Advanced Rocket Nozzles
Advanced Rocket Nozzles
Advanced Rocket Nozzles
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Gerald Hagemann*
DLR, German Aerospace Research Center, Lampoldshausen 74239, Germany
Hans Immich†
Daimler – Benz AG, Munich 81663, Germany
Thong Van Nguyen‡
GenCorp Aerojet, Sacramento, California 95813
and
Gennady E. Dumnov§
Keldysh Research Center, Moscow 125438, Russia
Several nozzle concepts that promise a gain in performance over existing conventional nozzles are
discussed in this paper. It is shown that signi cant performance gains result from the adaptation of the
exhaust ow to the ambient pressure. Special attention is then given to altitude-adaptive nozzle concepts,
which have recently received new interest in the space industry. Current research results are presented
for dual-bell nozzles and other nozzles with devices for forced ow separation and for plug nozzles with
external freestream expansion. In addition, results of former research on nozzles of dual-mode engines
such as dual-throat and dual-expander engines and on expansion – de ection nozzles are shown. In general,
ow adaptation induces shocks and expansion waves, which result in exit pro les that are quite different
from idealized one-dimensional assumptions. Flow phenomena observed in experiments and numerical
simulations during different nozzle operations are highlighted, critical design aspects and operation con-
ditions are discussed, and performance characteristics of selected nozzles are presented. The consideration
of derived performance characteristics in launcher and trajectory optimization calculations reveal sig-
ni cant payload gains at least for some of these advanced nozzle concepts.
a
Table 1 Performance losses in conventional rocket nozzles
Vulcain 1, SSME,
Losses % %
Chemical nonequilibrium 0.2 0.1
Friction 1.1 0.6
Divergence, nonuniformity of exit ow 1.2 1.0
Imperfections in mixing and combustion 1.0 0.5
Nonadapted nozzle ow 0 – 15 0 – 15
a
Other loss sources also shown in Fig. 1.
formance during the ascent of the launcher owing to their xed induce an oblique shock wave near the wall, which leads to a
geometry. Signi cant performance losses are induced during recompression of the ow.
the off-design operation of the nozzles, when the ow is over- The physical phenomenon of ow separation can be divided
expanded during low-altitude operation with ambient pressures into two simple phenomena.3 The rst is the turbulent bound-
higher than the nozzle exit pressure, or underexpanded during ary-layer separation from the nozzle wall, which is character-
high-altitude operation with ambient pressures lower than the ized by the ratio of the nozzle wall pressure just after the
nozzle exit pressure. Figure 2 shows photographs of nozzle separation, pp, to the nozzle wall pressure just before separa-
exhaust ows during off-design operation. In the case of over- tion, psep. This pressure ratio is referred to as the critical pres-
expanded ow, oblique shocks emanating into the ow eld sure ratio pcr = p sep / p p = p 1 / p 2. The second phenomenon is
adapt the exhaust ow to the ambient pressure. Further down- connected with the ow in the separated zone, which is char-
stream, a system of shocks and expansion waves leads to the acterized by a minor pressure gradient along the wall. The
characteristic barrel-like form of the exhaust ow. In contrast, analysis of model experimental data on separated turbulent su-
the underexpansion of the ow results in a further expansion personic ows shows that the pressure ratio p cr is equal for
of the exhaust gases behind the rocket. separated ows before an obstacle, and for separated nozzle
Off-design operations with either overexpanded or under- ows.3 For turbulent ows p cr shows a slight dependence on
expanded exhaust ow induce performance losses. Figure 3 Reynolds number, but depends strongly on Mach number. Fur-
shows calculated performance data for the Vulcain 1 nozzle as thermore, investigations of separated ows in rocket engine
function of ambient pressure, together with performance data nozzles showed that p cr is also a function of wall temperature,
for an ideally adapted nozzle. Flow phenomena at different gas composition, and nozzle wall roughness.
pressure ratios p c / pamb are included in Fig. 4. [The sketch with Various approaches to the prediction of ow separation are
ow phenomena for the lower pressure ratio p c / pamb shows a used in industry and research institutes. For example, the Kel-
normal shock (Mach disk). Depending on the pressure ratio, dysh Center’s method for the determination of the ow sepa-
this normal shock might not appear, see, e.g., Fig. 2.] The ration point in rocket nozzles is based on empirical relation-
Vulcain 1 nozzle is designed in such a manner that no uncon- ships obtained for p cr, p p / p w,e, and p w,e / p amb. Using these
trolled ow separation should occur during steady-state oper- relations, the pressure ratio p sep / p amb, the separation point, and
ation at low altitude, resulting in a wall exit pressure of p w,e the wall pressure distribution are determined. Aerojet uses tab-
’ 0.4 bar, which is in accordance with the Summer eld cri- ulated data for the prediction of ow separation from various
terion.8 The nozzle ow is adapted at an ambient pressure of experiments with conical and bell-shaped nozzles.4 At a known
p amb ’ 0.18 bar, which corresponds to a ight altitude of h ’ pressure ratio of chamber pressure to ambient pressure, a lower
15,000 m, and performance losses observed at this ambient and an upper separation pressure ratio follows from the data-
pressure are caused by internal loss effects (friction, diver- base that is different for conical and bell-shaped nozzles. A
gence, mixing), as shown in Fig. 1 and Table 1. Losses in separation criterion used at the European industry and scien-
performance during off-design operations with over- or un- ti c institutes describes the pressure ratio p sep / pamb as function
derexpansion of the exhaust ow rise up to 15%. In principle, of the local ow Mach number at the wall near the separation
the nozzle could be designed for a much higher area ratio to point. This analytical model was derived from various cold-
achieve better vacuum performance, but the ow would then and hot- ring tests of overexpanding nozzles.9
separate inside the nozzle during low-altitude operation with Various numerical simulations of ow problems featuring
an undesired generation of side-loads. ow separation have shown that the numerical prediction of
ow separation with state-of-the-art turbulence modeling re-
A. Flow Separation and Side-Loads sults in good agreement with experimental data.3,10 As an ex-
Flow separation in overexpanding nozzles and its theoretical ample, Figs. 6 and 7 compare experimental and numerical data
prediction have been the subject of several studies in the on wall pressure ratios in two overexpanding nozzles. Fur-
past,3,4,8 and different physical models and hypotheses for the thermore, former experiments of separated ows indicated that
prediction of ow separation have been developed. In strongly there is a random movement of the separation line in over-
overexpanding nozzles, the ow separates from the wall at a expanding rocket nozzles,9,11,12 resulting in the possible gen-
certain pressure ratio of wall pressure to ambient pressure, p w/ eration of side loads. The correct prediction of side-loads with
pamb. The typical structure of the ow eld near the separation models is uncertain, and still a subject of on-going research.
point is shown in Fig. 5, together with wall pressure data. Common side-load models described in Refs. 4 and 9 assume
Separation and the formation of a recirculation zone at the wall that the separation line in the nozzle has a maximum tilt angle
Fig. 5 Flow separation in overexpanding rocket nozzles, wall pressure pro le, and phenomenology.
HAGEMANN ET AL. 623
as compared with two baseline nozzles having the same area Fig. 11 Performance characteristics of a dual-bell nozzle. Per-
ratio as the dual-bell nozzle at its wall in ection and in its exit formance is compared with two baseline bell-type nozzles as func-
plane. Figures 10 and 11 illustrate the performance of a dual- tion of ight altitude (baseline nozzle 1: same area ratio as dual-
bell nozzle as a function of ight altitude in comparison with bell base nozzle; baseline nozzle 2: same area ratio as nozzle
extension).
both baseline bell-type nozzles. (Design parameters of the
dual-bell nozzle are taken from a launcher analysis published
in Ref. 18: Propellants hydrogen/oxygen, r̄ = 6, p c = 200 bar, plied to dual-bell nozzles with their wall in ection to estimate
r t = 0.07 m, wall in ection at area ratio «B = 30, and total the critical pressure ratio yields reasonable results with an ac-
area ratio « E = 100.) The pressure within the separated ow curacy of ’15%. Experimental data obtained at the Keldysh
region of the dual-bell nozzle extension at sea-level operation Center revealed that the critical pressure ratio at the wall in-
is slightly below the ambient pressure, inducing a thrust loss ection is about 15 – 20% less than the critical pressure ratio
referred to as ‘‘aspiration drag.’’ In addition, ow transition of a conventional nozzle, and that usual separation criteria
occurs before the optimum crossover point, which leads to must be corrected for an accurate prediction of the critical wall
further thrust loss as compared to an ideal switchover. The pressure ratio.
nonoptimum contour of the full owing dual-bell nozzle re- Flow transition behavior in dual-bell nozzles strongly de-
sults in further losses at high altitudes. pends on the contour type of the nozzle extension.3,6,18 A sud-
To gain insight into the performance and ow behavior of den transition from sea-level to vacuum operation can be, at
dual-bell nozzles at different ambient pressures, extensive nu- least theoretically, achieved by two different extensions, with
merical simulations with parametrical variations of contour de- a zero wall pressure gradient (constant pressure extension), or
sign parameters were performed.18 An optimized bell nozzle a positive wall pressure gradient (overturned extension). But a
with equal total length and area ratio was used as the reference critical analysis of the transition behavior considering decreas-
nozzle for comparison. As a result the vacuum performance of ing ambient pressures during the launcher ascent revealed that
the dual-bell nozzles has a degradation because of the imper- a considerable time with uncontrolled ow separation within
fect contour, and this additional loss has the same order of the nozzle extension exists even for these types of extensions.
magnitude as the divergence loss of the optimized bell nozzle. The duration of this period can be reduced drastically by throt-
The simulations of sea-level operation also revealed an ad- tling the chamber pressure.18
ditional performance loss because of aspiration drag, which is The main advantage of dual-bell nozzles as compared to
less than 3% for these dual-bell nozzles. This additional loss other means of controlling nozzle ow separation is its sim-
depends linearly on the ambient pressure and, therefore, it is plicity because of the absence of any movable parts and, there-
reduced during the ascent of the launcher. Furthermore, these fore, its high reliability. It is necessary to note that the external
simulations showed that the application of commonly used ow over the vehicle in ight reduces the pressure in the ve-
separation criteria derived for conventional nozzles and ap- hicle base region, where engines are installed. The ambient
HAGEMANN ET AL. 625
pressure triggering the ow transition is the vehicle base pres- conventional bell nozzle with equal design and operation data.
sure instead of the atmospheric pressure at the speci c ight A promising concept for controlled ow separation is therefore
altitude. As the base pressure is lower than the atmospheric temporary inserts, which are removed for vacuum operation.
pressure, the nozzle ow transition occurs at a lower altitude These inserts can be either ablative or ejectible. The inserts
than the one showed in Fig. 11, which slightly decreases the may have the form of a complete secondary nozzle23 (Fig. 12),
ef ciency of the dual-bell nozzle operation along the trajectory. or of small steps attached inside the nozzle wall. In case of
Despite the additional losses induced in dual-bell nozzles, ejectible inserts, a reliable mechanism is needed to provide a
they still provide a signi cant net impulse gain over the entire sudden and symmetrical detachment. In any case, shocks are
trajectory as compared to conventional bell nozzles. Indepen- induced during the transient ejection because the inserts act as
dent combined launcher and trajectory analyses performed by an obstacle in the supersonic exhaust ow. These shocks also
Dasa within the European Space Agency Future European interact with the nozzle walls and increase pressure loads on
Space Transportation Investigations Program (ESA FESTIP ) the wall and local heat uxes. A nonsymmetrical ejection
study19 and at DLR on SSTO vehicles powered with dual-bell would then result in the generation of side-loads. Furthermore,
nozzles result in a signi cant payload gain when compared the danger of a downstream collision with the nozzle wall
with a reference launcher equipped with conventional nozzles. arises because the inserts might also experience a transversal
movement toward the walls.
2. Nozzles with Fixed Inserts Recent hot- ring tests performed in Russia with a modi ed
A trip ring attached to the inside of a conventional nozzle RD-0120 engine, equipped with a secondary nozzle insert, re-
disturbs the turbulent boundary layer and causes ow separa- vealed a signi cant performance gain of 12% during the sea-
tion at higher ambient pressures. At higher altitudes with lower level operation at 100% chamber pressure, compared with the
ambient pressures the ow reattaches to the wall behind the original RD-0120 performance.23 Nominal chamber pressure
trip ring, and full owing of the nozzle is achieved. The tran- of this engine is p c = 206 bar, with an area ratio of « = 85.7.
sition from sea level to vacuum mode depends on the wall These full-scale, hot- ring tests demonstrated the durability of
pressure near the trip-ring location and on the disturbance in- materials, sealings, and the release mechanism and, thus, the
duced by the trip ring. The size of the trip ring is a compromise feasibility of this concept. Figure 12 shows the nozzle hard-
between stable ow separation during sea-level operation and ware and a sketch of the secondary nozzle mounted inside the
the induced performance loss during vacuum operation. In Ref. RD-0120 nozzle.23
9 it is reported that a trip-ring size of 10% of the local bound- The principle performance characteristics of this RD-0120
ary-layer thickness is suf cient to ensure stable ow separa- nozzle with ejectible insert are included in Fig. 13. The nozzle
tion. operation with insert results in a slight performance loss com-
In principle, this concept is similar to the dual-bell nozzle pared with an ideal bell nozzle with the same reduced area
concept with regard to performance characteristics, as shown
in Fig. 11 The sea-level performance of this nozzle concept is
lower than the performance of a conventional bell nozzle trun-
cated at the trip-ring location, because of the aspiration drag
in the separated ow region of the nozzle. Furthermore, at sea
level the bell nozzle with trip rings has even higher divergence
losses than a comparable dual-bell nozzle, because the nozzle
contour upstream of the obstacle differs from the optimal con-
tour for this low-area ratio, as a result of the bell nozzle design
for best vacuum performance. The additional losses induced
during vacuum operation is about 1%, compared with the per-
formance of the bell nozzle without an obstacle. Thus, the
additional losses are comparable to the additional losses in-
duced in dual-bell nozzles. As for dual-bell nozzles, the tran-
sition behavior of this nozzle concept is uncertain, but it will
be even more uncertain than for a dual-bell nozzle with a con-
stant pressure or overturned nozzle extension.
In principle, several altitude adaptations can be achieved Fig. 12 RD-0120 nozzle hardware with removed nozzle insert
with one nozzle by various trip rings, mounted one behind the and sketch of secondary nozzle mounted inside of the RD-0120
other. However, this results in increasing vacuum performance nozzle (photographs taken from Ref. 23).
losses. The trip rings can also be attached into existing nozzles
and, therefore, represent a low-cost concept, at least for test
purposes, with low technological risk. Trip rings have been
demonstrated to be effective for side-load reductions during
the transient startup of rocket engines.20 The main problems
with trip-ring nozzles are not only performance losses, but also
ring resistance in high-temperature boundary layers, the exact
circumferential xing, and the uncertainties in the transition
behavior. These uncertainties might be why active interest in
this nozzle concept in the 1970s, which is documented in var-
ious publications,9,20 22 has disappeared in recent years.
–
Nozzles of this type with extendible exit cones are currently to the previously discussed nozzle concepts, plug nozzles pro-
used only for rocket motors of upper stages to reduce the pack- vide, at least theoretically, a continuous altitude adaptation up
age volume for the nozzle, e.g., at present for solid rocket to their geometrical area ratio. Figures 15 and 16 show two
engines such as the inertial upper stage (IUS), or for the liquid different design approaches for circular plug nozzles, which
rocket engine RL10. The main idea of the extendible extension differ only in the chamber and primary nozzle layout. A con-
is to use a truncated nozzle with low expansion in low- ight ical central body is shown here, which could also be designed
altitudes and to have a higher nozzle extension at high alti- with more sophisticated contouring methods.29,30 Different de-
tudes. Figure 14 illustrates this nozzle concept. Its capability sign approaches include plug nozzles with a toroidal chamber
for altitude compensation is indisputable and the nozzle per- and throat (with and without truncation) and plug nozzles with
formance is easily predictable. The whole nozzle contour in- a cluster of circular bell nozzle modules or with clustered
cluding the extendible extension is contoured for maximum quasirectangular nozzle modules. The latter approach seems to
HAGEMANN ET AL. 627
be advantageous because further losses induced by the gaps ing and thermal expansion (side-loads and thrust vector devi-
between individual modules and the ow eld interactions ations); 2) the cooling of toroidal throat with tiny throat gaps;
downstream of the module exits can be minimized. It has been and 3) the control of combustion instabilities in the toroidal
shown that transition from a round to a square nozzle results combustion chamber. Another plug nozzle con guration is the
in a very small performance loss.31 In principle, the ow eld linear plug nozzle, which is foreseen for the propulsion system
development of a clustered plug nozzle with rectangular nozzle of the RLV X-33 concept.
modules is similar to that of a toroidal plug nozzle, but avoids
the inherent disadvantages of the toroidal plug design regard- 1. Circular Plug Nozzles
ing 1) the control of a constant throat gap during manufactur-
Figure 17 summarizes the principle ow phenomena of cir-
cular plug nozzles with full length and truncated central bodies
at different off-design (top and bottom) and design (center)
pressure ratios that were observed in experiments and numer-
ical simulations. For pressure ratios lower than the design pres-
sure ratio of a plug nozzle with a well-contoured central body,
the ow expands near the central plug body without separa-
tion, and a system of recompression shocks and expansion
waves adapts the exhaust ow to the ambient pressure pamb.
The characteristic barrel-like form with several in ections of
the shear-layer results from various interactions of compres-
sion and expansion waves with the shear layer, and turbulent
diffusion enlarges the shear layer farther downstream of the
throat. The existence of the overexpansion and recompression
Fig. 15 Principle design of plug nozzles, toroidal plug, full processes is inferred from up- and down-variations of plug
length. wall pressure pro les observed in various cold- ow tests and
numerical simulations, and will also be shown later for linear
plug nozzles.
At the design pressure ratio (see Fig. 17, left column, cen-
ter), the characteristic with the design Mach number should be
a straight line emanating to the tip of the central plug body,
and the shear layer is parallel to the centerline. However, for
circular plug nozzles designed with contouring methods pro-
posed in Refs. 29 and 30, no exact one-dimensional exit ow
pro le can be achieved, because both methods use the
Prandtl – Meyer relations that are only valid for planar ows.
Furthermore, nonhomogeneous ow in the throat region,
which is in general not considered within the contour design
process, also in uences the exit ow pro le. The wall pressure
distribution remains constant at pressure ratios above the de-
Fig. 16 Principle design of plug nozzles, clustered plug, 36 mod- sign pressure ratio, i.e., the plug nozzle behaves like a con-
ules, with truncated plug body. ventional nozzle, the loss of its capability of further altitude
Fig. 17 Flow phenomena of a plug nozzle with full length (left column) and truncated central body (right column) at different pressure
ratios pc/ p amb, off-design (top, bottom) and design (center) pressure ratio.
628 HAGEMANN ET AL.
Fig. 18 Performance of numerically simulated plug nozzle with Fig. 19 Performance of numerically simulated plug nozzle with
full-length central body. truncated central body.
adaptation being included. Figure 17 (left column, bottom) il- the ow eld of the toroidal plug nozzle was calculated with a
lustrates the ow eld at higher pressure ratios. Performance numerical method,10 and Fig. 20 shows the calculated Mach-
data of a typical plug nozzle are included in Fig. 18 and com- number distribution in the combustion chamber and nozzle.
pared to a conventional bell nozzle with equal area ratio. The Principal physical processes like expansion waves, shocks, and
design and combustion chamber parameters are included in the the recirculating base- ow region are in good agreement. Both
left column of Table 2.27 the experiment and the numerical simulation show that the
The truncation of the central plug body, which is an advan- ow separates from the conical plug body before reaching its
tage because of the huge length and high structural mass of truncated end. Recent results on experiments with plug nozzles
the well-contoured central body, results in a different ow and published in Ref. 28 also reveal separation of the ow from
performance behavior as compared to the full-length plug noz- the central plug body for conical contours. In contrast, no sep-
zle. At lower pressure ratios an open wake ow establishes, aration was observed for contoured central plug bodies de-
with a pressure level practically equal to the ambient pressure signed with the method proposed in Ref. 29. Recent numerical
(Fig. 17, right column, top). At a certain pressure ratio close simulations of contoured plug nozzles performed at DLR
to the design pressure ratio of the full-length plug nozzle, the within the ESA Advanced Rocket Propulsion Technology
base ow suddenly changes its character and turns over to the (ARPT) Program on advanced rocket propulsion also show
closed form, characterized by a constant base pressure that is that no ow separation occurs from well-contoured, full-length
no longer in uenced by the ambient pressure. Analyses indi- central plug bodies.26 Principal ow eld developments pre-
cate that shorter plug bodies with higher truncations trigger an dicted by these numerical simulations are again in a good
earlier change in wake ow at pressure ratios below the design agreement with experimental data published in Ref. 28. Within
pressure ratio. At the transition point the pressure within the the frame of this ESA ARPT Program performance and ow
wake approaches a value that is below ambient pressure, and behavior of clustered plug nozzles at different truncations are
the full base area induces a negative thrust (Fig. 17, right col- being examined by European industries [Société Européenne
umn, center). This thrust loss depends on the percentage of de Propulsion (SEP), Volvo, Dasa] and research institutes
truncation and the total size of the base area. Published ex- (ONERA, DLR) with subscale cold- ow plug models26 (Table
perimental data and numerical simulations reveal an increasing 3). Numerous experimental investigations on subscale models
thrust loss for shorter plug bodies, because the total base area of circular and clustered plug nozzles with cold- and hot-gas
increases. ow were performed at the Keldysh Center with and without
Beyond the transition point, the pressure within the closed external ows. Thrust characteristics, pressure and heat- ux
wake remains constant. At these lower ambient pressures, the distributions along the plug, and acoustic characteristics of the
base pressure is then higher than the ambient pressure, result- circular jet were investigated.
ing in a positive thrust contribution of the total base area. Per-
formance data of a numerically simulated truncated plug noz-
zle are shown in Fig. 19, and compared to the same plug 2. Linear Plug Nozzles
nozzle with full-length central body and a conventional bell Performance behavior and ow eld development for linear
nozzle. Design parameters for this truncated plug nozzle are plug nozzles as a function of ambient pressures are in principle
the same as for the full length plug (see Fig. 18 and Table 2). similar to those of circular plug nozzles (Fig. 17). However,
Figure 20 shows a typical photograph of a sea-level hot-run special attention must be paid to the in uence of both end
test with a truncated, toroidal subscale plug nozzle, performed sides, where the surrounding ow disturbs the expanding ow-
at DLR. 5 Nozzle design data are included in Table 2. For com- eld, resulting in an expansion of the ow normal to the main
parison of experimental results with numerical computations, ow direction and, therefore, in an effective performance loss.
HAGEMANN ET AL. 629
a
Table 3 Design parameters for linear and toroidal subscale plug nozzles
systems of recompression and expansion waves (Fig. 27). At The fuels are burned in two different combustion chambers,
higher altitude, the lower ambient pressure allows more gas with one located completely inside the other in the case of
expansion within the nozzle, resulting in a higher effective engines with dual-throat nozzles, or with a conventional bell
expansion area ratio. But in contrast to plug nozzles, the pres- thrust chamber surrounded by an annular thrust chamber in the
sure in the wake of the center plug is always less than the case of dual-expander engines. This type of engine has a built-
ambient pressure because of the aspiration effect. This occurs in acceleration – reduction capability, achieved by shutting
at low pressure ratios when the wake is opened and results in down one of two thrust chambers. The total engine thrust is
an aspiration loss. Furthermore, because the exhaust ow ex- then provided by the remaining thrust chamber with the use
pands to this base pressure rather than to the ambient pressure of the total nozzle exit area leading to an increase in speci c
level, wall pressures downstream are overexpanded. This re- impulse. Apart from the indicated bene ts of dual-mode en-
sults in an additional overexpansion loss. As the pressure ratio gines, which will be discussed in more detail later, their de-
increases, the wake region closes and is thus totally isolated
from the ambient environment (see Fig. 27). The behavior dur-
ing transition from open wake to closed wake is again equal
to plug nozzles, and the base pressure in the closed-wake re-
gion is essentially independent of the ambient pressure.
The E – D nozzle concept has also been a subject of numer-
ous analytical and experimental studies. Results from these
studies show that E – D nozzle capabilities for altitude com-
pensation are poor, and are in fact worse than those of plug
nozzles, because of aspiration and overexpansion losses.35 For
high-area-ratio nozzles with a relatively short length, an E – D
nozzle performs better than a comparable conventional bell
nozzle at the same length because of lower divergence and
pro le losses than the bell nozzle.
The advantages of the E – D nozzle concept include its small
engine envelope and no moving parts. However, like toroidal
plug nozzles, E – D nozzles have the disadvantage of having Fig. 28 Sketch of a dual-throat nozzle, view of combustion cham-
ber and throat region.
higher throat heat uxes relative to a conventional bell nozzle
with an equal throat area. The higher throat heat ux results
again from the relatively thin annular throat gap. This problem,
however, can be remedied with the modular thrust cell cluster
concept. Additional advantages of E – D nozzles with clustered
thrust cells are the same as those already discussed for the
clustered plug nozzle concept, with regard to ease in manu-
facturing, thrust vector control by throttling or shutting off an
individual or group of thrust cells, and lower nozzle throat
heating.
E. Dual-Mode Nozzles
Dual-mode rocket engines using one or two fuels offer a
trajectory-adapted dual-mode operation during the ascent of a
launcher, which may be of signi cant advantage for single-
stage Earth-to-orbit vehicles as compared to conventional
rocket engines with bell-type nozzles. This engine concept in-
volves the use of a dense propellant combination with mod-
erate performance during liftoff to provide high thrust during Fig. 30 Performance losses in dual-throat nozzles as a function
the initial ight phase, and a better performing propellant com- of nozzle design parameters (with r 4/r 1 = 3.52, r 2/r 1 = 1.16, and mÇ 2
bination in vacuum, which result in higher speci c impulse. = 0.0).
632 HAGEMANN ET AL.
velopment and construction require considerable technological the burn. In principle, the two operation modes are comparable
effort. to those of dual-throat nozzles.
Numerical simulations of the ow elds in dual-expander
1. Dual-Throat Nozzles nozzles during all operation modes have been performed at
A dual-throat nozzle con guration is shown in Fig. 28. It Aerojet,38 DLR,39 and the Keldysh Center. Aerojet simulations
consists of two conventional bell-thrust chambers, with one are based on an engine design using different propellants,
located completely inside the other. At low altitude, the outer hydrocarbon/oxygen for the inner chamber and hydrogen/ox-
thrust chamber operates with the inner thrust chamber running ygen for the outer chamber, whereas DLR simulations are
in parallel. In this operation mode the engine has a larger throat based on a combined vehicle/engine analysis using single-fuel/
providing a moderate expansion area ratio. During the mission, single-mixture ratio dual-expander engines with hydrogen/oxy-
the outer thrust chamber is shut off and operation continues gen. Table 4 summarizes combustion chamber parameters. Flow
with only the inner engine. In this con guration, ow from phenomena observed in numerical simulations are shown in Fig.
the inner engine expands and attaches supersonically to the 32a for the mode 1 operation with both thrust chambers burn-
outer engine, resulting in a higher expansion area ratio for the ing. Compression waves are induced near the inner nozzle lip
remainder of the burn. Flow phenomena in both operation as a result of the strongly inhomogeneous ow character at the
modes are included in Fig. 29. point where both exhaust gases of inner and outer combustion
Hot- red tests at Aerojet were conducted to provide heat chamber merge into the common divergent nozzle part. Up
transfer data that were very useful for the thermal analysis and and down variations of pressure ratios at the lip did not change
design of the dual-throat nozzle con guration.37 These tests this wave formation, even in cases of signi cantly lower exit
showed that ow separation occurred in the inner engine noz- pressures in the inner nozzle compared to the pressure eld of
zle at higher ratios of outer to inner chamber pressures during the outer nozzle at the lip. Further downstream the compres-
the rst operation mode with both chambers burning in par- sion waves interact with the wall, resulting in a re ection of
allel. The ow separation resulted in a higher heat load to the the compression waves back into the ow eld.
inner nozzle. Subscale tests performed at the Keldysh Center 3 During mode 2 operation a strong expansion of the outer
have shown that the additional loss caused by the nozzle con- nozzle ow is observed, when the expansion ratio suddenly
tour discontinuity during vacuum operation with active inner increases at the end of the nozzle lip. The ow is directed
chamber is in the range of 0.8 – 4%, depending on geometrical toward the axis of symmetry. Near the centerline the ow then
data (see Fig. 30). This high-performance loss results from the turns over to the axial direction, inducing a recompression
interaction of the inner chamber jet with the outer chamber shock. The static pressure rises signi cantly in this recom-
nozzle wall.3 The decrease of the jet incidence angle on the pression region on the centerline. A sub- and supersonic recir-
wall by means of gas injection through the outer chamber re- culation zone establishes in the inner chamber of the dual-
duces this performance loss by 0.4 – 0.7%.3 expander nozzle. Figure 32b emphasizes the essential ow
pattern in this operation mode.
2. Dual-Expander Nozzles These analyses have shown that dual-expander nozzles pro-
A dual-expander nozzle has two concentric thrust chambers duce high performance in both operation modes.38,39 Figure 33
and nozzles. It consists of a conventional bell thrust chamber summarizes performance behavior as a function of ight alti-
surrounded by an annular thrust chamber. Both chambers have tude for the dual-expander engine simulated at DLR. For com-
short primary nozzles, which end in a common divergent noz- parison, performance data are included for two reference bell
zle extension. Figure 31 shows a typical dual-expander nozzle
con guration. At low altitude, both thrust chambers operate,
sharing the same exit area, which results in a moderate expan-
sion area ratio. Part way into the mission, one thrust chamber
is shut off, allowing the other nozzle to use the whole exit
area, creating a high-expansion-area ratio for the remainder of
Fig. 31 Sketch of a dual-expander nozzle, view of combustion Fig. 32 Flow phenomena in a dual-expander nozzle during a)
chamber and throat region. sea-level- and b) high-altitude operation.
a
Table 4 Thrust chamber parameters for simulated dual-expander nozzles
Aerojet DLR
Operation mode First (in /out) Second (in /out) First (in /out) Second (in/out)
Propellants C3H 8– O 2/H 2– O 2 — /H 2– O 2 H 2– O 2/H 2– O 2 — /H 2– O 2
Chamber pressure pc 414/207 bar — /207 bar 200/200 bar — /200 bar
Mixture ratio r̄ 3.3/7 — /7 7/7 — /7
Exit area ratio « 69 146 58 116
a
Inner/outer chambers.
HAGEMANN ET AL. 633
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