FR2615903A1 - Aerobic heat engine, particularly for the propulsion of hypersonic aircraft - Google Patents

Abstract

Heat engine essentially characterised in that the turbine T precedes the compressor C and is separated from the latter by a cooling heat exchanger Ech supplied with cryogenic fluid. This engine thus operates according to a thermodynamic cycle which, with respect to conventional heat engines, is reversed. It is advantageous to take liquid hydrogen as the cryogenic fluid. The hydrogen leaving the exchanger Ech may then supply a combustion centre F1 situated upstream of the turbine, and a combustion centre F2 situated downstream of the compressor C. Application to hypersonic aircraft.

Description

AEROBIC THERMAL ENGINE. ESPECIALLY FOR PROPULSION
HYPERSONIOUES AIRCRAFT
The present invention relates to an aerobic heat engine, in particular for the propulsion of hypersonic aircraft, or even of orbital aircraft in their aerobic flight phase, namely capable of flying at Mach 6 or more.

 it will be in particular an engine of the type using at least two heat sources at different temperatures, namely at least one hot or intermediate temperature source and one cold source, and comprising at least one turbine driving a compressor, the kinetic energy of the ejection gases producing the thrust necessary for propulsion.

Different types of engines are currently envisaged in the field of applications defined above 1-) The turbojet engines with heating, in which,
conventionally, the compressor precedes the turbine
and is supplied by an atmospheric air inlet,
sources of heat (fireplaces powered for example
hydrogen) being provided on the one hand between the
compressor and turbine, on the other hand, for the
heats, between the turbine and the outlet nozzle 2) the turbojet engines to be heated with exchanger in
upstream, i.e. of the same type as under 1-), but with
a cooling heat exchanger, for example
air / cryogenic hydrogen type, located between
the air inlet and the compressor 3) the turbofuses, in which the compressor is
powered by the air inlet and the turbine by a
gas generator itself powered by an oxidant
and an on-board reducer, the gases coming from the
compressor and turbine being mixed and
burned in a hearth, before ejection by the outlet nozzle; 4) Combined motors, which consist of a
assembly of motors or parts of motors,
way to ensure good performance both in
specific impulse that in pushing, on all
flight domain. These motors can thus include
for example, a turbojet, a ramjet
and a rocket, usually with a blower or
compressor placed upstream.

The disadvantages of these different types of motors are as follows in configurations 1, 3 and 4 the compressors or blowers (in configuration 4) are subjected, from Mach numbers of the order of 3 to 4, to temperatures very important stops, which no longer guarantee the good mechanical strength of these organs. In addition, the peripheral speed of the wheels, relative to the square root of the stopping temperature, being greatly reduced, the pressure ratio of the compressors is considerably reduced, which affects performance.

 The reheated turbojet engines with upstream exchanger resist a little better at the high shutdown temperatures at the inlet, due to the presence at the head of the cooler exchanger, but we still reach quickly, in the last stages of the compressor, a temperature equal to or greater than 900 C, which again poses a problem of the materials' behavior. In addition, at high Mach numbers the exchanger, placed at a hot point of the thermodynamic cycle, is no longer interesting from a thermopropulsive point of view.

 As for the turbofan engines, they make it appear the difficulty of efficiently mixing at high speed the central fluid often rich in hydrogen, and the peripheral fluid coming from the compressor.

 Finally, the combined engines are characterized by their technological complexity, their difficulty of development and the persistence of thermal problems to which the turbojets are subjected.

The object of the invention is to remedy the drawbacks of the prior art which has just been mentioned, and, in general, to develop a propellant architecture suitable for hypersonic aerobic missions which has the following advantageous characteristics - simplicity and reliability - allowing efficient takeoff, at least in
conditions comparable to those of a turbojet
modern to warm - able to operate up to Mach numbers
vés, -of the order of 6 and more, and resist principa
Also at high shutdown temperatures at the entrance to the
engine.

For this purpose, an aerobic heat engine of the type
mentioned at the beginning is, in accordance with this
invention, essentially characterized, in its
simplest version, in that said hot spring
or at intermediate temperature is constituted by
atmospheric lrair and the cold source by a fluid
cryogenic, and in that, if we consider said mo
in the direction from the entry of the gases to their
outlet, said turbine is arranged upstream of said
compressor and is powered by the gases from said
hot spring or at intermediate temperature, a
chiller heat exchanger, supplied by said
cryogenic fluid and constituting said source
cold, being interposed between the turbine and the com
presser.

 Thus, it can be seen that, unlike the conventional arrangement in which the hot source is ensured by the combustion of a combustible mixture and the cold source is constituted by ambient air, according to the present invention the hot source is constituted by ambient air, and the cold source by an exchanger in which a cryogenic fluid circulates. It is therefore a "reverse cycle motor".

 The present invention therefore relates, firstly, to a reverse cycle heat engine which, in its simplest version, comprises a turbine directly fed by atmospheric air, from which it escapes at a very high pressure. lower in an exchanger traversed (advantageously against the current) by a cryogenic fluid, the compressor, driven by the turbine and arranged downstream, finally reducing the air pressure to the ambient value. The useful power conventionally comes from the difference between the power produced by the turbine and that consumed by the compressor.

 In such an engine according to the invention, the cryogenic fluid is not necessarily a fuel. It could, for example, be liquid nitrogen, or even liquid helium, which would in particular make it possible to obtain absolute operating safety in explosive atmospheres, or to avoid infrared detection, since there would then be no combustion, unlike the case of conventional heat engines.

 However, it will be particularly advantageous to use liquid hydrogen as cryogenic fluid, this fluid constituting a remarkable cold source, and which, in addition, in its gaseous form, after heating in the exchanger, can be used as fuel in at least one hearth which can itself supply turbines or propellant nozzles.

We will then obtain another type of new heat engine, having at least three main heat sources having different temperatures - a hot source from the combustion of hydrogen - an intermediate source constituted by the air in the
stopping conditions (the engine being assumed to be in displacement
cement); and - a cold source constituted by the hydro exchanger
liquid gene.

 One or more hot springs may be provided in the engine, namely that it may comprise, downstream of an air inlet, a first combustion chamber supplying said turbine and possibly, downstream of the compressor, a second combustion feeding an ejection nozzle.

 A heat engine according to the invention can therefore also be characterized in that it comprises, from upstream to downstream relative to the path of the gases - an air inlet - a first combustion chamber - a turbine - a cryogenic air / hydrogen heat exchanger - a compressor mechanically coupled to the turbine - a second combustion chamber; and - a gas ejection nozzle.

 It can be seen that this is again a reverse cycle engine, since the first combustion chamber and the turbine are located upstream of the compressor, having regard to the direction of flow of the gases in the engine.

 Such an engine will be particularly well suited for flights at Mach 6. Indeed, the turbine is then directly supplied with air at 1,300 or 1,500 C, a value used in conventional turbojets, (and it will be useless to cool the 'air intake). We note here an important advantage presented by an engine according to the invention over the conventional architecture with the compressor at the head. As it cannot withstand these temperatures, it is necessary to cool the air at the intake. This cooling must be very important, since the compressor must heat the air to a temperature which will remain much lower than that admitted at the inlet of the turbine.

 It therefore clearly appears that the adoption of the reverse cycle in an engine according to the invention will make it possible to achieve flight Machs greater than the limit Mach of the conventional turbojet engine with cooled intake.

 In any event, provision may also be made, in an inverse cycle engine of the type described above, for the air inlet, as well as said gas ejection nozzle, to be of variable geometry.

Embodiments of a heat engine according to the invention, as well as the corresponding operating diagrams and a certain number of alternative embodiments will now be described, by way of non-limiting examples, with reference to the figures in the drawing. annexed in which - Figure 1 shows schematically, in its
basic version, a reverse cycle heat engine
according to the present invention, its diagram of
theoretical operation H = f (S) being represented at
Figure 2 - Figure 3 is a diagram similar to that of
figure 2, but taking into account isentro yields
compressor and turbine spikes - Figure 4 shows the coefficient of performance curves
formance and specific power according to
temperature differences between inlet and outlet
of the turbine - FIG. 5 diagrammatically represents an engine
reverse cycle thermal system according to the invention,
type with two combustion chambers - Figures 6 and 7 show the diagrams of
operation H = f (S) of such a motor respectively
for a flight to Mach 0 and for a flight to large Mach
and - Figures 8 and 9 schematically represent at least
very reverse cycle motors according to the invention,
respectively a turbofuse and a motor with effect
pump produced at the compressor.

 The reverse cycle cryogenic motor of FIG. 1 comprises, from upstream to downstream, having regard to the direction of movement of the gases: a hot source constituted by an AR anti-radiator; a turbine T receiving air anti-radiator at temperature T1; a cooler heat exchanger Ech, taking in the expanded air in the turbine, and brought to temperature T2, a quantity of heat Qf bringing this air to temperature T3; and a compressor C coupled to the turbine T and bringing the air to the temperature T4, this air being brought back to the AR radiator. The exchanger is supplied countercurrently by a source Fc of cryogenic fluid, for example liquid nitrogen.

In the enthalpy / entropy diagram of FIG. 2, the operating points corresponding to locations 1, 2, 3 and 4 of the diagram in FIG. 1 have been indicated in 1, 2, 3 and 4, so that this diagram s practically explains itself - from 1 to 2, isentropic expansion of the air in the turbi
ne T (its efficiency as well as that of compressor C
are considered here = 1), from a temperature
T1, which supplies the turbine with energy Wt; - from 2 to 3, isobaric cooling in the exchanger,
up to temperature T3, with sampling of the
quantity of heat Qf - from 3 to 4, isentropic compression in the compressor
C, which takes from the turbine shaft T 1 'energy
gie Wc - from 4 to 1, isobaric heating of the tempera air
ture T4 at temperature T1 in the AR radiator,
the system thus withdrawing outside the quantity
heat to bring the main fluid to
the initial temperature.

The work available on the tree is Wt - Wc and the coefficient of performance Cp = Wt-Wc = T1
Qf T2
Such a motor can provide a very large energy (Wt - Wc), compared to the small amount of heat (Qf) supplied by the main fluid to the cold source.

 In FIG. 3, which takes account of the non-isentropic operations of the compressor and of the turbine, the same references and indications have been used as in FIG. 2 to designate the same points of the diagram, so that this can be explained by himself.

The curves representing the coefficient of performance CP and the specific power Wf as a function of
T1-T2 are given in FIG. 4. We can see that we can obtain - for a temperature T1 of 115 C at the high point of the cycle and a temperature T3 of - 173 C at the low point, with an aerodynamic efficiency rIA, fixed parts of 0.95, a specific power greater than 50 kJ / kg of the main fluid when T1-T2 is between 100 and 200 C.

 The coefficient of performance CP 'and the specific power W'f for T1 = 15 C have also been indicated in this figure. For these two examples, the efficiencies Pt of the turbine and Pc of the compressor have been assumed to be equal to 0.90. We have also represented (dashed curve) the coefficient of performance CP "corresponding to the cycle of FIG. 2 for T1 = 15. C.

 In FIG. 5, a diagrammatic description has yet been given of a thermal engine according to the invention of the general reverse cycle type but also comprising two combustion chambers consuming hydrogen and therefore being of the type with three heat sources main: the hot spring, the intermediate source and the cold source which were mentioned above.

 This engine therefore comprises, from upstream to downstream, an air intake EA, a first combustion chamber or hearth F1, a turbine T, a hydrogen heat exchanger designated Ech, a compressor C, a second combustion or hearth F2, and an outlet nozzle TS.

 In Fc, the source of cryogenic hydrogen is fed to the foci F1 and F2 after the hydrogen has been heated in the exchanger Ech to the temperature Tg, this hydrogen being stored at a temperature T8 of the order of -263 to -258'C.

Figures 6 and 7 show the enthalpy / entropy operating diagrams for such an engine assumed to be Mach of zero flight (Figure 6) and
Large flight mach (Figure 7). To facilitate understanding of these figures, the points of the diagram corresponding to the indices 0, 1 7 shown on the diagram of FIG. 5 have been identified by the indices i0, i 1 i7.

That said, the main structural and operating characteristics of such an engine could be as follows - Air intake EA [0-1]
EA air intake, which can be geometry
variable, must allow a good supply of
engine both at takeoff and at Mach number
high (6 and above).

We see that at takeoff (Figure 6) the points i0
and i1 of the cycle are confused, while in flight at Mach
high, we have a corresponding O-i1 compression phase
due to a significant increase in enthalpy due to
dynamic compression in the air intake EA
(figure 7). It should also be noted that on the
figure 6 the atmospheric pressure on the ground has been
represented by the isobar po, and that in figure 7
the pressure at flight altitude z has been represented
by the designated isobar pz - Foyer F1 [1-2]
The F1 fireplace remains in principle on as long as the
engine shutdown temperature remains below
advantageously at a temperature of the order of
800 to 1100 ° C, point i2 of Figures 6 and 7).

turbine T [2-3)
The expansion of the gases in the turbine T makes it possible to
supply the power on the compressor C shaft
located downstream. The presence of the turbine at the head
has a huge advantage from the point of view of
behavior of materials in relation to the turbojet
classical and in particular to the large Mach numbers (from
the order of 6) where the stop temperature Ti can be
siner 1530'C:
a) the turbine blades are very thick, con
compared to compressors, and lend themselves
easily to internal cooling using
cryogenic hydrogen, while the blades of
compressor, much thinner, very suitable
wrong.

b) in the future, advances in technology (blades
cooled, ceramic blades at least for the
first wheel) will exceed the temperature
1,600 C currently reached.

c) in a turbine, the shutdown temperature goes
falling from upstream to downstream when it is the
otherwise in compressors.

d) at takeoff, a temperature T2 = 730 to 830 C is
sufficient without significant alteration of the impulse
specific or thrust, which increases
lie about the longevity of the turbine.

e) the first wheel of a turbine is attacked with a
significant kinetic moment, which is not the case
for the first wheel of a compressor. He re
note that the stop temperature at the leading edge
blades of a turbine will be lower by about
200 to 300 C compared to that obtained on the
leading edges of the first wheel of a compres
sister.

 The turbine T can in a certain way be considered as the heat shield of the engine, preferably at the exchanger which cannot be cooled, provided that the first distributor is effectively cooled.

 The turbine distributor is well placed since, at low pressure, heat exchanges are attenuated.

 The use for the first impeller of blades rather with action than reaction (increase of the ratio tangential absolute speed of entry / speed of rotation) makes it possible to reduce the stopping temperature appreciably on the movable blade, this particularly in the case of clean air, which will no longer be the case behind combustion products with a high hydrogen content where Cp 8000 no longer makes it possible to significantly lower the relative shutdown temperatures (case of the turbofuse).

- Exchanger Ech [3-4]
A burnt gas / cryogenic hydrogen exchanger
that against the current to minimize production
entropy, allows to cool very significantly the
gases from the T turbine. This cooling takes place
kills substantially at constant pressure (at losses of
charge close).

The very high heat capacity of hydrogen
(cl ~ 14 to 15000 J / kg-C) allows effective cooling
the main circuit gases with a flow
of hydrogen ùc the order of magnitude of the one that goes
be burned in the set of the two rooms F1 and F2,
flow in principle less than 7% of the air flow,
value which corresponds substantially to the equality of
temperature variations in the hydrogen circuit
and in the main circuit T3-T4 = -Tg-T8. The
operating at wealth> 1 allows no vehicle
ler than reducing gases, which is favorable as
the mechanical strength of the materials.

- Compressor C [45]
Isobaric curves, in the diagram (H, S)
Figures 6 and 7, can be assimilated, in pre
better approximation, to deductive exponentials
from each other by translation along the axis of S.

It follows that the compression of the gases, after a
intense cooling, using power
developed by the turbine, will increase.
gas pressure significantly compared to that
existing at the entrance of the focus F1-: Pi5> Pi1 which, well
sure, promotes thrust.

- F2 fireplace [56)
The F2 type heating element will allow
burn at least some of the heated hydrogen
by the exchanger with the remaining oxygen available
after combustion in F1, this in order to reach the
maximum temperature Ti6 of the cycle.

- TS outlet nozzle [6]
Kinetic energy of gases in the nozzle
ejection TS, measured by the difference in enthalpy
<Hi6-H7), assuming the proper nozzle, will be fine
sure function of the maximum pressure pi6 of the cycle and
of the maximum temperature Ti6 of the cycle.

 As a variant of the engine which has just been described with reference to FIG. 5, it should be noted that, thanks to a bypass pipe, it would be possible, with low Mach numbers, to invert the framed sets of a broken line, namely the hearth F1-turbine T assembly and the Ech-compressor exchanger assembly C. This would make it possible to very significantly increase the specific impulse at takeoff, without altering the thrust.

 One could also consider, to extend the use of the reverse cycle engine up to Mach 7, to decompose the compressor C into two bodies, by interposing between them a main gas / hydrogen countercurrent exchanger.

 This would make it possible for the shutdown temperature at the level of the last stage of the compressor to remain compatible with satisfactory behavior of the materials.

 In Figure 8 there is shown schematically, as a variant, another thermal engine of the reverse cycle type, namely a turbofuse.

 In this figure we used the same references as in Figure 5 to designate the same parts of the engine. The second combustion chamber F2 is associated here with a mixing chamber M, supplied with pressurized oxygen from a cryogenic oxygen tank R. This makes it possible to burn a greater part of the hydrogen used and to increase the thrust .

It is thus possible to obtain an optimal specific impulse torque, particularly useful for a rapid climb takeoff; in general, this will optimize this couple depending on the mission.

FIG. 9 shows yet another variant of the reverse cycle thermal engine shown in FIG. 5. Here again, the same references have been used in these two figures to designate the same parts of the engine. As in the variant of FIG. 8, the second combustion chamber F2 is associated with a mixing chamber M, but its supply of oxidant is ensured here by a proboscis effect (symbolized in
ET), sucking in atmospheric air and thus making it possible to supply the combustion chamber F2 with the additional oxygen necessary for the combustion of all of the hydrogen.

 We can ensure, for example, that this trumpet effect arises at the outlet of compressor C.

The advantages of a reverse cycle engine according to the invention are as follows the reversal of the relative position of the turbine
and compressor, compared to turbojets
conventional, allows you to ensure that the pressure
maximum of the thermodynamic cycle is performed at
the inlet of the outlet nozzle, which is not the
advantage of these conventional turbojet engines.
particularly interesting because, at equal temperature
scratch and pressure in the F2 hearth, the thicknesses
crankcase will be more important in the turboreac
classic tor, especially at the hearth - the turbine placed at the head, just behind the entrance
air, behaves in large numbers of Mach like a
real heat shield for the rest of the engine - the thermodynamic cycle arising from the layout
particular of the turbine upstream of the compressor
maximizes the benefits of the source
cryogenic cold, removing calories from the system =
me at the low point of the thermodyna cycle temperatures
mique, especially at large Mach numbers - an engine according to the invention also allows a
efficient take-off, without accompanying trolley, and
this by sparing the turbine and the stoves, where a
increased longevity of this engine; we can consider
operate with low hydrogen consumption at
from Mach 2.5 approximately. In other words, and at
by way of comparison, the engine according to the invention
allows takeoff to obtain the same thrust as a
turbojet engine heats up, and works usefully
beyond the operating limit of the statoreac
subsonic combustion torches - there is still a significant thrust at Mach of
on the order of 6 to 7, while the performance of a sta
subsonic combustion toreactor, to such numbers
of Mach, weaken a reverse cycle motor conforming to this
invention will be particularly well suited to the
propulsion of orbital aircraft, as well as aircraft
hypersonic transport.

Claims (14)

 1. Aerobic heat engine, of the type using at least two heat sources at different temperatures, namely at least one hot source or at intermediate temperature and one cold source, and comprising at least one turbine (T) driving a compressor ( C), the kinetic energy of the ejection gases producing the thrust necessary for propulsion, characterized in that said hot source or at intermediate temperature is constituted by atmospheric air and the cold source by a cryogenic fluid, and in that that, if we consider said engine in the direction going from the entry of gases to their exit, said turbine (T) is arranged upstream of said compressor <C) - and is supplied by gases from said hot source or at intermediate temperature , a cooler heat exchanger (uh), supplied by said cryogenic fluid and constituting said cold source, being interposed between the turbine (T) and the compressor (C).  2. Heat engine according to claim 1, characterized in that said cryogenic fluid consists, in whole or in part, of liquid nitrogen.  3. A heat engine according to claim 1, characterized in that said cryogenic fluid consists, in whole or in part, of liquid hydrogen.  4. A heat engine according to claim 3, characterized in that it further comprises at least one combustion chamber tF1, F2) supplied with gaseous hydrogen by said cooler exchanger (uh).  5. A heat engine according to claim 4, characterized in that it comprises, downstream of an air inlet (EA), a first combustion chamber (F1) supplying said turbine (T).  6. Heat engine according to claim 4 or 5, characterized in that it comprises, downstream of the compressor (C), a second combustion chamber (F2)  feeding an ejection nozzle (TS).  7. Aerobic heat engine, of the type using at least two heat sources at different temperatures, namely at least one hot source or at intermediate temperature and one cold source, and comprising at least one turbine (T) driving a compressor ( C), the kinetic energy of the ejection gases producing the thrust necessary for propulsion, this engine comprising as an intermediate temperature source atmospheric air and as a cold source a cooling heat exchanger (Ech) supplied with liquid hydrogen, characterized in that it comprises, from upstream to downstream with respect to the gas path - an air inlet (EA) - a first combustion chamber (F1) - a turbine (T) - a heat exchanger air / cryogenic hydrogen  (Ech) - a compressor (C) mechanically coupled to the turbine  (T) - a second combustion chamber (F2) this and - a nozzle (TS) for ejecting the gases.  8 Heat engine according to claim 7, characterized in that said air inlet (EA) is of variable geometry.  9. Heat engine according to claim 7 or 8, characterized in that said gas ejection nozzle (TS) is of variable geometry.  10. A heat engine according to any one of claims 7 to 9, characterized in that it comprises a bypass pipe making it possible to invert the hearth assembly (F1) -turbine (T) and the exchanger assembly (Ech) -compressor (C).  11. selc-n heat engine any one of claims 7 to 9, characterized in that said compressor (C) comprises two bodies separated by a cooler heat exchanger.  12. A heat engine according to any one of claims 7 to 9, characterized in that said second combustion chamber (F2) is associated with a mixing chamber (M), which is supplied with oxygen pressurized by an oxygen tank cryogenic (R).  13. A heat engine according to any one of claims 7 to 9, characterized in that said second combustion chamber (F2) is associated with a mixing chamber (M), which is supplied with atmospheric air by a suction effect ( AND).  14. Hypersonic aircraft, characterized in that it comprises a heat engine in accordance with claim 1 or any one of claims 3 to 13.