
and (2) convection. Convective heat transfer peaks especially
around the throat area because of gas velocity and density.
The coolant-side heat transfer is determined by convection
and increases with coolant flow velocity and temperature
difference between coolant boundary layer and coolant bulk.
Equilibrium between the heat flows on both wall sides
must be obtained at wall temperatures compatible with the
wall material.
During a number of previous thrust-chamber tests run
within the allowable wall temperature range, the hot-side
heat flow indeed exceeded that on the coolant side.
In the first place one must try to keep the equilibrium wall
temperature below the melting temperature of customary
metals, such as copper or bronze.
A. The hot-side heat flux must be minimized.
1. Eliminate all heat transfer caused by flow perpendicu­
lar to the wall (minimum turbulence, no perpendicular
flow; walls as smooth as possible).
2. Reject radiative heat by reflective surfaces.
3. Minimize heat-exposed surface areas by avoiding protrusions,
bends, etc.
4. Maximize combustion-gas boundary-layer temperature
to reduce temperature difference of combustion gas
bulk and boundary layer (reduces radiative and convective
heat flow).
5. Reduce combustion-gas density (reduces convection).
B. The coolant-side heat flux must be maximized.
1. Provide very high coolant flow velocities for better
heat transfer.
2. Increase the heat dissipating surface areas by cooling
fins (for example, by internally grooved tubing, according
to Sztatecsny).
3. Provide for coolant flow mainly perpendicular to wall
(direct impingement, highly turbulent).
4. Increase coolant density (high pressure for gases, metallic
powder added to diesel fuel).
5. Increase temperature difference on coolant side by use
of cryo-coolants (for example, lox).
6. Increase coolant boundary-layer temperature for reasons
identical to those on the hot side.
If these steps necessitate uneconomical efforts or fail to obtain
wall equilibrium temperatures below 1000° C, then hightemperature-resistant
nozzle materials have to be used.
Based on this knowledge, SR-14 was built and
fired on 4 October 1934. During the second test, it
produced a thrust of 2 kg and obtained an exhaust
velocity of around 3000 m/sec for a chamber pressure
of 16 atm and highly fuel rich combustion;
the steady-state run-time, however, was not determined
very accurately. During a later test, with
30% fuel-rich combustion, a chamber pressure of
22 atm and a steady-state run duration of 63 sec, a
thrust of 4.5 kg and an exhaust velocity of 2760
m/sec were obtained. During both tests, the temperature
of the fuel and water coolant stayed
within allowable limits and the rocket engine was
undamaged.
Regrettably, the testing of this model was limited
to five runs. On 17 October 1934, Professor Rinagl
forbade further testing because the noise allegedly
annoyed the neighbors. The 135th and also the last
test, on 23 October 1934, was a demonstration run
for Count Max von Arco-Zinneberg; the test operation
was smooth and no hardware was damaged.
Based on his test experience, Sanger recorded the
following notes as patent claims:
1. High-pressure combustion chamber characterized by ducting
the propellants around the chamber in such a way
that they enter it in a preheated condition and cool the
chamber walls.