Performance at simulated high altitudes of a prevaporizing annular turbojet combustor having low pressure loss Page: 3 of 44

                                
NACA RM E56114

prevaporizer capacity in the experimental combustor for operation with
low-temperature fuel (800 F), in most aircraft applications fuel is
delivered to the combustor at temperatures in the range of 2500 to 3500 F.
At the test conditions investigated the combustor exhaust-temperature
profile followed the pattern generally desired at the turbine position.
INTRODUCTION
High-altitude operation of turbojet engines is frequently accompanied       C
by serious losses in combustion efficiency. It has been shown that at
high altitudes preheating the liquid fuel before injection into the com-
bustion chamber increases combustion efficiency significantly; use of a
gaseous fuel results in even greater gains in efficiency (ref. 1). Re-
search on an experimental turbojet combustor that incorporated a liquid-
fuel prevaporizer is reported herein.
A prevaporizing combustor incorporating a fuel-system, designed to
supply liquid fuel at sea level and low altitudes, preheated fuel with
an increasing vapor content up to a simulated altitude of 56,000 feet,
and 100-percent vaporized fuel at higher altitudes, is described in ref--
erence 2. The prevaporizing coils of this combustor were located at the
downstream end of the primary zone prior to the entry of secondary air.
This location was chosen for two reasons:  (1) to avoid quenching effects
in the burning zone due to cold prevaporizer walls, and (2) to minimize
pressure loss due to the coils by placing them in a low mass-flow region.
The combustor operated with a high combustion efficiency. While the
pressure losses were of the same-magnitude e_ncountered in current produc-
tion engines, redesign of the combustor liner was undertaken to explore
the possibilities of reducing pressure losses.
The reduced pressure-loss combustor had an air-entry pattern similar
to that of model 30 of reference 2 which incorporated the prevaporizing
system described in reference 2. Design modifications to reduce the
pressure loss were directed toward improvement of the combustor-liner
geometry with respect to the combustor housing and provision for adequate
open air-entry area. The fuel manifold and the upstream primary zone
walls were integrated into an annular, symmetrical wedge arrangement
that improved the entrance air diffusing passages. Modification of the
primary zone to obtain low pressure loss resulted in decreasing the air-
entry orifice coefficients, which in turn reduced the mass flow into the
primary region. The required primary flow was obtained by the use of
special airscoops that separated a small fraction of the air from the
mainstream, and then admitted this air fraction in a predetermined
manner.