Testing of hydrogen-fueled detonation
ramjet in aerodynamic wind tunnel at Mach
1.5 and 2.0
Cite as: AIP Conference Proceedings 2351, 030056 (2021); https://doi.org/10.1063/5.0051978
Published Online: 24 May 2021
V. I. Zvegintsev, V. S. Ivanov, S. M. Frolov, et al.
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AIP Conference Proceedings 2351, 030056 (2021); https://doi.org/10.1063/5.0051978
© 2021 Author(s).
2351, 030056
Testing of Hydrogen-Fueled Detonation Ramjet in
Aerodynamic Wind Tunnel at Mach 1.5 and 2.0
V. I. Zvegintsev1, a), V. S. Ivanov2, b), S. M. Frolov2, c), I. O. Shamshin2, d), and
A. E. Zangiev2, e)
1
Institute of Theoretical and Applied Mechanics, Siberian Branch of the Russian Academy of Sciences
(ITAM SB RAS), Novosibirsk, Russia
2
Semenov Institute of Chemical Physics, Russian Academy of Sciences (ICP RAS), Moscow, Russia
a)
Corresponding author: zvegin@itam.nsc.ru
b)
ivanov.vls@gmail.com
c)
smfrol@chph.ras.ru
d)
igor_shamshin@mail.ru
e)
sydra777@gmail.com
Abstract. The conceptual design of the hydrogen-fueled detonation ramjet (DR) of a new type for a cruising flight speed
of Mach 2 at sea level is developed using 3D numerical simulations of the operation process. The calculated effective
thrust of such a DR is shown to become positive at M = 1.3, i.e., the startup Mach number for such a DR can be lower
than M § 2.0, which is typical for ramjets operating on continuous-deflagration combustion. A DR demonstrator is
designed and manufactured. Its test fires are performed in a blowdown wind tunnel (WT) at free air jet Mach numbers
M = 2.0, 1.5, and 0.9. The result of test fires is the experimental proof of the possibility of arranging stable continuousdetonation combustion of hydrogen in the DR of the developed design at both Mach numbers exceeding 1.0. The
maximum measured values of the fuel-based specific impulse and total thrust were 1610 s and 650 N for the tests with
M = 1.5 and 1630 s and 860 N for the tests with M = 2.0.
INTRODUCTION
Modern high-speed unmanned aerial vehicles are powered, as a rule, with small-size turbojets or ramjets. The
efficiency of small-size turbojets when flying at a speed exceeding the sonic speed is known to decrease sharply
[1, 2]. Thus, one of the most important characteristics of small-size turbojet efficiency, that is the fuel-based specific
impulse equal to the ratio of thrust to the weight flow rate of fuel (kerosene), is approximately 1300 s at a flight with
Mach 1.0 and decreases to ~800 s at a flight with Mach 1.25.
Existing ramjets operating on a thermodynamic cycle with continuous-deflagration combustion of fuel at
constant pressure are effective in the range of aircraft flight Mach numbers ranging from approximately 2 to 6. One
of the most important characteristics of a ramjet-powered aircraft is the minimum Mach number (startup Mach
number) at which the autonomous flight of aircraft is possible after the initial acceleration provided by various
auxiliary boosting means. Existing ramjet-powered aircraft have the startup Mach number approximately equal to 2.
At a lower flight speed, the stagnation pressure of the approaching air stream is too low to ensure stable operation of
the air intake and the engine thrust is insufficient for the autonomous flight of the aircraft. To accelerate aircraft to
such a speed, auxiliary boosters, e.g., rocket engines with relatively low specific thrust performances, are usually
used, which leads to an increase in the starting mass and dimensions of the aircraft. The mass of the booster fuel
increases to about 20-40% of total aircraft mass to accelerate it to Mach 2.
The thrust performances of ramjets have reached their limits and their further increase is problematic. Therefore,
for obtaining a significant increase in the ramjet efficiency, the possibility of using a thermodynamic cycle with
International Conference on the Methods of Aerophysical Research (ICMAR 2020)
AIP Conf. Proc. 2351, 030056-1–030056-12; https://doi.org/10.1063/5.0051978
Published by AIP Publishing. 978-0-7354-4099-9/$30.00
030056-1
continuous-detonation rather than continuous-deflagration combustion of fuel is currently considered [3]. Such a
cycle is known to have a higher efficiency [4, 5]. Also, with continuous-detonation combustion, the chemical energy
of fuel is released in a narrow reaction zone in one or several detonation waves (DWs), continuously rotating in a
combustor. Since the DW at each time instant occupies only a small part of the combustor cross section, this allows
one to arrange the operation process at a relatively low flight speed of the aircraft with only a partial distortion of the
air flow in the intake. Another important advantage of continuous-detonation combustion is the possibility of
arranging the operation process with high combustion completeness in a combustor of a relatively small longitudinal
dimension close to the length of a self-sustaining DW. The disadvantage of continuous-detonation combustion is
obviously its unsteady character. However, due to high propagation velocities of the DWs continuously rotating in
the combustor, pulsations of flow parameters possess high frequencies (kilohertz and higher), and the operation
process appears as quasi-steady.
There are several reviews of physical processes in engines utilizing continuous-detonation combustion often
referred to as rotating detonation engines (RDEs) [6–9]. The first studies of the RDE come back to [10, 11].
Currently, research on RDEs is progressing worldwide. Continuous detonation combustion of hydrogen–oxidizer
mixtures was studied experimentally and computationally elsewhere [12–21] in combustors of different scale and
design. Various modes of continuous-detonation combustion are reported including modes with single and several
DWs rotating in the combustor referred to as the continuous spin detonation (CSD) mode, as well as the
longitudinally pulsating detonation (LPD) mode arising at certain limiting conditions. In the LPD mode, the
detonation is reinitiated at a position close to a combustor outlet and propagates axially upstream as a supersonic
reaction front covering the entire cross-section of a combustor [12, 14]. The gaseous fuels other than hydrogen used
so far in experiments on continuous-detonation combustion with air are ethylene, acetylene, syngas (H2 + CO), etc.
(see, [6, 8]). As for liquid fuels, the pioneering results of experiments on continuous-detonation combustion of
aviation kerosene in air in a detonation afterburner are reported recently in [22, 23].
In the open literature there are two types of publications on the DRs: those powered by the rocket-type RDEs
(see, e.g., [6, 24]) and air-breathing RDEs [25-28]. In our earlier works [29–32], results of calculations and test fires
of large-scale hydrogen-fueled air-breathing DRs with an annular combustor under the conditions of an approaching
supersonic air stream are reported. In [30–32], the test fires of a hydrogen-fueled DR demonstrator are performed in
the short-duration (blowdown) WTs at M = 4-8 with a stagnation temperature of 300 and 1500 K. It is concluded
that the transition from the conventional deflagrative combustion to the detonative combustion in DRs can
potentially resolve the problems of rapid turbulent and molecular mixing of fuel with air and achieving high
combustion completeness at the shortest distances from fuel injectors. A strong shock wave leading the detonation,
either circulating continuously across the airflow in the annular combustor or periodically reinitiating and
propagating upstream the airflow, induces enormous transient shear stresses in cross streams of fuel and air and in
transverse shock waves reflecting from combustor walls causing aerodynamic breakup of fuel jets/sprays and
mixing of reactants. As for chemical transformations, they proceed in a shock compressed gas in the self-ignition
mode at a remarkably high rate.
In view of the specific features of continuous-detonation combustion mentioned above, it can be expected that
the startup Mach number for air-breathing DRs is less than for ramjets operating on continuous-deflagration
combustion. This fundamental issue is addressed in our paper by considering hydrogen as a fuel. Thus, in this work,
we focus on: (1) developing the conceptual design of the air-breathing DR, which could ensure the startup Mach
number significantly less than M = 2 at sea level (to reduce the starting mass of an aircraft with a booster); (2)
design of an air-breathing DR demonstrator; and (3) manufacturing and testing the air-breathing DR demonstrator in
a blowdown WT at approaching air stream Mach numbers M = 2, 1.5, and 0.9.
CONCEPTUAL DESIGN OF DETONATION RAMJET
First, we search for the conceptual design of DR with continuous-detonation combustion of hydrogen–air
mixture, which ensures a flight with a cruising Mach number M = 2 at sea level with a zero angle of attack. As a
result of multivariate gas dynamic calculations (25 DR configurations all in all), the exterior and interior appearance
of the DR shown in Fig. 1 with the rear position of the combustor is formed. The DR consists of a center body with
a front cone, cylindrical part and rear cone, an annular intake with a gasdynamic isolator, and a flow splitter dividing
the incoming flow into two annular channels: outer combustor and inner bypass. The external diameters of the
combustor and the cylindrical part of the center body are 120 and 90 mm, respectively. The width of the annular gap
at the combustor entrance is 7 mm. The internal and external diameters of the annular gap at the entrance to the
030056-2
bypass channel are 80 and 102 mm, respectively. The combustor inlet is displaced downstream from the annular
intake by 60 mm. The length of the combustor and bypass channel is 140 mm. The inner and outer diameters of the
combustor exit are 82 and 120 mm, respectively. The inner and outer diameters of the bypass channel exit are
42 and 80 mm, respectively. It is implied that hydrogen could be injected from the outer combustor wall through a
belt of radial holes located at a certain distance downstream from the combustor entrance.
The length of the DR combustor is chosen based on the implication that it should be about twice longer than the
length of the rotating DW. The geometry of the front part of the center body is selected so that the flow entering the
combustor is uniform except for the near-wall boundary layer, whereas the boundary layer thickness at the entrance
to the bypass channel does not exceed its height. From now on, this conceptual design will be referred to as the dualduct design of the DR [37, 38].
Under flight conditions, the approaching supersonic gas stream is partly decelerated in the oblique shock wave
attached to the front cone of the center body and in the near-wall boundary layer, and then accelerates in the fan of
rarefaction waves with partial recovery of flow parameters and enters the annular intake. The length of the
cylindrical part of the center body is selected so that the fan of rarefaction waves does not enter the intake. At the
entrance to the annular intake, the maximum local flow Mach number is M § 2.2, the mean static pressure is
approximately 0.9ܲ௦௧ , and the total mass flow rate of gas reaches 94% of the value calculated based on the velocity
and density in the approaching gas stream. After entering the annular intake and passing the isolator, the gas flow is
split into two parts: one part enters the annular combustor and the other enters the annular bypass channel. Upon
ignition, a continuous-detonation combustion of the reactive mixture is established in the annular combustor, which
provides acceleration of the detonation products downstream with the formation of a quasi-steady exhaust jet and
the creation of thrust.
FIGURE 1. Dual-duct conceptual design of the detonation ramjet
FIGURE 2. Calculated dependence of the DR effective thrust on the flight Mach number
The DW is inclined to the combustor axis and is terminated downstream by a bow shock. The maximum static
pressure in the DW is about 3.5 MPa. Continuous-detonation combustion in the annular combustor is organized so
that neither combustible mixture nor detonation products penetrate upstream the intake. In the calculations for flight
conditions with Mach 2, a steady-state operation process is obtained with one DW, continuously rotating in the
annular combustor. The tangential and axial components of the DW velocity vector are about 1900 and 520 m/s,
030056-3
respectively, which gives a normal detonation velocity of §1970 m/s. The length of the DW in the axial direction is
about 70 mm, i.e., nearly half the length of the combustor.
The gas flow entering the bypass channel includes a near-wall boundary layer formed on the center body. This
eliminates the negative effect of the boundary layer on filling the combustor with reactive mixture. On the one hand,
the gas flow directed to the bypass channel provides cooling of the combustor internal wall and, on the other hand,
prevents the gasdynamic effect of continuous-detonation combustion of the reactive mixture on the gas flow at the
entrance to the intake. The continuous-detonation operation process is accompanied by generation of gasdynamic
disturbances in the form of shock waves traveling upstream towards the intake, which cause an increase in pressure
at the entrance to the intake. This can lead to the intake unstart and to the failure of continuous-detonation
combustion. Since the front edge of the wall separating the combustor from the bypass channel is displaced deep
into the isolator, such shock waves, when they exit from the combustor, are effectively attenuated and transformed
into weak pressure waves that do not affect much the intake operation but contribute to the creation of additional
thrust when the waves and jets are directed to the bypass channel.
The cross sections of the intake, combustor and bypass channel are selected to ensure a steady-state operation
process with at least one DW and on-design operation of the intake without unstart. The outlet cross sections are
selected to ensure as complete expansion of the detonation products and gas passing through the bypass channel to
the atmospheric pressure as possible.
The effective thrust Feff is defined as the integral of pressure and friction forces over all solid surfaces of the DR
during the continuous-detonation operation process. The total thrust R is defined as the difference between the
effective thrust Feff and the aerodynamic drag force Fd of the DR in the absence of combustion. It should be noted
that the thrust of the DR is created only by 35% of the gas entering the intake, while 65% of the gas expels through
the bypass channel. The calculated value of mass flow rate of gas through the DR under flight conditions with
M = 2.0 is 3.51 kg/s. The steady-state values of the calculated Feff and R for M = 2.0 are 510 and 1040 N
respectively.
CALCULATED STARTUP MACH NUMBER
To determine startup Mach number for the DR, a series of calculations of the operation process in the combustor
with a change in the Mach number of the approaching gas stream is performed. Figure 2 shows the calculated
dependence of the effective thrust of the DR on the flight Mach number. As expected, the maximum value of the
effective thrust (510 N) is attained at M = 2, since the DR is designed precisely for such a cruising flight speed.
Thus, for such a flight speed, the outlet sections of the combustor and bypass channel are optimized to ensure as full
expansion of the detonation products and gas passing through the bypass channel as possible. With an increase in the
flight Mach number, the axial length of the DW increases and at M § 2.7 it becomes comparable with the combustor
length. At M • 2.8, the continuous-detonation operation process is getting unstable: it is blown-off irreversibly
without residual combustion of hydrogen. It follows from Fig. 2 that the calculated effective thrust of the DR
becomes positive at M • 1.3, i.e., the minimum calculated value of the startup Mach number is 1.3.
DETONATION RAMJET DEMONSTRATOR AND TEST RIG
Based on the conceptual design of the DR obtained computationally, we developed design documentation for the
dual-duct DR demonstrator. Figure 3 shows a general view of the DR and a longitudinal section of its 3D model
730 mm long with an annular combustor 120 mm in outer diameter and 210 mm in length.
According to the design documentation, a dual-duct DR demonstrator is manufactured. The walls of the
combustor and the bypass channel made of stainless-steel sheet 1 mm thick are attached to the center body by three
pylons in the front and rear parts of the structure. Hydrogen is supplied to the combustor from the manifold on the
outer wall through a belt of 120 equidistant holes 0.8 mm in diameter at 10 mm downstream from the combustor
entrance. Hydrogen supply pressure in test fires does not exceed 1.1 MPa. To initiate the continuous-detonation
operation process, a detonation initiator fueled by a hydrogen–oxygen mixture is used. A similar initiator was used
in experiments [14, 30–32] and proved its effectiveness.
Test fires of the DR demonstrator are performed in a blowdown WT (Fig. 4) initially designed for tests with free
air jets of Mach numbers up to M = 0.9 [39]. The volume of and the maximum air pressure in the WT receiver are
10.4 m3 and 1.6 MPa, respectively. To obtain a supersonic free air jet, the WT is modified for installing replaceable
converging-diverging profiled supersonic nozzles with design Mach numbers M = 1.5 and 2.0.
030056-4
(ɚ)
(b)
FIGURE 3. 3D model of the DR: (ɚ) general view, (b) longitudinal section
Additionally, calculations of the supersonic air flow around the DR demonstrator placed in free air jets of the
WT nozzles are conducted. The calculated values of the aerodynamic drag force of the DR Fd in free air jets at
M = 2.0 and 1.5 are -700 and -450 N, respectively. Note that the drag of the fairing of the DR demonstrator used to
hide all supports and fittings is not included in these values.
Hydrogen is supplied to the DR demonstrator from a fuel receiver through a system of fast-response valves. The
hydrogen mass flow rate is calculated based on the rate of pressure drop in a receiver 0.64 m3 in volume measured
with a low-frequency static pressure sensor (CourantDA 4.0 MPa).
The DR demonstrator is mounted on the thrust table along the axis of a supersonic nozzle with a zero angle of
attack. The static pressure PC in the combustor is measured with a low-frequency static pressure sensor (CourantDA
1.6 MPa). Pressure pulsations P’C in the combustor are measured by a PCB 113B24 high-frequency pressure sensor.
Both pressure sensors are installed remotely using the waveguide tubes. The static pressure sensor is installed at the
end of a 2-meter long waveguide tube in a gasdynamic damper 180 cm3 in volume. The pressure pulsation sensor is
installed in a waveguide tube 4 mm in diameter at 0.8 m from its fastening to the external wall of the combustor. To
avoid the interference of incident and reflected pressure waves during pressure pulsation measurements, a damping
volume in the form of the same tube 25 m long is installed behind the sensor. Other parameters measured in the test
fires include force F acting on the DR demonstrator measured by the load cell (Tenzo-M T2, ±2000 N), the pressure
at the exit of supersonic nozzle, PST,NOZ, measured by a low-frequency static pressure sensor (CourantDA 250 kPa).
The maximum estimated errors of static pressure and force measurements are 5%. The operation time of the WT
from on to off is about 4 s. During a period from detonation initiation to hydrogen cut-off, the static pressure at the
exit of the WT nozzle changes from 105 to 90 kPa, which ensures the on-design operation of the DR intake.
RESULTS OF TEST FIRES AT ɆACH 2
Below we present the results of test fires of the DR demonstrator at M = 2.0 with a stagnation pressure of 700–
800 kPa and a stagnation temperature of about 270 K. In the test fires, the air-to-fuel equivalence ratio in the
combustor, ĮC, is varied from 0.8 to 2.1. The tasks of the test fires are: (1) to confirm the possibility of arranging
continuous-detonation combustion in the DR demonstrator, and (2) to determine the thrust performance of the DR
demonstrator.
As an example, Fig. 5 shows the primary records of the measured parameters in one of the test fires. It can be
seen from Fig. 5 that by the time the hydrogen supply starts (H2,ON = 1.2 s), the force acting on the DR demonstrator
is negative (-1460 N), and after ignition (IGN = 1.4 ms) it increases up to -600 N. The average static pressure in the
combustor PC increases from 50 to 177 kPa, and regular pressure pulsations of amplitude ~50-100 kPa with a
frequency f ൎ 824 Hz appear in the combustor. The operation process continues until the hydrogen supply is cut off
(H2,OFF = 2,6 s), and then the combustor is purged with air for about 1.5 s. Defining total thrust R as the difference
between the readings of the force sensor after and before ignition, one obtains R § 850 N. Effective thrust Feff,
defined as the sum of total thrust R and the calculated aerodynamic drag force of the DR (Fd § -700 N, see above), is
then equal to Feff § 150 N.
030056-5
(a)
(b)
FIGURE 4. Test rig with the blowdown WT: (ɚ) 3D model and (b) photograph
Fuel-based specific impulse ISP defined as the ratio of total thrust R to the hydrogen weight flow rate, is then
equal to ‫ܫ‬ௌ௉ ൎ 1350 s. Table 1 shows the main parameters and the results of several representative test fires. Records
of sensors in Fig. 5 correspond to test fire No. 2 in Table 1. When ĮC changes from 0.8 to 1.65, total thrust ܴ
decreases from 860 to 605 N, and fuel-based specific impulse ISP increases from 1160 to 1625 s. At ĮC § 2.1,
hydrogen combustion is not registered. The important result of test fires at M = 2.0 is the experimental proof of the
possibility of arranging the steady-state continuous-detonation combustion of hydrogen in the DR demonstrator of
the developed design. In test fires, the near-limit LPD mode, previously observed in [12, 14, 23], and the CSD mode
are registered. In the LPD mode, detonation is periodically reinitiated in a fresh mixture in the vicinity of the
combustor outlet, and the generated DW propagates upstream towards the belt of hydrogen supply holes, causing
longitudinal pressure pulsations with frequency f.
FIGURE 5. Examples of primary records of all sensors measuring flow parameters in one of test fires:
(ɚ) ܲுଶ , (b) ܲௌ்ǡேை௓ , (c) ‫ܨ‬, (d) തതത
ܲ஼ , and (e) ܲԢ஼ .
030056-6
Figure 6 shows fragments of records of the pressure pulsation sensor for the LPD mode (Fig. 6a) and for the
mixed LPD/CSD mode (Fig. 6b). The LPD mode is characterized by a signal with regular triangular pressure
pulsations having a steep front and constant frequency (~800 Hz). In the mixed LPD/CSD mode, the signal has a
more complicated shape: on the background of a frequency of about 850 Hz, low-frequency (~100 Hz) pressure
pulsations with pronounced periods of high-frequency pulsations (2 kHz) are observed. The mixed LPD/CSD mode
occurs when the hydrogen weight flow rate in the DR demonstrator is below a certain value corresponding to
ĮC • 1.9. In this case, due to a decrease in the depth of penetration of hydrogen jets, the mixture formation zone
could be shifted to the outer wall of the combustor, where more favorable conditions are apparently created for the
propagation of CSD. Another possible reason could be the effect of pylons supporting the flow splitter.
TABLE 1. Main parameters and results of DR demonstrator test fires at M = 2.0.
No.
1
2
3
4
5
PH2,
kPa
1065
903
735
574
438
GH2,
g/s
75
63
51
37
29
ĮC
0.80
0.97
1.19
1.65
2.10
PC,
kPa
160
177
176
185
f,
Hz
664
824
1080/2635*
847/1850*
F,
R,
N
N
-600
860
-590
850
-705
755
-855
605
Fuel-lean limit
ISP,
s
1160
1350
1490
1625
Feff,
N
160
150
55
-95
*two pronounced pressure pulsation frequencies were observed.
(a)
(b)
FIGURE 6. Fragments of records of a pressure pulsation sensor in the combustor
(ɚ) LPD mode in test fire No. 2 and (b) for combined LPD/CSD mode in test fire No. 4
In all test fires, the measured value of force ‫ ܨ‬acting on the DR demonstrator is negative (see Table 1). This is,
firstly, due to the high aerodynamic drag of the mounting system of the DR demonstrator on the thrust table, which
is estimated as - 700 N (in addition to the aerodynamic drag of the DR demonstrator itself). Secondly, during highspeed video recording, it is found that the intake of the DR demonstrator operates in an off-design mode. As a matter
of fact, Fig. 7 shows the video frames of test fires Nos. 1 to 4, where hydrogen combustion zones are clearly visible
in front of the entrance to the DR intake. All frames correspond to the steady-state operation of the DR. Presumably
for the same reasons, the fuel-based specific impulse in these test fires attains relatively low values of
ISP = 1160–1625 s.
RESULTS OF TEST FIRES AT ɆACH 1.5
Below we present the results of test fires of the DR demonstrator at M = 1.5 with a stagnation pressure of 300400 kPa and a stagnation temperature of about 270 K. In the test fires, the air-to-fuel equivalence ratio in the
combustor ĮC varies from 0.77 to 1.93. The tasks of the test fires are: (1) to confirm the possibility of arranging
continuous-detonation combustion in the DR demonstrator of the developed design at an off-design flight Mach
number of M = 1.5, and (2) to determine the thrust performance of the DR demonstrator at such a speed.
Table 2 shows the main parameters and results. When ĮC changes from 0.77 to 1.6, total thrust ܴ changes from
605 N at ĮC = 0.77 to 440 N at ߙ஼ = 1.6 with the maximum value of 650 N at ĮC = 0.97, and fuel-based specific
impulse ISP varies from 1040 s at ĮC = 0.77 to 1560 s at ĮC = 1.6 with the maximum value of 1610 s at ĮC = 1.41.
The absolute values of total thrust R decreases by approximately ~200 N as compared to test fires at M = 2.0. At
M = 1.5, effective thrust Feff is positive in almost the entire range of ĮC. Thus, it is experimentally proved that the
startup Mach number for the DR can be on the level of and even less than M = 1.5, that is, the calculation is
030056-7
confirmed by experiment (see Fig. 2). The important result of the test fires at M = 1.5 is the experimental proof of
the possibility of arranging the steady-state continuous-detonation combustion of hydrogen in the DR of the
developed design at an off-design flight speed. In the test fires with 0.77 ” ĮC ” 1.6, the near-limit LPD mode is
registered. The maximum values of the frequency of longitudinal pressure pulsations (݂ = 700–730 Hz) are obtained
at 0.77 ” ĮC ” 1.05. At ĮC > 1.05, the pressure pulsation frequency first decreases to 550–600 Hz, and at ĮC • 1.93,
hydrogen combustion is not registered.
Apparently, one of the reasons that the effective thrust of the DR demonstrator at M = 1.5 turns out to be higher
than at M = 2.0 is more stable operation of the DR intake. Figure 8 shows the frames of video recording
corresponding to the DR steady-state operation in test fires Nos. 1, 3, 5, 6, 7, and 8 at M = 1.5. If, with ĮC = 0.77, the
glow of combustion products is clearly visible at the entrance to the DR intake, then with ĮC • 0.83 this effect is
absent.
RESULTS OF TEST FIRES AT ɆACH 0.9
In test fires with M = 0.9 (stagnation pressure and temperature are 150 kPa and about 270 K, respectively), no
steady-state operation of DR demonstrator is attained. Tests show that mixture ignition in the combustor leads to
intake unstart and the development of external combustion of hydrogen–air mixture around the DR demonstrator
(Fig. 9). Readings of the static and pulsation pressure sensors indicate no combustion inside the DR. Decrease in
hydrogen weight flow rate does not lead to visible improvements in the operation process due a drop in heat release.
One of the possible solutions for an efficient design of detonation ramjet for flight conditions at M ” 1.0 is the use of
pulsed detonation engines [39].
TABLE 2. Main parameters and results of DR demonstrator test fires at M = 1.5.
No.
1
2
3
4
5
6
7
8
9
ܲ‫ ʹܪ‬, kPa
‫ ʹܪܩ‬, g/s
ߙ‫ܥ‬
886
847
803
714
655
589
513
455
387
60
58
56
48
44
39
33
29
24
0.77
0.80
0.83
0.97
1.05
1.19
1.41
1.60
1.93
തതതത
ܲʙʠ ,
kPa
130
130
129
139
140
146
141
134
݂, Hz
‫ܨ‬,
N
706
732
732
706
732
601
549
549
-395
-385
-355
-350
-370
-410
-480
-560
ܴ,
N
‫ ܲܵܫ‬,
s
605
1040
615
1080
645
1180
650
1385
630
1465
590
1530
520
1610
440
1560
Fuel-lean limit
FIGURE 7. Frames of video records of test fires Nos. 1 to 4 at Ɇ = 2.0
030056-8
‫ ݂݂݁ܨ‬,
N
155
165
195
200
180
140
70
-10
FIGURE 8. Frames of video recording of test fires Nos. 1, 3, 5, 6, 7 and 8 at Ɇ = 1.5
FIGURE 9. Frames of video recording of test fires at Ɇ = 0.9
DISCUSSION OF EXPERIMENTAL RESULTS
Comparison of Tables 1 and 2 shows that with an increase in the Mach number of the approaching air stream the
pressure in the combustor increases from 130–145 kPa at M = 1.5 to 160–180 kPa at M = 2.0. For fuel-rich and
near-stoichiometric mixtures with 0.77 ” ĮC ” 1.0, the frequency of longitudinal pressure pulsations in the near-limit
LPD mode is 700–800 Hz both at M = 1.5 and at M = 2.0. At M = 2.0 and ĮC • 1.2, due to the appearance of the
mixed LPD/CSD mode in the combustor, the frequency of longitudinal pressure pulsations in the LPD mode
increases to 800–900 Hz and a pressure pulsation mode with a frequency of 2600–1800 Hz, corresponding to the
CSD mode, is added. In this case, an irregular effect is manifested: despite an increase in ĮC (decrease in fuel
supply) and a decrease in thrust, the pressure in the combustor increases. This effect could be attributed to nonlinear
pressure piling in the waveguide tube, at the end of which a static pressure sensor is installed [40] and will be
addressed in future studies.
In general, the results obtained demonstrate the possibility of decreasing the Mach number of DR operation to at
least M = 1.5. This possibility is associated with a combination of two factors: the use of a dual-duct flow path and
continuous-detonation combustion in the DR. As for the first factor, even with ĮC = 1, the total air-to-fuel
equivalence ratio at the entrance to the DR intake is close to 3, i.e., most of the air passes through the bypass channel
without participating in combustion, and the heat release in the combustor does not lead to intake unstart even at a
relatively low flight speed. As for the second factor, the heat release zone in the DW at each instant of time occupies
only a small part of combustor cross-section, which contributes to the on-design operation of the DR intake.
030056-9
Figure 10 shows the generalized dependences of thrust performance on the air-to-fuel equivalence ratio for
M = 1.5 (Fig. 10a) and M = 2.0 (Fig. 10b). Circles and squares correspond to fuel-based specific impulse (right
ordinate axis) and total thrust (left ordinate axis), respectively. The dashed horizontal line shows the calculated drag
force Fd of the DR demonstrator. The difference between the thrust curve and Fd line corresponds to estimated
effective thrust Feff. Comparison between Figs. 10a and 10b indicates that the effective thrust becomes negative at
ĮC • 1.3 for M = 2.0 and is positive at all values of ĮC almost up to the blow-off conditions for M = 1.5. The reason
for thrust decrease at M = 2.0 is the unstable intake operation with losses in pressure and air flow rate (see Fig. 7).
The fuel-based specific impulse at M = 1.5 and 2.0 is almost the same despite its increase was expected for M = 2.0
due to elevated stagnation pressure. In addition to test fires for the basic DR design of Fig. 2, there were tests for a
DR demonstrator with a combustor of shorter length (by 80 mm) but with same exit areas. The operation process in
such a DR was much more unstable, and the domain of stable operation process with detonative combustion in terms
of hydrogen weight flow rate was narrow. Open squares and circles in Fig. 10 show the total thrust and specific
impulse for such a combustor. A shorter combustor is seen to exhibit lower thrust performance presumably due to
lower combustion efficiency.
(a)
(b)
FIGURE 10. Measured dependences of the fuel-based specific impulse, total thrust, and effective thrust of the DR demonstrator
on the air-to-fuel equivalence ratio for (a) M = 1.5 and (b) M = 2.0
CONCLUSIONS
The conceptual design of a hydrogen-fueled dual-duct DR for the cruising flight speed of Mach 2 at sea level is
developed. The calculated effective thrust of such a DR is shown to become positive at M = 1.3, i.e., its startup
Mach number can be lower than M = 2.0, which is typical for ramjets with continuous-deflagration combustion.
A DR demonstrator for cruising flight speed of Mach 2 at sea level is designed and manufactured. Its test fires
are performed in a blowdown WT at approaching air stream Mach numbers M = 2, 1.5, and 0.9. The most important
result of the test fires at M = 2 is the experimental proof of the possibility of arranging steady-state continuousdetonation combustion of hydrogen in the DR of the developed design. The near-limit LPD mode and the CSD
mode are registered.
The important result of the test fires at M = 1.5 is the experimental proof of the possibility of arranging steadystate continuous-detonation combustion of hydrogen in the DR of the developed design at off-design flight speed.
Thus, it is experimentally proved that the startup Mach number of the DR can be at the level of and even less than
M = 1.5. In the test fires with M = 1.5, the near-limit LPD mode is registered.
For all the Mach numbers, the thrust performance and economic characteristics of the DR are obtained. The
maximum measured values of the effective thrust, as well as the estimated values of the total thrust and fuel-based
specific impulse are 160 N, 860 N and 1630 s for test fires at M = 2 and 200 N, 650 N and 1610 s for test fires at
M = 1.5, respectively. The effective thrust of the DR demonstrator is shown to be positive for the air-to-fuel
equivalence ratio in the combustor ĮC < 1.3 at M = 2.0 and ĮC < 1.6 at M = 1.5.
030056-10
The experimental results obtained demonstrate the possibility of decreasing the Mach number of DR operation to
at least M = 1.5. This possibility is associated with a combination of two factors: the use of a dual-duct design and
continuous-detonation combustion in the DR. As for the first factor, even with ĮC = 1, the overall air-to-fuel
equivalence ratio at the entrance to the DR intake is close to 3, i.e., most of the air passes through the bypass channel
of the DR without participating in the combustion process, and the heat release in the combustor does not lead to
intake unstart even at a relatively low flight speed. As for the second factor, the heat release zone in the DW at each
instant of time occupies only a small part of the combustor cross-section, which contributes to the on-design
operation of the DR intake. Another important advantage of continuous-detonation combustion is the possibility of
arranging an operation process with a high completeness of combustion in a combustor with a relatively small
longitudinal dimension corresponding to the length of a self-sustaining DW. This allows one to significantly reduce
viscous losses in the engine path compared to a ramjet with continuous-deflagration combustion.
ACKNOWLEDGMENTS
This work is supported by Russian Science Foundation (project 18-73-10196).
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