supersonic combustion of kerosene/h2-mixtures in a model scramjet

SUPERSONIC COMBUSTION
OF KEROSENE/H 2 -MIXTURES IN A MODEL
SCRAMJET COMBUSTOR
C. GRUENIG * and F. MAYINGER
Institute A for Thermodynamics, Technical University Munich,
D-85747 Garching / Germany
Liquid hydrocarbon supersonic combustion has been experimentally investigated. Kerosene was
burnt in a steady, vitiated Mach 2.15 - air flow of a model scramjet combustor. The fuel is injected
into the supersonic air stream by means of pylons. The effervescent atomisation method has been
employed such that the liquid fuel is injected as a spray. By means of the Mie scattering technique
the fuel jet structure was visualised and the evaporation rate estimated.
The mechanisms of ignition and combustion of the kerosene/H 2 -mixture were studied and
compared with the case of hydrogen combustion. Combustor ignition limits have been determined.
Fuel-specific combustion phenomena are discussed. It was found that for the kerosene combustion a
gas dynamic feedback mechanism strongly affects the supersonic combustion process.
Keywords: Supersonic combustion; kerosene; scramjet; self-ignition; gas dynamic feedback
1 INTRODUCTION
Supersonic combustion ramjets are an attractive option to propel future hypersonic flight systems
(e.g. Hunt and Rausch, 1998; Scuderi et al., 1998; Maita and Kubota, 1998; Grallert and Herrmann,
1998). Most research activities focus on hydrogen-fuelled scramjets as hydrogen is the fuel of
choice for high speed flight because of its properties such as its high heat of combustion, its high
molecular diffusion velocity, its ability turn burn at lowest pressures, and – what is important for
the application in hypersonic flight vehicles – its high specific heat, i.e., its cooling capacity. In
fact, the cooling capacity is one of the most important porperties that strongly infuences the fuel
selection for high speed flight systems.
For flight Mach numbers below appr. 8 liquid hydrocarbon fuels possess sufficient cooling
capacity (Hunt and Rausch, 1998; Hunt et al., 1997). Hence, below M 0 = 8 the advantages of liquid
hydrocarbon fuels (high volumetric specific impulse ρ fuel ∗I sp which is of special importance for
narrow body and/or small scale flight systems) can be utilised. Since the early nineties there has
been an increase in research in the field of hydrocarbon supersonic combustion (e.g. Avrashkov et
al., 1990; Avrashkov et al., 1992; Vinogradov et al., 1992; Romankov and Starostin, 1993; Owens
et al., 1997; Owens et al., 1998, Bouchez et al., 1998; Semenov et al., 1998).
* Corresponding Author: Email: carolus.gruenig@avl.com

In this paper we describe research activities that in particular supplement the work by Avrashkov
et al. (1990 and 1992) at the Moscow Aviation Institute. The basis of the present paper has been
laid by a joint co-operation with this institute (Gruenig et al., 1996). The aim is to get insight in the
processes of the kerosene/air mixing, the fuel evaporation, and the ignition and supersonic
combustion. Furthermore, fuel-specific combustion aspects are to be explored.
2 EXPERIMENTAL DETAILS
The experiments have been performed in the supersonic flow channel of the Institute A for
Thermodynamics at the Technical University Munich. This is a blow-down facility supplied with
reservoir conditions p 0 = 13.5 bar, T 0 = 293 K and m& max = 1.5 kg/s. Figure 1 shows a schematic of
the flow channel.
FIGURE 1
Schematic of supersonic combustion test facility
In the preheater the air total temperature of is raised by precombustion of hydrogen. To
compensate for the oxygen consumption in the preheater, additional oxygen is supplied to the air
flow approximately 100 hydraulic diameter upstream of the preheater. The preheated air passes
through a Laval-nozzle wherein it is accelerated to M = 2.15. Typical flow conditions at the
entrance of the supersonic test combustor are summarised in Table I. These conditions
approximately correspond to flight Mach number M 0 = 6.5 at altitude h = 23.5 km on a
representative trajectory of a transatmospheric accelerator, assuming a typical inlet compression
ratio (Billig, 1996).
M
[-]
T 0
[K]
T
[K]
m&
[kg/s
]
p
[bar]
u
[m/s]
2.15 1700 1000 0.33 1.0 1310 3.3∗10 5 0.705 0.188 0.107
Re
[-]
w N2
[-]
w O2
[-]
w H2O
[-]
TABLE 1
Flow conditions at entrance of test combustor

Due to the precombustion there is a significant amount of water in the test air. By means of
spectral analysis the test air was checked for residual radical concentrations. For T 0,air < 1750 K no
radicals could be detected. However, minor concentrations can not be totally excluded.
Nevertheless, even taking into account possible radical contaminations the measurements will yield
valuable results as Mitani (1995) showed that radical contaminations have the same effects on the
scramjet behaviour as a higher temperature of radical-free air.
The supersonic test combustor (see Figure 1) is of rectangular cross section (width = 25 mm,
height = 27.5 mm at the entrance). The overall length is 645 mm. The bottom combustor wall
consists of modules of varying expansion angles. Modules with expansions angles of 0°, 2°, 4°, 6°
and 12° are available so that different combustor area ratios A exit /A entrance can be provided. The fuel
(kerosene Jet A-1) is injected 75 mm downstream of the combustor entrance by means of a pylon.
Further downstream provision is made to temporarily inject N 2 from both the top and bottom
combustor wall to support the combustor ignition if necessary.
Along the top combustor wall the static pressure p w is measured at seven locations. The static
pressure in the flow is mainly influenced by the combustion heat release and the combustor
geometry (Shapiro, 1953; Zierep, 1975). The heat removal by radiative and convective heat transfer
to the combustor walls is much smaller than the heat input by combustion (Gruenig, 1998).
Moreover, pressure measurements in case of inert gas injection into the preheated air stream
showed only a slight pressure level increase. Thus, the axial static pressure distribution indicates
areas and intensity of the heat release.
FIGURE 2
Optical setup for Mie scattering technique
The combustion chamber is equipped with quartz windows to give optical access to the flow
phenomena. To visualise the liquid fuel jet, the Mie scattering technique has been employed. For
theory and background of this measurement technique the reader is referred to the numerous
literature (e.g. Mayinger, 1994; Kelvin, 1970; van de Hulst, 1981; Gousbet et al., 1991). The optical
setup is shown in Figure 2. A Nd-YAG pulsed laser beam is formed into a light sheet (δ lightsheet
approximately 0.3 mm) and passed through the combustor. The two-dimensional scattering signal is

detected by a 14 bit intensified CCD camera. By this, spatially and time-resolved (t pulse = 7 ns)
images of the liquid fuel distribution are achieved.
To determine location and intensity of the reaction zone the CH selffluorescence is detected. The
fluorescence signal of the A 2 Σ−X 2 Π system with its band head at λ = 431.5 nm is spectrally filtered
by 4 dielectrically coated mirrors with FWHM = 20 nm and T > 92 % at λ = 430 nm and detected
with the ICCD camera. The exposure time is adjusted to t exposure = 100 ms.
3 RESULTS AND DISCUSSION
3.1 Formation of the Kerosene/H 2 -Mixture
To enhance mixing, droplet atomisation, and fuel evaporation the effervescent atomisation method
(Lefebvre, 1989) was employed. The principle of this spray formation technique is sketched in
Figure 3.
FIGURE 3
Principle of effervescent fuel dispersion technique
The liquid is saturated with gas bubbles. In the injector orifice the liquid is squeezed by the gas
bubbles. This is important for the atomisation because it is well known that the drop sizes produced
in a spray are roughly proportional to the square root of the initial thickness of the ligaments from
which they are formed (Rizkalla and Lefebvre, 1975). When the bubbles emerge from the orifice
they „explode“, shattering the surrounding liquid shreds and ligaments, thus, producing fine
droplets.
This atomisation technique was first applied to supersonic kerosene combustion by Avrashkov et
al. (1990) and further studied by Sabelnikov et al. (1998/1 and 1998/2). In Sabelnikov et al.
(1998/1) it was shown that the type of gas affects the supersonic mixing and combustion process.
They suggest to work with hydrogen rather than air or nitrogen. In our work we choose hydrogen as
the bubbling gas. The mixer, which was developed for the formation of the two-phase mixture, is
also shown in Figure 3. The aim is to produce a homogeneous bubble flow with a gas content
γ approximately 0.1 mass%. In the mixer the kerosene passes a perforated sheet of metal (δ sheet =
0.25 mm, ∅ orifice = 50 µm), which the hydrogen is supplied through.
The effect of the gas addition is demonstrated in Figure 4 where a jet of pure kerosene as well as
a kerosene/H 2 -mixture is vertically injected into room temperature atmosphere.

FIGURE 4 Demonstration of fuel dispersion method: vertical injection of kerosene and kerosene/H 2 jet
(p inject = 10.0 bar, m& kerosene = 5.0 g/s, m& kerosene/H2 = 3.2 g/s)
The injection pressure is rather low, p inject = 10.0 bar. The jet of pure kerosene remains compact
for about 80 mm and disintegrates only due to instability. Large drops are produced. The
kerosene/H 2 -jet forms a spray right after the injection. A large number of small fuel droplets is the
result.
3.2 Fuel/Air Mixing and Evaporation
To study the fuel mixing and evaporation in the Mach 2.15 air flow spatially and time-resolved
images of the liquid fuel Mie scattering signal have been taken. As an example Figure 5 shows the
axial distribution of a non-reacting, liquid fuel jet along the first 150 mm distance from the injection
pylon.

FIGURE 5 Structure of a non-reacting kerosene/H 2 mixing jet along the first 150 mm mixing length, fuel
normally injected into a Mach 2.15 air stream, Mie scattering single pulse images (t pulse = 7 ns, M air = 2.15,
p air = 0.7 bar, T 0, air = 560 K, T air = 300 K, m& air = 0.3 kg/s, m& kerosene/H2 = 3.2 g/s, γ = 0.1 %, p inject = 10.0 bar,
A exit /A entrance = 1.13)
A kerosene mass flow rate m& = 3.2 g/s containing 0.1 mass-% H 2 is normally injected into the
Mach 2.15 air stream. The air stream is slightly preheated (T 0, air = 560 K) to ensure that the ambient
static temperature of the fuel jet (T air = 300 K at the combustor entrance) is above the freezing point
of the kerosene Jet A-1 (T freeze, Jet A-1 = 222 K), but still below the boiling point of the kerosene
(T boiling, Jet A-1 = 440 .. 540 K).
From the mixing jet visualisation it can be seen:
• The fuel/air mixing jet occupies only part of the flow channel. Because of the relatively low
injection pressure (p inject = 10.0 bar), and thus the low momentum ratio, the trajectory of the fuel
jet remains at the pylon top level. The fuel spreads approximately 10 mm around the jet
trajectory.
• The liquid fuel is clearly separated from the surrounding air, which is entrained into the mixing
jet by large-scale coherent structures. At the given air stream static temperature these coherent
structures exist throughout the imaged combustor region.
• Widespread regions of diffuse Mie scattering signal indicate that a large amount of the liquid
fuel exists as small droplets. Individual drops can not be resolved with the employed detection
optics, which has a resolution of approximately 5 µm. Hence, the drop diameter are of the same
order of magnitude.
Apart from just visualising the mixing jet structure the Mie scattering technique can also be used
to discriminate between liquid and vapour phase of the fuel because the Rayleigh scattering signal
of the kerosene vapour is orders of magnitude weaker than the Mie scattering signal. Also the
strong dependence of the scattering signal on droplet size, polarisation plane and detection angle
vanishes in the case of irregular droplet shapes and a wide range of droplet size (Self, 1976) and,
hence, the Mie scattering signal can be assumed to be proportional to size and number of the liquid
fuel droplets. Therefore, the measured Mie scattering signal can be related to the amount of fuel
which still exists in the liquid phase at the measurement plane. By this, the fuel evaporation can be
studied.

Figure 6, showing instantaneous liquid fuel jet cross-sections at three axial positions for different
air flow temperatures, provides a visual impression of the kerosene evaporation (Note that the given
temperature values do not include the temperature rise by the oblique shocks due to the flow
disturbance by pylon and fuel jet. The fuel will be exposed to higher static temperatures.).
FIGURE 6 Liquid fuel evaporation: single-pulse Mie scattering measurements of jet cross-sections
along the combustor for different air flow temperatures (M air = 2.15, m& air = 0.3 kg/s, m& kerosene/H2 = 3.2 g/s, γ
= 0.1 %, p inject = 10.0 bar, A exit /A entrance = 1.13)
As before in Figure 5, the time-resolved images of the liquid fuel jet cross-sections show that the
mixing process is highly turbulent. Moreover, it can be seen that for T air = 290 K the amount of
liquid phase kerosene remains essentially unchanged. For increasing air flow temperatures first the
small fuel droplets evaporate. As expected, to evaporate the large-scale coherent structures it needs
a longer time and/or higher ambient temperatures.
To get a statistically more accurate estimation of the evaporation rate it is necessary to average a
number of measurements. Figure 7 shows the amount of liquid phase fuel crossing the measurement
plane in dependence of the air flow temperature where 10 Mie signal measurements have been
averaged.

FIGURE 7 Mass flow rate of the liquid phase fuel across the measurement plane in dependence of the air
stream temperature at the combustor entrance, data derived from Mie scattering measurements averaged over
10 laser shots
For T air = 290 K only a small part of the fuel evaporates. Assuming an oblique shock angle of
approximately 40° − earlier Schlieren measurements with similar fuel injectors revealed a shock
strength of that order of magnitude (Gruenig et al., 1996; Gruenig, 1998) − the fuel jet bow shock
and its reflection will raise the static temperature to T air about 450 K, which slightly touches the
kerosene boiling point. Increasing the air stream temperature leads to a rapid fuel evaporation. For
T air = 470 K about 50 % of the injected fuel evaporates within the first 150 mm of the mixing path.
For T air = 570 K nearly all fuel is evaporated within that distance. The fuel evaporation time scale
τ evaporate is of the order of 10 -4 s. This is about an order of magnitude below the fuel residence time
within the combustor (compare with Table 1) and, hence, a prerequisite for the supersonic
combustion of the kerosene is fulfilled.
3.3 Self-Ignition and Supersonic Kerosene Combustion
To self-ignite and burn the injected kerosene in the given combustor (A exit /A entrance = 1.13), the air
static temperature in the combustor entrance plane must be raised to T air appr. 1050 K. Again,
assuming the shock strength as mentioned above, the fuel jet will be exposed to an ambient
temperature of T air appr. 1400 K.
To visualise the supersonic reaction zone, first a spectrum of the flame emission has been taken
to select an appropriate spectral range. The measured spectrum is given in Figure 8.

FIGURE 8 Flame emission spectrum of the supersonic combustion of the kerosene/H2-mixture
First, it will be noted that there is no broadband continuum emission, i.e., thermal soot
incandescence. The absence of soot incandescence, which is typical for hydrocarbon diffusion
flames, indicates that the fuel is reacting under a premixed regime. This, in turn, indicates a very
intense fuel/air mixing. A homogeneous premixture is established before the self-ignition takes
place.
The flame emission mainly consists of band spectra from different reaction intermediaries: OH-,
CH- and C 2 -radicals. The most prominent peak at λ = 431.5 nm belongs to the A 2 Σ−X 2 Π system of
the CH-radical. This band has been chosen to locate the supersonic reaction zones and to estimate
the reaction rate. By spectrally filtering, as described in Chapter 2, the supersonic combustion
process has been imaged. In Figure 9 the supersonic reaction zone is shown.

FIGURE 9 Supersonic reaction of the kerosene/H2-mixture, CH-selffluorescence (Mair = 2.15, T 0, air =
1740 K, T air = 1030 K, air = 0.3 kg/s, kerosene/H2 = 3.2 g/s, ÀÃ Ã Ã ÈÃ ÀÃ Ã Ã S inject = 10.0 bar,
Aexit/Aentrance = 1.13)
The majority of the fuel self-ignites approximately 250 mm downstream the fuel injection. The
main reaction zone ends about 150 mm before the combustor exit, indicating high global
combustion efficiency. In the top left-hand part of Figure 9 the flow region of the first 150 mm
mixing length has been enlarged. In this region relatively low CH-selffluorescence signal levels are
detected. To visualise these low CH concentrations, the grey scale has been adjusted from the
minimum and to the maximum signal level in that region. It can be seen that comparatively weak
reactions take place well before the main part of the fuel starts to undergo reaction. There are small
quantities of CH radicals even in the wake of the pylon. These weak reactions take place in local
areas where the flow conditions (fuel/air mixture ratio and temperature) suffice for self-ignition.
The same behaviour has been discovered for the case of supersonic hydrogen combustion as well
(Gruenig et al., 1996). It can also be discerned that the weak reactions are induced by oblique
shocks due to the flow deflection at the second combustor segment. As the amount of fuel that
undergoes reaction is small, there is no much heat released into the flow. But what is more
significant for the global reaction, small amounts of free radicals − such as the observed CHradicals
− are produced, which are known to accelerate the self-ignition and subsequent reactions
(Suttrop, 1972; Suttrop, 1974).
In the top right-hand part of Figure 9 the CH selffluorescence emitted from the free jet leaving
the supersonic combustor is shown. Again, the detected signal level is weak as compared to the
main reaction zone. These CH selffluorescence signals do not necessarily signify exothermic
reactions. It is known that at high temperatures free radicals survive in medium concentrations
(Algermissen and Noetzold, 1970). Subjected to the high ambient temperatures, these radicals are
thermally excited and, thus, fluoresce.
3.4 Influence of the Gas Content on Evaporation and Combustion
Test runs with pure kerosene, i.e., without hydrogen addition, revealed that for the given fuel
injection pylon the hydrogen addition only improved the jet characteristic in the vicinity of the
injection. This is shown in Figure 10.

FIGURE 10 Influence of the kerosene gas content on the fuel jet structure in the vicinity of the pylon,
photographic image (Mair = 2.15, pair = 0.7 bar, T0, air = 300 K, air = 0.3 kg/s, kerosene = kerosene/H2
= 3.2 g/s, pinject, kerosene/H2 = 10.0 bar, pi
The fuel mass flow rate is the same for both cases. The pure kerosene jet is deflected into the
aerodynamic wake of the injection pylon. The fuel is turned towards the top combustor wall. The
kerosene/H 2 -mixture, on the other hand, shows an improved penetration into the supersonic air
flow. Furthermore, the fuel is distributed more uniformly across the channel cross section.
By injecting the two-phase mixture into the supersonic flow rather than the pure liquid fuel, a
higher momentum ratio q fuel /q air is achieved, as the two-phase mixtures necessitates a higher fuel
injection velocity u fuel to yield the same fuel mass flow rate. This, together with the liquid
dispersion, positively affects the fuel jet structure in the vicinity of the pylon..
Further downstream, in the fuel jet region visualised in Figure 5, no significant evaporation
enhancement could be measured. Also, there was no detectable difference in the ignition and
combustion of the kerosene.
This is in contrast to the experimental observations reported by Avrashkov et al. (1990),
Romankov and Starostin (1993), and Semenov et al. (1998). However, in these articles the fuel
injectors were directed downstream, i.e., the fuel jets were injected more or less parallel to the air
flow direction. In our work the fuel is injected normal to air flow. This leads to the following
conclusion. The mixing process is determined by turbulent action as well as by the fuel dispersion
mechanism. In our case, the shear action of the air stream and the subsequent turbulent mass
transport seem to play the dominant role.
3.5 Comparison of Kerosene and Hydrogen Supersonic Combustion
Figure 9 gave an impression of the supersonic combustion process for Φ = 0.22. This value was
near the maximum equivalence ratio, that could be realised without thermally choking the flow for
the given combustor area ratio A exit /A entrance . To achieve higher equivalent ratios Φ the combustor
area must be increased, by this raising the flow velocity and Mach number so that a higher amount
of heat can be taken up by the supersonic flow. However, as the flow expansion also reduces the
static temperature of the flow, the temperature level at the combustor entrance must be increased. In
other words, the minimum air stream temperature to self-ignite the fuel in the supersonic combustor
T min depends on the combustor geometry, i.e., the combustor area ratio A exit /A entrance .
The relationship between the minimum combustor entrance temperature that ensures stable fuel
self-ignition T min and the combustor area ratio A exit /A entrance is given in Figure 11. The diagram

shows experimentally obtained values for both kerosene and hydrogen supersonic combustion. The
combustor equivalence ratio was held constant at Φ = 0.22.
FIGURE 11 Combustor ignition temperatures in dependence of area ratio Aexit/Aentrance for kerosene/H2
and hydrogen combustion (Φ = 0.22)
On the left-hand side of the graph the flow total temperature is given, on the right-hand side the
corresponding static temperature at the combustor entrance plane. Neither the temperature decrease
due to the flow expansion nor the increase due to the action of the shock waves is included in the
given temperature values (Moreover, there is an additional temperature reduction due to the
injection and mixing of the cold fuel as well as the fuel evaporation in case of kerosene injection).
As expected, the test data for both the kerosene and the hydrogen fuel show the same trend: the
larger the channel expansion, the higher the air flow temperature level that must be provided for
fuel ignition and combustion.
For the case of hydrogen combustion it will be noted, that the combustor entrance temperatures
are well below the hydrogen self-ignition temperature T ignition,H2 . The shaded temperature difference
∆T = T min – T ignition (plus the temperatur decrease due to flow expansion ∆T expansion and fuel/air
mixing ∆T mixing ) is produced by the oblique shock system.
For the case of kerosene combustion the combustor entrance temperatures are above the selfignition
temperatures T ignition,kerosene . Obviously, here the self-ignition temperatures T ignition is not the
critical physical parameter but the ignition delay τ ignition . For the given test parameter the fuel
residence time in the combustor is of the order τ residence = 10 -4 -10 -3 s (compare Table 1). Hence, the
ignition delay should be approximately one order of magnitude smaller, i.e. τ ignition = 10 -5 -10 -4 s.

In Figure 12 the ignition delays for hydrogen and kerosene Jet A-1 are given as a function of
temperature. The data are taken from Algermissen and Noetzold (1970) and Spadaccini and
TeVelde (1980) respectively.
FIGURE 12 Temperature dependence of stoichiometric kerosene/air and hydrogen/air systems (data taken
from Spadaccini and TeVelde (1980) and Algermissen and Noetzold (1970) respectively) as well as the
resulting minimum air stream temperatures for both types of
For a given temperature level the ignition delay of hydrocarbons are much longer than those of
hydrogen. This is true in general even taking into account that the determination of the ignition
delay strongly depends on the specific experimental conditions and assumptions made. In Figure 12
the range of self-ignition delay necessary for combustor operation as well as the resulting minimum
temperature level is given.
It can be seen, that in the case of hydrogen combustion the ignition delay is sufficiently short at
T ignition,H2 ≈ 860 K. For the kerosene/air-system a static temperature level T ≈ 1400 K is necessary.
This is also sketched in Figure 11. Again, as in the case of hydrogen combustion the shaded
temperature difference plus ∆T expansion and ∆T mixing are provided by the action of the oblique shocks.
3.6 Gas Dynamic Feedback Mechanism
It has been shown, that for the given application of a scramjet combustion chamber, wherein the
flame stabilisation relies on fuel self-ignition, physically different fuel properties are responsible for
the combustor ignition:
• For hydrogen the fuel self-ignition temperature T ignition is the critical parameter.
• For hydrocarbon fuels it is the self-ignition delay τ ignition that governs the combustor ignition
process.

This substantial difference results in a completely different operational behaviour of the
supersonic combustor at an air flow temperature range, that is critical for the fuel self-ignition. To
investigate this, for both types of fuel a test series has been conducted, where the air stream
temperature was gradually raised above the temperature threshold which is critical for combustor
ignition.
Figure 13 shows the corresponding experimental results for the case of hydrogen combustion.
The wall static pressure distributions indicate regions and intensity of heat release. Additionally, the
reaction zones visualised through the OH fluorescence are shown.
FIGURE 13 H2 combustion process for increasing air stream temperatures: axial wall static pressure
distributions and OH fluorescence recordings ( air = 0.3 kg/s, ÀÃ ÃÃ$H[LW$HQWUDQFHÃ Ã
At low combustor entrance temperatures (T air = 550 K) there is only a small reaction taking place
at the rear part of the combustor which can be seen from the axial wall pressure distribution.
Apparently, only after the air stream has passed a number shock waves the self-ignition temperature
is reached. Increasing the combustor entrance temperature causes an intensivation and upstreamdisplacement
of the reaction until a final state is reached where the combustion is fully developed.
This can be seen from the wall pressure as well as the reaction zone distribution (The pressure rise
due to combustion is added to the pressure level of the incoming flow which in turn depends on the
flow enthalpy, i.e., the flow temperature. Therefore, the absolute pressure level is still rising for the
increased air flow temperatures even though the combustion process is not intensified any more.).
Thus, for the case of hydrogen combustion, the reaction starts slowly, continually gaining
intensity as the air stream temperature is increased (It should be noted that the combustor ignition
temperature T min for hydrogen combustion, which was given in Figure 11, has been defined as the
air flow temperature where at least 90% of the reaction intensity are reached.). This continuous start
of the hydrogen combustion process is in contrast to the kerosene combustion where the supersonic
reaction abruptly starts. There is a discrete temperature threshold above which the kerosene-fuelled
combustor ignites. Just at this threshold the combustor shows an intermittent combustion mode;

ignition and flame-out alternate several times per second. Stable combustor operation requires a
further small increase of the air stream temperature.
In Figure 14 this fuel-specific combustor behaviour is visualised by plotting the normalised
reaction intensity I/I 0 against the combustor entrance temperature T air .
FIGURE 14 Normalised reaction intensity I/I0 as a function air stream temperature Tair for kerosene and
hydrogen combustion
It can be seen that for hydrogen there is a modest gradient d(I/I 0 )/dT air whereas for kerosene this
gradient is rather strong. This can be explained by means of a gas dynamic feedback mechanism
which is sketched in Figure 15.
FIGURE 15 Gas dynamic feedback loop for the combustor ignition
By raising the air flow temperature the ignition delay τ ignition is shortened (see Figure 12). This
reduces the time scale ratio τ ignition /τ residence where τ residence is the characteristic fuel residence time
within the combustor (The time scale ratio τ ignition /τ residence can also be thought of as a global
combustor Damköhler number). At some temperature level there will be part of the fuel igniting,
thus inputting heat into the flow. The combustion heat release causes the supersonic flow to be
decelerated (increase of τ residence ) as well as a temperature rise (reduction of τ ignition for the remaining
fuel). Both mechanisms reduce the time scale ratio τ ignition /τ residence which, in turn, favours the fuel
ignition, combustion, and heat release, which again leads to an reduction of τ ignition /τ residence , and so
on. This feedback process takes place until an equilibrium state according to the boundary
conditions is reached.

Obviously, the stronger the temperature dependence of the ignition delay, dτ ignition /dT, the more
pronounced this gas dynamic feedback mechanisms will be. For kerosene/air-systems the gradient
dτ ignition /dT is larger than for hydrogen/air-systems (see Figure 12). Hence, for the ignition
behaviour of kerosene-fuelled combustors the gas dynamic feedback is more significant.
Finally, it will be noted that due to the importance of the time scale ratio τ ignition /τ residence for the
supersonic kerosene combustion, there is a hysteresis effect which helps to self-stabilise the reaction
as follows. Once the combustion has been initiated, the time scale ratio τ ignition /τ residence is on the
stable side, i.e., the fuel residence time is sufficiently longer than the ignition delay. Therefore, the
ignition delay τ ignition can become longer, i.e., the air stream temperature can be reduced below the
critical value T min , until at a lower temperature level T flame-out the ratio τ ignition /τ residence switches back
to the unstable side again.
For T flame-out < T air < T min this hysteresis effect can be and was utilised for the ignition of the
combustion chamber. By temporarily introducing a flow disturbance into the supersonic flow − in
our case wall injection of nitrogen in the middle of the combustor (see Figure 1) − oblique shock
waves are generated, which reduce the flow velocity and increase the downstream flow static
temperature as long as the flow disturbance exists (i.e., heat release by chemical reaction and
oblique shocks produce the same thermo-fluiddynamical effects in a supersonic flow). The time
scale ratio τ self-ignition /τ residence reaches the stable side, self-ignition occurs, and now the heat release
into the supersonic flow ensures the time scale ratio to remain on the stable side. The flow
disturbance can be removed without causing a flame-out in the combustor. The same hysteresis
effects has also been reported and utilised for the combustor ignition by Avrashkov et al. (1990),
Avrashkov et al. (1992), Romankov and Starostin (1993), Sabelnikov et al. (1998/1), and Semenov
et al. (1998).
4 SUMMARY
Mechanisms related to the supersonic combustion of liquid hydrocarbons have been investigated.
The following main conclusions can be drawn:
• By the addition of small amounts of hydrogen to the kerosene the liqid fuel jet is dispersed and a
fine spray produced. However, this additional fuel jet dispersion is not necessary for the
supersonic combustion if the fuel is injected normally into the cross flow.
• The reaction is initiated by self-ignition mechanisms. The necessary temperature level is partly
achieved by the oblique shock waves in the supersonic flow. With increasing combustor area
ratio, i.e., expansion of the supersonic flow, the minimum air stream static temperature at the
combustor entrance, which ensures stable combustor operation, rises.
• The choice of fuel − either hydrogen or liquid hydrocarbons − affects the operational behaviour
of the supersonic combustor as far as combustor ignition is concerned. For hydrogen selfignition
the self-ignition temperature is the critical condition to be met. For liquid hydrocarbons
the self-ignition delay governs the combustor ignition. Whereas the supersonic hydrogen
combustion gradually starts when the flow temperature is raised above the self-ignition limit, for
kerosene a discrete air stream temperature threshold exists, above which the combustor fully
ignites.
• This fuel-specific combustor ignition behaviour can be explained by means of a gas dynamic
feedback mechanism.
• Because for kerosene combustion it is the time scale ratio τ ignition /τ residence that governs the
combustor ignition behaviour, the air stream temperature can be reduced below the combustor

ignition level T min once the combustor has ignited. Below T flame-out the time scale ratio
τ ignition /τ residence reaches its unstable regime again and the flame extinguishes.
• The gas dynamic self-stabilisation can be used for the combustor ignition. Oblique shock waves
induced in the supersonic flow are thermo-fluiddynamically equivalent to the heat release by
chemical reactions. Therefore, in the temperature range T flame-out < T air < T min the combustor can
be ignited by temporarily introducing a flow disturbance into the flow.
NOMENCLATURE
A [m²] area
h [m] altitude
I [-] reaction intensity
I sp [s] specific impulse
m& [kg/s] mass flow rate
M [-] Mach number
M 0 [-] flight Mach number
p [bar] static pressure
p 0 [bar] total pressure
q [kg/m s²] momentum
Re [-] Reynolds number
t [s] time
T [%] transmission
T [K] static temperature
T 0 [K] total temperature
u [m/s] velocity
w i [-] mass fraction of species i
x [m] axial co-ordinate
γ [%] gas content (ratio of H 2 mass flow rate to kerosene mass flow rate)
δ [mm] thickness
Φ [-] equivalence ratio
λ [nm] wave length
ρ [kg/m 3 ] density
τ [s] time scale

ACKNOWLEDGEMENT
The authors would like to thank the National German Research Council (DFG) for funding these
research activities. The project was part of the Special Collaborative Research Centre (SFB) 255
„Hypersonic Transport Systems“.
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