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infrared integrated stealth optimization design of helicopter engine intake and exhaust system
Aerospace Science and Technology 95 (2019) 105483
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Radar/infrared integrated stealth optimization design of helicopter engine intake and exhaust system Zeyang Zhou ∗ , Jun Huang, Jinjun Wang School of Aeronautic Science and Engineering, Beihang University, Beijing 100191, China
a r t i c l e
i n f o
Article history: Received 15 July 2019 Received in revised form 8 October 2019 Accepted 13 October 2019 Available online 17 October 2019 Keywords: Radar cross section Infrared radiation Comprehensive stealth Helicopter Engine intake and exhaust system
a b s t r a c t Along with the diversification and sophisticated development of detection methods in modern warfare, helicopters are increasingly subject to unilateral or simultaneous threats from radar and infrared detectors. In order to improve the survivability and operational effectiveness of the helicopter, a comprehensive stealth approach based on Pareto solution is presented. Considering the geometric constraints and aerodynamic characteristics of the engine intake and exhaust system, the model of the system is established by the full factorial design, the internal, central and external flow fields are constructed, then the highprecision computational fluid dynamics method is used to simulate the total flow field under the rotor downwash airflow in hovering state. The radar cross section of the system is evaluated by the physical optics and physical theory of diffraction. Based on the Monte Carlo and ray tracking method, the infrared signature of the system is calculated and analyzed in detail. Under the comprehensive evaluation and selection of comprehensive stealth approach, the optimization model of the system is continuously established and updated. The ultimate design has achieved good results in both radar cross section reduction and infrared radiation suppression and the proposed method is effective and efficient for radar/infrared integrated stealth of helicopter engine intake and exhaust systems. © 2019 Elsevier Masson SAS. All rights reserved.
1. Introduction With the continuous replacement and joint application of infrared (IR) imagers and radar detection equipment in modern warfare [1], the survivability of helicopters has never been more challenging. The unilateral adoption of RCS reduction or IR suppression technology cannot guarantee the comprehensive stealth performance and integrated operational effectiveness of helicopters [2,3]. Therefore, it is very important to carry out research on helicopter radar/infrared integrated stealth. The radar stealth performance of aircraft is affected by many factors, such as surface configuration, tilt angle, azimuth and radar band [4,5]. In order to achieve low radar scattering of the target helicopter, physical optics (PO) and method of equivalent current are used to calculate radar cross section (RCS) of the target [4]. Some reduction schemes based on radar absorbing material have been proposed and used to improve the radar stealth performance of helicopter air intakes [6,7]. The singularity of physical diffraction theory has been studied and used to solve wedge diffraction, while the calculation results are compared with the triangular column
*
Corresponding author. E-mail address: [email protected] (Z. Zhou).
https://doi.org/10.1016/j.ast.2019.105483 1270-9638/© 2019 Elsevier Masson SAS. All rights reserved.
RCS and the data of method of moment [8,9]. The electromagnetic performance of the radome is analyzed by the method of physical optics [10]. Facing the continuous development of radar systems, the improvement of helicopter survivability is becoming more and more important, and radar stealth has become one of the significant tactical indicators of modern helicopters. Aircrafts or helicopters sometimes fly at low altitudes or shuttle between obstacles to avoid the radar system, but they will still be the target of infrared imaging equipment or heat-seeking weapons [11,12]. The suppression methods such as concealing, camouflage, and downwash are implemented or experimented to study the infrared characteristics of helicopters [13–15]. Multistage ejector nozzles are tested and investigated to learn the infrared signature suppression [16,17]. Based on the basic principle of heat source method of heat transfer, an approach for calculating the temperature field of helicopter is proposed [18]. Different nozzle shapes are tested to study the performance of the small turbojet engine and the characteristics of infrared radiation [19]. The reverse Monte Carlo (MC) method and ray tracing method are used to calculate the infrared characteristics of the ejector-mixed lobes infrared suppressor [20–22]. In modern local wars, infrared countermeasures are intensifying and infrared imaging is becoming more and more accurate, so the helicopter’s excellent infrared stealth capability is particularly important.
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Fig. 1. Schematic of radar/infrared integrated stealth for engine intake and exhaust system radar.
However, the complexity and diversity of modern and future battlefields [23–26] have greatly reduced the viability of traditional helicopters [27–31]. The aircrafts with unilateral stealth performance will still be targeted when facing both radar and thermal imaging or more detection equipment [32–35]. In order to pursue a good exhaust cooling effect [23,24], the design and installation of the infrared suppressor will change its shape characteristics at the installation of the helicopter body [42,43], thereby affecting the electromagnetic scattering characteristics of the aircraft under certain azimuth angles [25,26]. The helicopter full stealth has to emphasize the radar stealth design of various components and systems in important or even omnidirectional angles [29,30], which will also affect the intake/exhaust and cooling effects of the infrared suppressor [31–34]. Therefore, this paper attempts to develop a radar/infrared integrated stealth study of helicopter infrared suppressors to explore the interaction between radar stealth and infrared stealth of the engine intake/exhaust system. It is increasingly important to enhance the integrated stealth capability of helicopters [44–47], especially radar and infrared comprehensive performance. This manuscript presents the optimization method in section 2. The models and calculation patterns are provided in section 3. After discussing the results, the paper is concluded in section 5. 2. Comprehensive stealth approach The addition of the infrared suppressor is mainly to improve the IR radiation characteristics of the engine high-temperature exhaust by ingesting the surrounding cold air including the rotor downwash flow and the airflow in front of the helicopter as shown in Fig. 1, but this also affects the radar stealth performance of the helicopter rear fuselage where α is the azimuth angle and β is the elevator angle. The integrated design of the infrared suppressor increases the size of original slim rear fuselage beam, especially the secondary ejector mixing device. In order to improve the infrared/radar integrated stealth performance of the engine intake and exhaust system, the multiobjective function is defined as follows:
min : f i (m) ,
i = 1, 2, 3
M = {m1 , m2 , ..., mn }
(1)
where m is the model of the system, f i (m) represents the i-th target performance, and M refers to the set of different m.
Each target performance could be described as follows:
⎧ ⎪ ⎨ f 1 (m) = σ f 2 (m) = I n ⎪ ⎩ f 3 (m) = I h
◦ 0 ≤ α ≤ 180◦ , β = 0◦ −30◦ ≤ αn ≤ 30◦ HH
(2)
where αn is the observation angle in the normal observation field, HH indicates that the radar wave is horizontally polarized [36], the radar wave frequency is 10 GHz, σ is the far field RCS mean of m in the horizontal observation field of 0◦ < α < 180◦ , I n is the mean of infrared radiation intensity of m in the normal observation field of −30◦ < αn < 30◦ , I h is the mean of infrared radiation intensity of m in the horizontal observation field of 0◦ < α < 180◦ . 2.1. Constraints The equality constraints could be expressed as follows:
⎧ ⎪ ⎨ R 0 ( M i ) − R 0 M j = 0 L0 (M i ) − L0 M j = 0 ⎪ ⎩ R 1 (M i ) − R 1 M j = 0 ⎧ ⎪ j =0 ⎨ M b ( M i ) − Mcb M L1 (M i ) − L1 M j = 0 ⎪ ⎩ V 0 (M i ) − V 0 M j = 0
&
∀i , j ∈ {1, 2, ..., k}
(3)
where R 0 is the inner radius of the first stage lobe inlet, L 0 is the length of the first stage lobe straight section, R 1 is to the outer radius of the first stage lobe, M b is the central body model of the first stage lobe entrance, L 1 is the first stage lobe length, V 0 is the gas velocity of the first stage lobe entrance and k is the optimized generations. The inequality constraints could be expressed as:
⎧ ⎪ ⎨ L ( M i ) − L max ≤ 0 H ( M i ) − H max ≤ 0 & ⎪ ⎩ W ( M i ) − W max ≤ 0 T M j +1 − T M j ≤ 0 ∀i ∈ {1, 2, ..., k} V M j +1 − V M j ≥ 0 ∀ j ∈ {1, 2, ..., k − 1}
(4)
where L, H and W respectively represent the length, height and width of m, L max , H max and W max are respectively the corresponding size limits, T refers to the average temperature of the nozzle end face and V is the average velocity of the nozzle end face exhaust airflow.
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2.2. Optimal scheme and Pareto solution When the optimal solution exists, it has the following form (COM test):
⎧ f i (M ) ≥ f i (M ∗ ) ⎪ ⎪ ⎨
T M ∗j +1 ≤ T M ∗j ⎪ ⎪ ⎩ V M j +1 − V M j ≥ 0
∀ i ∈ {1, 2, 3} ∀ j ∈ {1, 2, ..., k − 1}
(5)
where M ∗ is the optimal model and satisfies all equality and inequality constraints. When the multi-target performance cannot obtain the minimum value at the same time, the feasible solution (Pareto solution) has the following form:
⎧ f i (M ∗) ≤ f i (M ) ⎪ ⎪ ⎪
⎨ ∗ T M j +1 ≤ T M ∗j ⎪
⎪ ∗ ⎪ ⎩V M V M ∗j ≥ j +1
∃ i ∈ {1, 2, 3} ∃ j ∈ {1, 2, ..., k − 1}
Fig. 2. Validation of the PO+PTD algorithm on the RCS calculation model of the IR system.
E s (r ) =
S
(6)
×
This indicates that there is at least one value of i or j satisfy the constraints of the comprehensive stealth performance.
1
E (r ) =
j ωε · 4π
S
3 − k2 R 2 + j3kR − jkR e R × R × J s r 5 R
1 + jkR
+2 Js r H (r ) =
1 4π
R3
−1 − jkR R3
S
e − jkR dS
(7)
e − jkR R × J s r
S
dS
(8)
e − jkR
Js =
×e
, Zi
0
, Zd
nˆ r · E 0 kˆ − nˆ r · kˆ E 0
− jk −ˆr +kˆ ·r
I=
ˆ × Rˆ × J s r R
R
dS
(9)
(10)
where n is the unit normal vector of the outer normal direction of r at the surface of the scatterer, Z i is the illuminated area, and Z d is the dark area. When the incident wave is a plane wave:
(13)
nˆ r · kˆ Eˆ 0 e − j2k·r dS
(14)
S
σ=
4π
λ2
| I |2
(15)
For solution of the edge diffraction, the diffraction coefficient of PTD could be found in Ref. [30,37]. The validation of the PO+PTD algorithm for RCS calculation is given in Fig. 2, showing that the two RCS curves are generally consistent including trend, mean, peak size, and peak position. The average value of RCS in the 0-180◦ azimuth of the PO+PTD curve is 0.4145 dB.m2 smaller than that of FEKO (PO+MOM), which indicates that the RCS calculation result is reliable.
The solution of the hot and cold air mixing problem in the intake and exhaust system is the basis for the calculation of infrared radiation. According to the radiation transfer equation (RTE):
ds
+ (a + σ S ) I (r , s)
σ T 4 σS = an + π 4π
4π
2
E i r = | E 0 | e − jk·r
dS
Considering the characteristics of plane waves, there are:
dI (r , s)
2n × H
(12)
2.4. Infrared radiation calculation
Based on the assumptions of physical optics, there are:
dS
RCS value could be determined as:
where r is the coordinate vector of the field point, r is the coordinate vector of the source point, and R is the distance between the field point and the source point, ω is the electromagnetic wave angular frequency, is the dielectric permittivity. Then E (r ) could be transformed into the following form:
−k2 E s (r ) = j ωε · 4π
R
rˆ × rˆ × S
The far-field RCS index of the system could be calculated by PO+PTD (physical theory of diffraction, PTD) method. According to the Stratton-Chu integral equation, the electric and magnetic fields could be obtained as follows:
e
− jk R +kˆ ·r
Then the integral term could be recorded as the following form:
I=
2.3. RCS calculation method
−k2 j ωε · 2πη
ˆ × Rˆ × nˆ r · E 0 kˆ − nˆ r · kˆ E 0 × R
(11)
where k refers to the wave vector. Considering the approximate condition of PO, the electric field formula can be written as:
I r , s s, s d
(16)
0
where r is the position vector, s is the direction vector, s is the length along the path, a is the absorption coefficient, s is the scattering direction vector, σ S is the scattering coefficient, n is the refractive index, I is the radiation intensity, σ is the StefanBoltzmann constant, T is the local temperature, is the scattering phase function, is the solid angle.
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Assuming that the reaction in the combustion chamber is complete, the scattering of the medium is not considered and the refractive index is equal to 1, the RTE could be simplified to:
dL (r , s) ds
+ aL (r , s) = a
σT4 π
(17)
Calculation of thermal radiation using discrete coordinates model:
∇ · [L (r , s) s] + aL (r , s) = a
σT4 π
(18)
Based on the Monte Carlo and ray tracing method, the spectral radiance of the detector receiving point is as follows:
L σ = L 0σ τ1σ τ2σ · · · τnσ + L b1σ (1 − τ1σ ) τ2σ τ3σ · · · τnσ + · · ·
+ Lnbσ−1 1 − τ(n−1)σ τnσ + Lnbσ (1 − τnσ )
(19)
where L 0σ is spectral radiance of wall reverse rays, L bi σ is spectral radiance, τi σ is i-th layer spectral transmittance. The irradiance of the detector receiving point could be determined as:
E=
N NB
L nσ ,(i , j ) · cos θ j · j · σi · 100
(20)
i =1 j =1
where E is the radiant illumination, N B is the total number of wave bands, N is the total number of rays contributing to the measurement point, θ j is the angle between the center of the j-th solid angle and the surface normal of its measuring point, j is the j-th solid angle, σi is the width of the i-th wave band. Then the radiation intensity at the receiving point is:
Iσ = Eσ (R) · R2
(21)
where I is the radiation intensity and R is the linear distance between the helicopter and the detector. Due to the continuous heating of high temperature gas, the temperature of the exhaust pipe nozzle is high and its infrared radiation is prominent. For a given infrared band, the radiance and radiance are calculated as:
⎧ ⎪ ⎪ ⎪ ⎨ ⎪ ⎪ ⎪ ⎩
λ2
ε ε L λ = M λ1 ∼λ2 = π π
M λ dλ
(22)
λ1
I λ = L λ A cos θ
where λ is the wavelength, M λ1 ∼λ2 is the radiation emission degree of the black body between the bands λ1 ∼ λ2 , A is the area of the nozzle bin, M λ is the spectral radiation emission degree of the black body. Introducing the Planck formula:
Mλ =
c1
1
(23)
λ5 ec2 /(λT ) − 1
Then the radiation emittance of blackbody in λ1 ∼ λ2 band could be obtained as:
M λ1 ∼λ2 =
15σ T 4
c 2 /(λ 1T )
π4 c 2 /(λ2 T )
c2 =
hc k
,
basically consistent with the data points of the literature [25], including the start and end positions, the maximum value, the mean value, and the turning point where scale of the engine exhaust model is 1.5, indicating that the presented calculation method is credible and accurate. 2.5. Method flow chart Fig. 4 shows that the integrated stealth optimization design process of the whole system mainly includes three parts: initial model design and performance evaluation; flow field, RCS and infrared radiation calculation of the model population, comprehensive performance evaluation and update of optimal models; ultimate model determination, method termination and recycling. 3. Model design and calculation method Taking the RAH-66 stealth helicopter as the reference object [30,39], the model of the infrared/radar integrated stealth engine intake and exhaust system is designed in a 1:1 ratio as shown in Fig. 5, where M f represents the front air intake model of the system, M b is the central body model at the first stage inlet, M c is the circular lobe model and acts as the first stage blending structure, and M r refers to the rectangular lobe acts as a second-stage blending structure, and D e is the normal direction of the nozzle end face. Ambient air enters the engine compartment through the front air inlet, and the hot gas after combustion reaches the center body at the entrance of the first-stage lobes [38], and mixes with the cold air taken in by the first-stage air inlet to flow to the second-stage lobes. The cold air taken in from the two intake ports of the second stage is again mixed and discharged outward through the nozzle, thereby achieving the effect of suppressing infrared radiation of the exhaust gas. 3.1. Geometric model establishment
x3 ex − 1
dx
(24)
The radiation constants have the following relationship:
c 1 = 2π hc 2 ,
Fig. 3. Ref. [25] data to verify the presented IR signature calculation method.
σ=
2π 5 k 4 15c 2 h3
(25)
where h is the Planck constant, h = 6.626176 × 10−34 J·s, k = 1.38 × 10−23 J/K, c is the speed of light. The validation of the presented IR calculation method is shown in Fig. 3. The IR curve is
According to the design of the system, the air inlet is opened on the side of the fuselage, the rear half of the back and the upper surface of the tail beam as shown in Fig. 6. The exhaust port is opened on the side edge of the tail beam and the nozzle end face is parallel to the horizontal plane. The infrared suppressor is designed in the inner space of the fuselage. The model of this system could be represented by the following parameters or sub modules:
m = M b , M f , Nl , M c , W h , W t , M r , D e
(26)
Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
Fig. 4. Infrared/radar integrated stealth optimization method flow chart.
Fig. 5. Comprehensive stealth model design for the engine intake and exhaust system.
Fig. 6. The model of the infrared/radar integrated stealth system.
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Fig. 7. Flow field construction and meshing of the system.
Fig. 8. Ashok experiment results of the IR suppression device to verify the CFD method [17].
Fig. 9. Verification of grid number and convergence time, Nl = 1.
where Nl is the number of lobe stages, W h is the head width of the tail beam upper surface, and W t is the tail width of the tail beam upper surface. 3.2. Computational fluid dynamics method For the temperature distribution characteristics of the system in the hovering state, the whole flow field is constructed and divided into three parts as shown in Fig. 7. The external flow field is a large cylindrical as a whole and the top surface has a conical design. The central flow field is a flat cylindrical, including the paddle plane and the fuselage surface. The fuselage surface excluding the ducted tail rotor adopts 1:1 modeling and the paddle plane is used to simulate the down washing effect of the rotor. The internal flow field is the internal space of the infrared suppressor, including the first/secondary lobe, the center body, the inlets, the exhaust port and the first/secondary exhaust pipe. High-precision unstructured grid technology is used to divide each region, and the quality and quantity of the grid are strictly controlled to ensure the convergence of the calculation results of the computational fluid dynamics method. The grid size is set as shown in Table A1 of the Appendix A. The Ashok experiment uses a simple 5-stage round tube ejector as an example to study infrared radiation suppression [17]. The CFD method in this paper simulates the suppressor system as shown in Fig. 8, and the results of variation of mass entrainment with nozzle distance from bottom funnel agree well with the experimental data where H overlap / D nz = 4 and Re = 7073. Fig. 9 presents the verification of grid number and convergence time of the CFD method where the total number of grids increased from 3 million to 8.12 million at unequal intervals. The total number of iteration steps is 10,000 steps and is guaranteed to be completed within 24 hours or it is considered not to converge. The convergence time is replaced by the number of steps at which convergence begins. It could be seen that the temperature T and convergence steps begin to stabilize from the 4.91 million grid. In order to ensure the reliability of the CFD method and the speed of
Fig. 10. RCS evaluation method of the closed system. (For interpretation of the colors in the figure(s), the reader is referred to the web version of this article.)
the entire optimization process, the number of grids in this paper is controlled between 5.2 and 5.7 million, and the total number of iteration steps is 10,000 steps. 3.3. RCS evaluation method To reduce the number of grids in the RCS calculation, the closed model of the system is extracted as shown in Fig. 10, where the blue part maintains a certain gap with the corresponding fuselage skin to avoid direct heating of the skin. 3.4. Infrared radiation observation field setting According to the indexes of infrared radiation intensity, two observation fields are set as shown in Fig. 11. One of the observation fields is located in the horizontal plane and the other observation field is perpendicular to the x-axis where the nozzle end face is designed to be parallel to the horizontal plane. 4. Results and analysis The different front air intake models are shown in Fig. 12, including the inclined triangular air inlet (M f 1 ), the smooth convex
Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
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Fig. 11. Infrared radiation observation field setting diagram.
Fig. 12. Different front inlet models for RCS calculation.
Fig. 13. RCS distribution of the system model under M f 1 .
curved air inlet (M f 2 ), the trapezoidal air inlet (M f 3 ), the curved air inlet with forward lip (M f 4 ) and the tilted polygonal air inlet (M f 5 ). The M f 1 is similar to the diamond inlet design and does not form a strong reflection on the forward radar wave. The RCS results of the system model with M f 1 are shown in Fig. 13. The RCS peak appears at azimuth 270◦ because of the plane reflection of the vertical left side plate, but this part surface is hidden inside the fuselage and does not contribute to the fuselage RCS. A large RCS peak occupies a size of 15.5447 dB.m2 in the lateral 92.45◦ , which has a significant effect on the radar scattering characteristics of the fuselage. 4.1. Radar stealth optimization Fig. 14 supports that the RCS peak of the M f 3 curve is the largest and the overall value is also large, which is mainly caused by its mediocre trapezoidal shape design. The peak of M f 2 is similar to that of M f 1 , but the RCS of M f 2 is significantly higher than that of M f 1 in the azimuth range of 110◦ ∼150◦ . In the azimuth range of 0◦ ∼20◦ and 110◦ ∼120◦ , the RCS distribution of M f 5 model is obviously higher than that of M f 1 . The RCS values of M f 4 are larger than that of M f 1 and M f 5 in the azimuth angles
Fig. 14. The system model RCS comparison between various M f .
Table 1 RCS indicators of the system model with various M f .
σ /dB.m2
Mf1
Mf2
Mf3
Mf4
Mf5
−5.4302
−3.4146
0.0856
−1.8767
−3.8909
of 0◦ ∼30◦ and 75◦ ∼160◦ , because both M f 1 and M f 5 use a tilted plane design and M f 4 is a regular surface configuration. Table 1 presents the RCS indicators of the system model with various M f . The initial design of the front air intake model has a large impact on the RCS index, and the maximum difference could reach 5.5158 dB.m2 . The M f 3 suppressor model harvests the largest RCS mean in the range of 0◦ to 180◦ azimuth of the horizontal field, up to 0.0856 dB.m2 because its trapezoidal air intake side panel produces more strong scattering sources. M f 1 performs best as a result of its stealth design and the weaker edge diffraction. Fig. 15 reveals that the overall RCS curves of W h1 and W h2 are not much different, and the RCS peaks are almost equal and the peak positions are consistent. Compared with the RCS curve of W h3 , that of W h4 shows more minimum values in the azimuth
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15.4374 dB.m2 . The value of W t determines the tail width and inclination of the upper surface of the tail beam, affecting the radar wave illumination area and the angle between each bin and the radar wave. When both W h and W t increase a lot, it is a conventional non-radar stealth design, which results in very strong specular scattering of the tail beam side plates. 4.2. Infrared radiation suppression
Fig. 15. The system model RCS comparison between various W h .
Table 2 RCS indicators of the system model with various W h .
Value/mm σ /dB.m2
W h1
W h2
W h3
W h4
467.142 −5.4302
457.142 −5.3983
477.142 −5.4778
487.142 −5.4598
Fig. 16. The system model RCS comparison between various W t .
angles of 0◦ ∼30◦ and 110◦ ∼150◦ . But overall, the four RCS curves have no substantial drastic reduction effect. Table 2 manifests that the value of W h2 is reduced by 10 mm compared with that of W h1 , while its σ value reaches −5.3983 dB.m2 . As the value of W h increases, the RCS index first decreases and then increases, and a minimum value of σ is obtained at 477.142 mm. The size of W h affects the inclination of the side panels of the tail beam and the upper side panels on the rear fuselage, while this side panels will form a strong surface scattering source and more edge diffraction, which in turn affects the lateral RCS distribution of the system model. Fig. 16 provides that the four RCS curves have basically the same trend, and the peak occurs at the azimuth angle of 94.50◦ , wherein the RCS peak of W t1 is 15.5575 dB.m2 , and that of W t2 is
Fig. 17 shows that the air inlet in the direction of the back opening could receive the down wash airflow well and the light blue wake trace outside the nozzle indicates that the exhaust gas flow still has a good velocity, further illustrating the better aerodynamic performance of the initial suppressor model. The static pressure on the top and back of the fuselage is large and that near the side ridge line of the fuselage and the front inlet side ridge is small. Different from Nl = 2, the suppressor model with N l = 1 closes the two second-stage inlets, removing the second-stage pipe and the rectangular lobe structure as shown in Fig. 18, where the airflow at the inlet of the first stage lobe has a temperature of 900 K and a velocity of 20 m/s negative along the x-axis. The static pressure on the upper surface of the pipe facing the incoming flow is large and gradually decreases. A low pressure zone appears in the pipe turning point and in the straight pipe section of the outlet. There is a high temperature area of about 840 K at the end of the nozzle face because the mixed airflow still has a high speed and reaches the nozzle first along the upper side of the pipeline, resulting in a shorter mixing time of the airflow. A large area of about 750 K on the outer side of the middle part of the nozzle for the reason that the down-washing action causes the partially mixed airflow to flow along the outside of the tube wall to the end of the nozzle. For the suppressor model with N l = 2, the nozzle temperature distribution has been greatly improved due to the ejecting effect of the two secondary air inlets on the down wash fluid, plus the larger size rectangular lobes re-blending the first-stage mixed airflow. The high temperature airflow is longitudinally dispersed by the rectangular lobes along the nozzle to ensure that each stream is further mixed and the temperature is not high and is no longer concentrated. Various M c models are presented in Fig. 19, where the straight sections of these models are the same in diameter and length. M c1 , M c5 and M c6 are circle lobed structure while both M c3 and M c4 are serrated crown structures. Different blenders affect the cooling effect of high temperature gas, which in turn affects the infrared radiation intensity of the system [40,41]. Fig. 20 indicates that different M c models have an apparent effect on the static temperature distribution of the nozzle face while the secondary ejector technique is used. Compared with the high
Fig. 17. Flow field characteristics in hovering state, down washing speed V d = 10 m/s, Nl = 1.
Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
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Fig. 18. The static pressure distribution of pipe and the total temperature of nozzle, V d = 10 m/s.
Fig. 19. Various sub models of the first stage blender.
Fig. 20. Nozzle static temperature distribution under various M c , V d = 10 m/s, Nl = 2. Table 3 Average temperature of nozzle end face under different W t values.
W t /mm T /K
1
2
3
4
250.455 522.5354
240.455 523.2303
260.455 510.2144
270.455 511.7061
temperature regions of M c1 , that of M c4 nozzle are significantly more because of its insufficient gas mixing. However, the suppressor model with M c6 obviously has a better cooling effect owing to the joint design of internal and external spoilers based on M c1 . Due to the presence of the external spoiler which is longer than the internal spoiler, some of the high-temperature gas could be mixed along the axially deeper area with the cold air at the back of the external spoiler, thereby forming more axes vortex, then the temperature of the first stage mixed gas stream is further reduced. Table 3 presents the mean temperature of the nozzle face with various W t values. As the W t value increases from 240.455 mm to 270.455 mm, the average temperature of the nozzle face decreases from 523.2303 K to 510.2144 K and then rises to 511.7061 K,
showing that the effect of W t setting on the nozzle temperature is obvious and a minimum value appears in the given range. W t value determines the inclination of the outer wall of the second-stage pipe, affecting the internal space of the secondary pipe and the movement of the airflow from the second stage inlets. Fig. 21 provides that the infrared radiation in the 3∼5 μm band is obviously higher than the other two bands, occupying the main position, and the radiation intensity in the 8∼14 μm band is nearly 30% higher than that of 1∼3 μm. Since the suppressor model adopts the radar/infrared stealth design, the radiation direction of the nozzle is greatly limited. Considering the shielding effect of the upper and lower side plate of the tail beam in the normal observation field, the radiation angle is in the range of −120◦ ∼0◦ and the corresponding scope of αn is −30◦ ∼90◦ . Fig. 22 shows that the infrared radiation using secondary ejector technology is significantly suppressed compared to the singlestage infrared suppressor. The maximum infrared radiation intensity of Nl = 1 is 567.0247 W/sr, while that of N 1 = 2 is less than
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Fig. 21. Comparison of infrared radiation intensity of nozzle face under different wavelength bands, hovering state, V d = 10 m/s, Nl = 1, normal observation field.
Fig. 22. Contrast of infrared radiation intensity of nozzle face in hovering state, V d = 10 m/s, 3∼5 μm wave band, normal observation field.
Fig. 23. Comparison of infrared radiation intensity of exhaust gas under different W h and W t , V d = 10 m/s, 3∼5 μm wave band, level observation field.
200 W/sr, because the combined design of the secondary air inlet, the secondary pipe and the rectangular lobes plays a significant role in further reducing the temperature of the primary mixed gas stream. The maximum radiation intensity of M c1 is 189.8461 W/sr, and that of M c5 and M c6 are 191.0054 W/sr and 186.8582 W/sr, respectively, because M c5 has no chamfer and spoiler, but M c6 has external spoilers design to better mix hot and cold air compared to M c1 . The design of M c affects the aerodynamic, blending and cooling of the airflow within the system. Fig. 23 supports that the infrared radiation distribution under various W h is obviously different, where the radiation intensity values at most azimuth angles of W h2 and W h3 are obviously larger than that of W h1 and W h4 . As the value of W h increases, the maximum value of radiation decreases first, then increases
and then decreases. The maximum value of the four curves differs by 20.0747 W/sr, indicating that the effect of W h setting on infrared radiation suppression is still quite large. Increasing W h allows the secondary fresh air to reach the front part of the spout faster, which is used to cool the high temperature mixed airflow of the first stage. Observing the four radiation distribution curves under W t parameters, the radiation distributions of W t3 and W t4 are similar and much higher than the other two. The maximum radiation intensity in the four curves is 85.3225 W/sr, and the difference between the maximum values is 17.2134 W/sr. A suitable W t value provides a good control of the infrared radiation level of the exhaust gas, because the W t value directly determines the inclination of the tail beam side plates and the internal space of the secondary
Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
Fig. 24. The history curve of the index
σ and the value T .
Fig. 25. The history curve of the index I n and the value I h .
pipe, which affects the total amount of secondary mixed gas and the flow of the airflow from the second secondary inlet into the secondary conduit. However, an excessive W t value will reduce the speed of the secondary airflow from the second secondary inlet, which is not conducive to the ejector effect. 4.3. Comprehensive results comparison Fig. 24 reveals that the RCS indicator fluctuates greatly at the beginning of the two phases of the design, with the maximum of 7.6194 dB.m2 and the minimum less than −5 dB.m2 , while the performance of this indicator is also changing in most of the subsequent phases, but M ∗ tends to be stable. During the whole optimization process, the fluctuations of T value is very obvious from the initial 592.4057 K to about 520 K. Although the T value of M ∗ is not always in a down state due to the COM and Pareto solution, it still has a good improvement. Fig. 25 manifests that the I n index has been in a steady downward trend. Although the I n performance of ordinary individuals fluctuates most in the 2nd and 3rd generations, that of the optimized individual M ∗ decreased from the beginning of more than 500 W/sr to 158.3156 W/sr because of the optimal selection of
11
COM, showing that the suppressor itself has good infrared radiation suppression effect. The indicator I h has also been in a good improvement, from 204.9310 W/sr to around 75 W/sr, with large fluctuations occurring in the 2nd, 3rd and 5th generations, implying that the system has excellent infrared stealth capability in horizontal observation fields. Fig. 26 indicates that the RCS curve of the optimization model M ∗ is located below that of the initial model m0 , and the RCS peak of the initial model has been greatly weakened after optimization, where m0 has a peak value of 30.1275 dB.m2 at azimuth of 98.25◦ , while that of M ∗ is only −3.5863 dB.m2 . In the azimuth range of 72◦ ∼139.5◦ , the RCS value of M ∗ is significantly smaller than that of the initial model. The high temperature region of m0 nozzle face is concentrated in the middle and the tail part, and the maximum temperature could exceed 800 K and exist in a large area. The nozzle face of M ∗ does not have a large area of high temperature, and the original hot zone is dispersed by the rectangular lobe M r . Only a very small area of M ∗ nozzle face has a temperature of 800 K, most of the hot and concentrated places have a temperature of about 630 K and the average temperature of the nozzle face is just 522.5 K, showing that the optimized suppressor model has obvious effect on exhaust gas cooling at nozzle face. Fig. 27 suggests that the high temperature zone of M ∗ is much reduced compared to m0 , including hot gas flow in the pipe, the spout and the large area below the spout in the central flow field. Due to the addition of two secondary air inlets, the upper part of the high temperature gas in the M ∗ pipe is gradually cooled, and the mixing of the narrow rectangular lobes further reduces the temperature of the secondary air flow. On the section y = 0.45 m and x = −9.5 m, the length and area of the hot exhaust gas are greatly reduced after optimization, and the exhaust gas outside the spout of M ∗ is almost no longer extremely high, showing a good cooling performance of the engine intake and exhaust system. After optimization, the M c , W h and W t of the suppressor model changes greatly as shown in Table 4. M ∗ finally adopts the secondary ejector technology, the first-stage blender uses M c6 , and the second-stage blender adopts M r1 , while the value of W h is reduced from 800.142 mm to 487.142 mm and that of W t is reduced from 510.455 mm to 250.455 mm, where the other designs remain unchanged. Table 5 provides that the σ value of M ∗ is reduced by 13.0740 dB.m2 compared with that of m0 , because of the overall radar stealth design and the optimization of M f , W h and W t , indicating that the optimized system has a good radar stealth characteristics in the level plane. The decrease of I n value is 374.8010 W/sr after optimization, and the I n value of the secondary suppressor is reduced by 20.1777 W/sr, which indicates that the suppression effect of the suppressor itself is greatly improved. The I h value is reduced by 129.5646 W/sr, showing that the infrared stealth performance of the system in the level observation field are significantly improved. The T value is reduced
Fig. 26. Comparison before and after optimization.
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Fig. 27. Performance comparison before and after optimization.
Acknowledgements
Table 4 Parameters or sub models comparison before and after optimization.
m0 M∗
Mb
Mf
Nl
Mc
W h /mm
W t /mm
Mr
D e /◦
M b0 M b0
Mf1 Mf1
1 2
M c1 M c6
800.142 487.142
510.455 250.455
M r0 M r1
−90 −90
This work was supported by the Excellence Foundation of BUAA for PhD and the National Natural Science Foundation of China (Grant No. 91641123). Appendix A
Table 5 Comparison of indicators before and after optimization.
m0 M∗
σ /dB.m2
I n /W/sr
I h /W/sr
T /K
7.6142 −5.4598
533.1166 158.3156
204.9310 75.3664
592.4057 522.5354
by 69.8703 K, implying that the average temperature of the nozzle face is well controlled. These results show that the proposed COM method could achieve better results for both the RCS reduction and the infrared radiation suppression of the engine intake and exhaust system.
Table A1 Grid size distribution of each part of the flow field. Area
Max size/mm
Area
Max size/mm
Part min size Second lobe Nozzle First lobe inlet First pipe Internal fluid field Rotor face Center fluid field Outer boundary surface
0.5 10 15 20 30 50 500 500 8000
First lobe First inlet Center body Second inlet Second pipe Fuselage Interface in Center fluid field Interface in External fluid field External fluid field
5 15 20 20 30 100 500 500 8000
5. Conclusions Aiming at the focus of radar/infrared integrated stealth of helicopter engine intake and exhaust system, a preliminary stealth suppressor model design is carried out, and a series of work of RCS reduction and infrared radiation suppression are carried out. The most important points are: (1) Compared with traditional non-radar stealth design and flow field construction, the initial model of the suppressor based on the secondary ejector and radar stealth is designed, and a reasonable flow field including internal, central, and external zone is established and a high-precision unstructured grid is provided. (2) According to the characteristics of the hovering state of the helicopter, the proposed CFD method is used to simulate the flow field characteristics of the system. The downwash flow is a good measure to suppress infrared radiation and the air inlets on the fuselage back are popular for radar stealth. (3) The radar scattering characteristics of the system are calculated by PO+PTD and the IR radiation intensity is determined by MC+RTM (ray tracking method, RTM). Reasonable M f , W h and W t are the main measures of low RCS design for the system, while the main factors affecting the IR characteristics are N l , M c , W h and W t parameters or sub models. (4) The optimization design scheme of the helicopter engine intake and exhaust system is obtained by the proposed approach and the radar/infrared integrated stealth performance of the system has been well improved. Declaration of competing interest The authors declare that there is no conflict of interest.
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Aerospace Science and Technology www.elsevier.com/locate/aescte
Radar/infrared integrated stealth optimization design of helicopter engine intake and exhaust system Zeyang Zhou ∗ , Jun Huang, Jinjun Wang School of Aeronautic Science and Engineering, Beihang University, Beijing 100191, China
a r t i c l e
i n f o
Article history: Received 15 July 2019 Received in revised form 8 October 2019 Accepted 13 October 2019 Available online 17 October 2019 Keywords: Radar cross section Infrared radiation Comprehensive stealth Helicopter Engine intake and exhaust system
a b s t r a c t Along with the diversification and sophisticated development of detection methods in modern warfare, helicopters are increasingly subject to unilateral or simultaneous threats from radar and infrared detectors. In order to improve the survivability and operational effectiveness of the helicopter, a comprehensive stealth approach based on Pareto solution is presented. Considering the geometric constraints and aerodynamic characteristics of the engine intake and exhaust system, the model of the system is established by the full factorial design, the internal, central and external flow fields are constructed, then the highprecision computational fluid dynamics method is used to simulate the total flow field under the rotor downwash airflow in hovering state. The radar cross section of the system is evaluated by the physical optics and physical theory of diffraction. Based on the Monte Carlo and ray tracking method, the infrared signature of the system is calculated and analyzed in detail. Under the comprehensive evaluation and selection of comprehensive stealth approach, the optimization model of the system is continuously established and updated. The ultimate design has achieved good results in both radar cross section reduction and infrared radiation suppression and the proposed method is effective and efficient for radar/infrared integrated stealth of helicopter engine intake and exhaust systems. © 2019 Elsevier Masson SAS. All rights reserved.
1. Introduction With the continuous replacement and joint application of infrared (IR) imagers and radar detection equipment in modern warfare [1], the survivability of helicopters has never been more challenging. The unilateral adoption of RCS reduction or IR suppression technology cannot guarantee the comprehensive stealth performance and integrated operational effectiveness of helicopters [2,3]. Therefore, it is very important to carry out research on helicopter radar/infrared integrated stealth. The radar stealth performance of aircraft is affected by many factors, such as surface configuration, tilt angle, azimuth and radar band [4,5]. In order to achieve low radar scattering of the target helicopter, physical optics (PO) and method of equivalent current are used to calculate radar cross section (RCS) of the target [4]. Some reduction schemes based on radar absorbing material have been proposed and used to improve the radar stealth performance of helicopter air intakes [6,7]. The singularity of physical diffraction theory has been studied and used to solve wedge diffraction, while the calculation results are compared with the triangular column
*
Corresponding author. E-mail address: [email protected] (Z. Zhou).
https://doi.org/10.1016/j.ast.2019.105483 1270-9638/© 2019 Elsevier Masson SAS. All rights reserved.
RCS and the data of method of moment [8,9]. The electromagnetic performance of the radome is analyzed by the method of physical optics [10]. Facing the continuous development of radar systems, the improvement of helicopter survivability is becoming more and more important, and radar stealth has become one of the significant tactical indicators of modern helicopters. Aircrafts or helicopters sometimes fly at low altitudes or shuttle between obstacles to avoid the radar system, but they will still be the target of infrared imaging equipment or heat-seeking weapons [11,12]. The suppression methods such as concealing, camouflage, and downwash are implemented or experimented to study the infrared characteristics of helicopters [13–15]. Multistage ejector nozzles are tested and investigated to learn the infrared signature suppression [16,17]. Based on the basic principle of heat source method of heat transfer, an approach for calculating the temperature field of helicopter is proposed [18]. Different nozzle shapes are tested to study the performance of the small turbojet engine and the characteristics of infrared radiation [19]. The reverse Monte Carlo (MC) method and ray tracing method are used to calculate the infrared characteristics of the ejector-mixed lobes infrared suppressor [20–22]. In modern local wars, infrared countermeasures are intensifying and infrared imaging is becoming more and more accurate, so the helicopter’s excellent infrared stealth capability is particularly important.
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Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
Fig. 1. Schematic of radar/infrared integrated stealth for engine intake and exhaust system radar.
However, the complexity and diversity of modern and future battlefields [23–26] have greatly reduced the viability of traditional helicopters [27–31]. The aircrafts with unilateral stealth performance will still be targeted when facing both radar and thermal imaging or more detection equipment [32–35]. In order to pursue a good exhaust cooling effect [23,24], the design and installation of the infrared suppressor will change its shape characteristics at the installation of the helicopter body [42,43], thereby affecting the electromagnetic scattering characteristics of the aircraft under certain azimuth angles [25,26]. The helicopter full stealth has to emphasize the radar stealth design of various components and systems in important or even omnidirectional angles [29,30], which will also affect the intake/exhaust and cooling effects of the infrared suppressor [31–34]. Therefore, this paper attempts to develop a radar/infrared integrated stealth study of helicopter infrared suppressors to explore the interaction between radar stealth and infrared stealth of the engine intake/exhaust system. It is increasingly important to enhance the integrated stealth capability of helicopters [44–47], especially radar and infrared comprehensive performance. This manuscript presents the optimization method in section 2. The models and calculation patterns are provided in section 3. After discussing the results, the paper is concluded in section 5. 2. Comprehensive stealth approach The addition of the infrared suppressor is mainly to improve the IR radiation characteristics of the engine high-temperature exhaust by ingesting the surrounding cold air including the rotor downwash flow and the airflow in front of the helicopter as shown in Fig. 1, but this also affects the radar stealth performance of the helicopter rear fuselage where α is the azimuth angle and β is the elevator angle. The integrated design of the infrared suppressor increases the size of original slim rear fuselage beam, especially the secondary ejector mixing device. In order to improve the infrared/radar integrated stealth performance of the engine intake and exhaust system, the multiobjective function is defined as follows:
min : f i (m) ,
i = 1, 2, 3
M = {m1 , m2 , ..., mn }
(1)
where m is the model of the system, f i (m) represents the i-th target performance, and M refers to the set of different m.
Each target performance could be described as follows:
⎧ ⎪ ⎨ f 1 (m) = σ f 2 (m) = I n ⎪ ⎩ f 3 (m) = I h
◦ 0 ≤ α ≤ 180◦ , β = 0◦ −30◦ ≤ αn ≤ 30◦ HH
(2)
where αn is the observation angle in the normal observation field, HH indicates that the radar wave is horizontally polarized [36], the radar wave frequency is 10 GHz, σ is the far field RCS mean of m in the horizontal observation field of 0◦ < α < 180◦ , I n is the mean of infrared radiation intensity of m in the normal observation field of −30◦ < αn < 30◦ , I h is the mean of infrared radiation intensity of m in the horizontal observation field of 0◦ < α < 180◦ . 2.1. Constraints The equality constraints could be expressed as follows:
⎧ ⎪ ⎨ R 0 ( M i ) − R 0 M j = 0 L0 (M i ) − L0 M j = 0 ⎪ ⎩ R 1 (M i ) − R 1 M j = 0 ⎧ ⎪ j =0 ⎨ M b ( M i ) − Mcb M L1 (M i ) − L1 M j = 0 ⎪ ⎩ V 0 (M i ) − V 0 M j = 0
&
∀i , j ∈ {1, 2, ..., k}
(3)
where R 0 is the inner radius of the first stage lobe inlet, L 0 is the length of the first stage lobe straight section, R 1 is to the outer radius of the first stage lobe, M b is the central body model of the first stage lobe entrance, L 1 is the first stage lobe length, V 0 is the gas velocity of the first stage lobe entrance and k is the optimized generations. The inequality constraints could be expressed as:
⎧ ⎪ ⎨ L ( M i ) − L max ≤ 0 H ( M i ) − H max ≤ 0 & ⎪ ⎩ W ( M i ) − W max ≤ 0 T M j +1 − T M j ≤ 0 ∀i ∈ {1, 2, ..., k} V M j +1 − V M j ≥ 0 ∀ j ∈ {1, 2, ..., k − 1}
(4)
where L, H and W respectively represent the length, height and width of m, L max , H max and W max are respectively the corresponding size limits, T refers to the average temperature of the nozzle end face and V is the average velocity of the nozzle end face exhaust airflow.
Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
3
2.2. Optimal scheme and Pareto solution When the optimal solution exists, it has the following form (COM test):
⎧ f i (M ) ≥ f i (M ∗ ) ⎪ ⎪ ⎨
T M ∗j +1 ≤ T M ∗j ⎪ ⎪ ⎩ V M j +1 − V M j ≥ 0
∀ i ∈ {1, 2, 3} ∀ j ∈ {1, 2, ..., k − 1}
(5)
where M ∗ is the optimal model and satisfies all equality and inequality constraints. When the multi-target performance cannot obtain the minimum value at the same time, the feasible solution (Pareto solution) has the following form:
⎧ f i (M ∗) ≤ f i (M ) ⎪ ⎪ ⎪
⎨ ∗ T M j +1 ≤ T M ∗j ⎪
⎪ ∗ ⎪ ⎩V M V M ∗j ≥ j +1
∃ i ∈ {1, 2, 3} ∃ j ∈ {1, 2, ..., k − 1}
Fig. 2. Validation of the PO+PTD algorithm on the RCS calculation model of the IR system.
E s (r ) =
S
(6)
×
This indicates that there is at least one value of i or j satisfy the constraints of the comprehensive stealth performance.
1
E (r ) =
j ωε · 4π
S
3 − k2 R 2 + j3kR − jkR e R × R × J s r 5 R
1 + jkR
+2 Js r H (r ) =
1 4π
R3
−1 − jkR R3
S
e − jkR dS
(7)
e − jkR R × J s r
S
dS
(8)
e − jkR
Js =
×e
, Zi
0
, Zd
nˆ r · E 0 kˆ − nˆ r · kˆ E 0
− jk −ˆr +kˆ ·r
I=
ˆ × Rˆ × J s r R
R
dS
(9)
(10)
where n is the unit normal vector of the outer normal direction of r at the surface of the scatterer, Z i is the illuminated area, and Z d is the dark area. When the incident wave is a plane wave:
(13)
nˆ r · kˆ Eˆ 0 e − j2k·r dS
(14)
S
σ=
4π
λ2
| I |2
(15)
For solution of the edge diffraction, the diffraction coefficient of PTD could be found in Ref. [30,37]. The validation of the PO+PTD algorithm for RCS calculation is given in Fig. 2, showing that the two RCS curves are generally consistent including trend, mean, peak size, and peak position. The average value of RCS in the 0-180◦ azimuth of the PO+PTD curve is 0.4145 dB.m2 smaller than that of FEKO (PO+MOM), which indicates that the RCS calculation result is reliable.
The solution of the hot and cold air mixing problem in the intake and exhaust system is the basis for the calculation of infrared radiation. According to the radiation transfer equation (RTE):
ds
+ (a + σ S ) I (r , s)
σ T 4 σS = an + π 4π
4π
2
E i r = | E 0 | e − jk·r
dS
Considering the characteristics of plane waves, there are:
dI (r , s)
2n × H
(12)
2.4. Infrared radiation calculation
Based on the assumptions of physical optics, there are:
dS
RCS value could be determined as:
where r is the coordinate vector of the field point, r is the coordinate vector of the source point, and R is the distance between the field point and the source point, ω is the electromagnetic wave angular frequency, is the dielectric permittivity. Then E (r ) could be transformed into the following form:
−k2 E s (r ) = j ωε · 4π
R
rˆ × rˆ × S
The far-field RCS index of the system could be calculated by PO+PTD (physical theory of diffraction, PTD) method. According to the Stratton-Chu integral equation, the electric and magnetic fields could be obtained as follows:
e
− jk R +kˆ ·r
Then the integral term could be recorded as the following form:
I=
2.3. RCS calculation method
−k2 j ωε · 2πη
ˆ × Rˆ × nˆ r · E 0 kˆ − nˆ r · kˆ E 0 × R
(11)
where k refers to the wave vector. Considering the approximate condition of PO, the electric field formula can be written as:
I r , s s, s d
(16)
0
where r is the position vector, s is the direction vector, s is the length along the path, a is the absorption coefficient, s is the scattering direction vector, σ S is the scattering coefficient, n is the refractive index, I is the radiation intensity, σ is the StefanBoltzmann constant, T is the local temperature, is the scattering phase function, is the solid angle.
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Assuming that the reaction in the combustion chamber is complete, the scattering of the medium is not considered and the refractive index is equal to 1, the RTE could be simplified to:
dL (r , s) ds
+ aL (r , s) = a
σT4 π
(17)
Calculation of thermal radiation using discrete coordinates model:
∇ · [L (r , s) s] + aL (r , s) = a
σT4 π
(18)
Based on the Monte Carlo and ray tracing method, the spectral radiance of the detector receiving point is as follows:
L σ = L 0σ τ1σ τ2σ · · · τnσ + L b1σ (1 − τ1σ ) τ2σ τ3σ · · · τnσ + · · ·
+ Lnbσ−1 1 − τ(n−1)σ τnσ + Lnbσ (1 − τnσ )
(19)
where L 0σ is spectral radiance of wall reverse rays, L bi σ is spectral radiance, τi σ is i-th layer spectral transmittance. The irradiance of the detector receiving point could be determined as:
E=
N NB
L nσ ,(i , j ) · cos θ j · j · σi · 100
(20)
i =1 j =1
where E is the radiant illumination, N B is the total number of wave bands, N is the total number of rays contributing to the measurement point, θ j is the angle between the center of the j-th solid angle and the surface normal of its measuring point, j is the j-th solid angle, σi is the width of the i-th wave band. Then the radiation intensity at the receiving point is:
Iσ = Eσ (R) · R2
(21)
where I is the radiation intensity and R is the linear distance between the helicopter and the detector. Due to the continuous heating of high temperature gas, the temperature of the exhaust pipe nozzle is high and its infrared radiation is prominent. For a given infrared band, the radiance and radiance are calculated as:
⎧ ⎪ ⎪ ⎪ ⎨ ⎪ ⎪ ⎪ ⎩
λ2
ε ε L λ = M λ1 ∼λ2 = π π
M λ dλ
(22)
λ1
I λ = L λ A cos θ
where λ is the wavelength, M λ1 ∼λ2 is the radiation emission degree of the black body between the bands λ1 ∼ λ2 , A is the area of the nozzle bin, M λ is the spectral radiation emission degree of the black body. Introducing the Planck formula:
Mλ =
c1
1
(23)
λ5 ec2 /(λT ) − 1
Then the radiation emittance of blackbody in λ1 ∼ λ2 band could be obtained as:
M λ1 ∼λ2 =
15σ T 4
c 2 /(λ 1T )
π4 c 2 /(λ2 T )
c2 =
hc k
,
basically consistent with the data points of the literature [25], including the start and end positions, the maximum value, the mean value, and the turning point where scale of the engine exhaust model is 1.5, indicating that the presented calculation method is credible and accurate. 2.5. Method flow chart Fig. 4 shows that the integrated stealth optimization design process of the whole system mainly includes three parts: initial model design and performance evaluation; flow field, RCS and infrared radiation calculation of the model population, comprehensive performance evaluation and update of optimal models; ultimate model determination, method termination and recycling. 3. Model design and calculation method Taking the RAH-66 stealth helicopter as the reference object [30,39], the model of the infrared/radar integrated stealth engine intake and exhaust system is designed in a 1:1 ratio as shown in Fig. 5, where M f represents the front air intake model of the system, M b is the central body model at the first stage inlet, M c is the circular lobe model and acts as the first stage blending structure, and M r refers to the rectangular lobe acts as a second-stage blending structure, and D e is the normal direction of the nozzle end face. Ambient air enters the engine compartment through the front air inlet, and the hot gas after combustion reaches the center body at the entrance of the first-stage lobes [38], and mixes with the cold air taken in by the first-stage air inlet to flow to the second-stage lobes. The cold air taken in from the two intake ports of the second stage is again mixed and discharged outward through the nozzle, thereby achieving the effect of suppressing infrared radiation of the exhaust gas. 3.1. Geometric model establishment
x3 ex − 1
dx
(24)
The radiation constants have the following relationship:
c 1 = 2π hc 2 ,
Fig. 3. Ref. [25] data to verify the presented IR signature calculation method.
σ=
2π 5 k 4 15c 2 h3
(25)
where h is the Planck constant, h = 6.626176 × 10−34 J·s, k = 1.38 × 10−23 J/K, c is the speed of light. The validation of the presented IR calculation method is shown in Fig. 3. The IR curve is
According to the design of the system, the air inlet is opened on the side of the fuselage, the rear half of the back and the upper surface of the tail beam as shown in Fig. 6. The exhaust port is opened on the side edge of the tail beam and the nozzle end face is parallel to the horizontal plane. The infrared suppressor is designed in the inner space of the fuselage. The model of this system could be represented by the following parameters or sub modules:
m = M b , M f , Nl , M c , W h , W t , M r , D e
(26)
Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
Fig. 4. Infrared/radar integrated stealth optimization method flow chart.
Fig. 5. Comprehensive stealth model design for the engine intake and exhaust system.
Fig. 6. The model of the infrared/radar integrated stealth system.
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Fig. 7. Flow field construction and meshing of the system.
Fig. 8. Ashok experiment results of the IR suppression device to verify the CFD method [17].
Fig. 9. Verification of grid number and convergence time, Nl = 1.
where Nl is the number of lobe stages, W h is the head width of the tail beam upper surface, and W t is the tail width of the tail beam upper surface. 3.2. Computational fluid dynamics method For the temperature distribution characteristics of the system in the hovering state, the whole flow field is constructed and divided into three parts as shown in Fig. 7. The external flow field is a large cylindrical as a whole and the top surface has a conical design. The central flow field is a flat cylindrical, including the paddle plane and the fuselage surface. The fuselage surface excluding the ducted tail rotor adopts 1:1 modeling and the paddle plane is used to simulate the down washing effect of the rotor. The internal flow field is the internal space of the infrared suppressor, including the first/secondary lobe, the center body, the inlets, the exhaust port and the first/secondary exhaust pipe. High-precision unstructured grid technology is used to divide each region, and the quality and quantity of the grid are strictly controlled to ensure the convergence of the calculation results of the computational fluid dynamics method. The grid size is set as shown in Table A1 of the Appendix A. The Ashok experiment uses a simple 5-stage round tube ejector as an example to study infrared radiation suppression [17]. The CFD method in this paper simulates the suppressor system as shown in Fig. 8, and the results of variation of mass entrainment with nozzle distance from bottom funnel agree well with the experimental data where H overlap / D nz = 4 and Re = 7073. Fig. 9 presents the verification of grid number and convergence time of the CFD method where the total number of grids increased from 3 million to 8.12 million at unequal intervals. The total number of iteration steps is 10,000 steps and is guaranteed to be completed within 24 hours or it is considered not to converge. The convergence time is replaced by the number of steps at which convergence begins. It could be seen that the temperature T and convergence steps begin to stabilize from the 4.91 million grid. In order to ensure the reliability of the CFD method and the speed of
Fig. 10. RCS evaluation method of the closed system. (For interpretation of the colors in the figure(s), the reader is referred to the web version of this article.)
the entire optimization process, the number of grids in this paper is controlled between 5.2 and 5.7 million, and the total number of iteration steps is 10,000 steps. 3.3. RCS evaluation method To reduce the number of grids in the RCS calculation, the closed model of the system is extracted as shown in Fig. 10, where the blue part maintains a certain gap with the corresponding fuselage skin to avoid direct heating of the skin. 3.4. Infrared radiation observation field setting According to the indexes of infrared radiation intensity, two observation fields are set as shown in Fig. 11. One of the observation fields is located in the horizontal plane and the other observation field is perpendicular to the x-axis where the nozzle end face is designed to be parallel to the horizontal plane. 4. Results and analysis The different front air intake models are shown in Fig. 12, including the inclined triangular air inlet (M f 1 ), the smooth convex
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Fig. 11. Infrared radiation observation field setting diagram.
Fig. 12. Different front inlet models for RCS calculation.
Fig. 13. RCS distribution of the system model under M f 1 .
curved air inlet (M f 2 ), the trapezoidal air inlet (M f 3 ), the curved air inlet with forward lip (M f 4 ) and the tilted polygonal air inlet (M f 5 ). The M f 1 is similar to the diamond inlet design and does not form a strong reflection on the forward radar wave. The RCS results of the system model with M f 1 are shown in Fig. 13. The RCS peak appears at azimuth 270◦ because of the plane reflection of the vertical left side plate, but this part surface is hidden inside the fuselage and does not contribute to the fuselage RCS. A large RCS peak occupies a size of 15.5447 dB.m2 in the lateral 92.45◦ , which has a significant effect on the radar scattering characteristics of the fuselage. 4.1. Radar stealth optimization Fig. 14 supports that the RCS peak of the M f 3 curve is the largest and the overall value is also large, which is mainly caused by its mediocre trapezoidal shape design. The peak of M f 2 is similar to that of M f 1 , but the RCS of M f 2 is significantly higher than that of M f 1 in the azimuth range of 110◦ ∼150◦ . In the azimuth range of 0◦ ∼20◦ and 110◦ ∼120◦ , the RCS distribution of M f 5 model is obviously higher than that of M f 1 . The RCS values of M f 4 are larger than that of M f 1 and M f 5 in the azimuth angles
Fig. 14. The system model RCS comparison between various M f .
Table 1 RCS indicators of the system model with various M f .
σ /dB.m2
Mf1
Mf2
Mf3
Mf4
Mf5
−5.4302
−3.4146
0.0856
−1.8767
−3.8909
of 0◦ ∼30◦ and 75◦ ∼160◦ , because both M f 1 and M f 5 use a tilted plane design and M f 4 is a regular surface configuration. Table 1 presents the RCS indicators of the system model with various M f . The initial design of the front air intake model has a large impact on the RCS index, and the maximum difference could reach 5.5158 dB.m2 . The M f 3 suppressor model harvests the largest RCS mean in the range of 0◦ to 180◦ azimuth of the horizontal field, up to 0.0856 dB.m2 because its trapezoidal air intake side panel produces more strong scattering sources. M f 1 performs best as a result of its stealth design and the weaker edge diffraction. Fig. 15 reveals that the overall RCS curves of W h1 and W h2 are not much different, and the RCS peaks are almost equal and the peak positions are consistent. Compared with the RCS curve of W h3 , that of W h4 shows more minimum values in the azimuth
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15.4374 dB.m2 . The value of W t determines the tail width and inclination of the upper surface of the tail beam, affecting the radar wave illumination area and the angle between each bin and the radar wave. When both W h and W t increase a lot, it is a conventional non-radar stealth design, which results in very strong specular scattering of the tail beam side plates. 4.2. Infrared radiation suppression
Fig. 15. The system model RCS comparison between various W h .
Table 2 RCS indicators of the system model with various W h .
Value/mm σ /dB.m2
W h1
W h2
W h3
W h4
467.142 −5.4302
457.142 −5.3983
477.142 −5.4778
487.142 −5.4598
Fig. 16. The system model RCS comparison between various W t .
angles of 0◦ ∼30◦ and 110◦ ∼150◦ . But overall, the four RCS curves have no substantial drastic reduction effect. Table 2 manifests that the value of W h2 is reduced by 10 mm compared with that of W h1 , while its σ value reaches −5.3983 dB.m2 . As the value of W h increases, the RCS index first decreases and then increases, and a minimum value of σ is obtained at 477.142 mm. The size of W h affects the inclination of the side panels of the tail beam and the upper side panels on the rear fuselage, while this side panels will form a strong surface scattering source and more edge diffraction, which in turn affects the lateral RCS distribution of the system model. Fig. 16 provides that the four RCS curves have basically the same trend, and the peak occurs at the azimuth angle of 94.50◦ , wherein the RCS peak of W t1 is 15.5575 dB.m2 , and that of W t2 is
Fig. 17 shows that the air inlet in the direction of the back opening could receive the down wash airflow well and the light blue wake trace outside the nozzle indicates that the exhaust gas flow still has a good velocity, further illustrating the better aerodynamic performance of the initial suppressor model. The static pressure on the top and back of the fuselage is large and that near the side ridge line of the fuselage and the front inlet side ridge is small. Different from Nl = 2, the suppressor model with N l = 1 closes the two second-stage inlets, removing the second-stage pipe and the rectangular lobe structure as shown in Fig. 18, where the airflow at the inlet of the first stage lobe has a temperature of 900 K and a velocity of 20 m/s negative along the x-axis. The static pressure on the upper surface of the pipe facing the incoming flow is large and gradually decreases. A low pressure zone appears in the pipe turning point and in the straight pipe section of the outlet. There is a high temperature area of about 840 K at the end of the nozzle face because the mixed airflow still has a high speed and reaches the nozzle first along the upper side of the pipeline, resulting in a shorter mixing time of the airflow. A large area of about 750 K on the outer side of the middle part of the nozzle for the reason that the down-washing action causes the partially mixed airflow to flow along the outside of the tube wall to the end of the nozzle. For the suppressor model with N l = 2, the nozzle temperature distribution has been greatly improved due to the ejecting effect of the two secondary air inlets on the down wash fluid, plus the larger size rectangular lobes re-blending the first-stage mixed airflow. The high temperature airflow is longitudinally dispersed by the rectangular lobes along the nozzle to ensure that each stream is further mixed and the temperature is not high and is no longer concentrated. Various M c models are presented in Fig. 19, where the straight sections of these models are the same in diameter and length. M c1 , M c5 and M c6 are circle lobed structure while both M c3 and M c4 are serrated crown structures. Different blenders affect the cooling effect of high temperature gas, which in turn affects the infrared radiation intensity of the system [40,41]. Fig. 20 indicates that different M c models have an apparent effect on the static temperature distribution of the nozzle face while the secondary ejector technique is used. Compared with the high
Fig. 17. Flow field characteristics in hovering state, down washing speed V d = 10 m/s, Nl = 1.
Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
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Fig. 18. The static pressure distribution of pipe and the total temperature of nozzle, V d = 10 m/s.
Fig. 19. Various sub models of the first stage blender.
Fig. 20. Nozzle static temperature distribution under various M c , V d = 10 m/s, Nl = 2. Table 3 Average temperature of nozzle end face under different W t values.
W t /mm T /K
1
2
3
4
250.455 522.5354
240.455 523.2303
260.455 510.2144
270.455 511.7061
temperature regions of M c1 , that of M c4 nozzle are significantly more because of its insufficient gas mixing. However, the suppressor model with M c6 obviously has a better cooling effect owing to the joint design of internal and external spoilers based on M c1 . Due to the presence of the external spoiler which is longer than the internal spoiler, some of the high-temperature gas could be mixed along the axially deeper area with the cold air at the back of the external spoiler, thereby forming more axes vortex, then the temperature of the first stage mixed gas stream is further reduced. Table 3 presents the mean temperature of the nozzle face with various W t values. As the W t value increases from 240.455 mm to 270.455 mm, the average temperature of the nozzle face decreases from 523.2303 K to 510.2144 K and then rises to 511.7061 K,
showing that the effect of W t setting on the nozzle temperature is obvious and a minimum value appears in the given range. W t value determines the inclination of the outer wall of the second-stage pipe, affecting the internal space of the secondary pipe and the movement of the airflow from the second stage inlets. Fig. 21 provides that the infrared radiation in the 3∼5 μm band is obviously higher than the other two bands, occupying the main position, and the radiation intensity in the 8∼14 μm band is nearly 30% higher than that of 1∼3 μm. Since the suppressor model adopts the radar/infrared stealth design, the radiation direction of the nozzle is greatly limited. Considering the shielding effect of the upper and lower side plate of the tail beam in the normal observation field, the radiation angle is in the range of −120◦ ∼0◦ and the corresponding scope of αn is −30◦ ∼90◦ . Fig. 22 shows that the infrared radiation using secondary ejector technology is significantly suppressed compared to the singlestage infrared suppressor. The maximum infrared radiation intensity of Nl = 1 is 567.0247 W/sr, while that of N 1 = 2 is less than
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Fig. 21. Comparison of infrared radiation intensity of nozzle face under different wavelength bands, hovering state, V d = 10 m/s, Nl = 1, normal observation field.
Fig. 22. Contrast of infrared radiation intensity of nozzle face in hovering state, V d = 10 m/s, 3∼5 μm wave band, normal observation field.
Fig. 23. Comparison of infrared radiation intensity of exhaust gas under different W h and W t , V d = 10 m/s, 3∼5 μm wave band, level observation field.
200 W/sr, because the combined design of the secondary air inlet, the secondary pipe and the rectangular lobes plays a significant role in further reducing the temperature of the primary mixed gas stream. The maximum radiation intensity of M c1 is 189.8461 W/sr, and that of M c5 and M c6 are 191.0054 W/sr and 186.8582 W/sr, respectively, because M c5 has no chamfer and spoiler, but M c6 has external spoilers design to better mix hot and cold air compared to M c1 . The design of M c affects the aerodynamic, blending and cooling of the airflow within the system. Fig. 23 supports that the infrared radiation distribution under various W h is obviously different, where the radiation intensity values at most azimuth angles of W h2 and W h3 are obviously larger than that of W h1 and W h4 . As the value of W h increases, the maximum value of radiation decreases first, then increases
and then decreases. The maximum value of the four curves differs by 20.0747 W/sr, indicating that the effect of W h setting on infrared radiation suppression is still quite large. Increasing W h allows the secondary fresh air to reach the front part of the spout faster, which is used to cool the high temperature mixed airflow of the first stage. Observing the four radiation distribution curves under W t parameters, the radiation distributions of W t3 and W t4 are similar and much higher than the other two. The maximum radiation intensity in the four curves is 85.3225 W/sr, and the difference between the maximum values is 17.2134 W/sr. A suitable W t value provides a good control of the infrared radiation level of the exhaust gas, because the W t value directly determines the inclination of the tail beam side plates and the internal space of the secondary
Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
Fig. 24. The history curve of the index
σ and the value T .
Fig. 25. The history curve of the index I n and the value I h .
pipe, which affects the total amount of secondary mixed gas and the flow of the airflow from the second secondary inlet into the secondary conduit. However, an excessive W t value will reduce the speed of the secondary airflow from the second secondary inlet, which is not conducive to the ejector effect. 4.3. Comprehensive results comparison Fig. 24 reveals that the RCS indicator fluctuates greatly at the beginning of the two phases of the design, with the maximum of 7.6194 dB.m2 and the minimum less than −5 dB.m2 , while the performance of this indicator is also changing in most of the subsequent phases, but M ∗ tends to be stable. During the whole optimization process, the fluctuations of T value is very obvious from the initial 592.4057 K to about 520 K. Although the T value of M ∗ is not always in a down state due to the COM and Pareto solution, it still has a good improvement. Fig. 25 manifests that the I n index has been in a steady downward trend. Although the I n performance of ordinary individuals fluctuates most in the 2nd and 3rd generations, that of the optimized individual M ∗ decreased from the beginning of more than 500 W/sr to 158.3156 W/sr because of the optimal selection of
11
COM, showing that the suppressor itself has good infrared radiation suppression effect. The indicator I h has also been in a good improvement, from 204.9310 W/sr to around 75 W/sr, with large fluctuations occurring in the 2nd, 3rd and 5th generations, implying that the system has excellent infrared stealth capability in horizontal observation fields. Fig. 26 indicates that the RCS curve of the optimization model M ∗ is located below that of the initial model m0 , and the RCS peak of the initial model has been greatly weakened after optimization, where m0 has a peak value of 30.1275 dB.m2 at azimuth of 98.25◦ , while that of M ∗ is only −3.5863 dB.m2 . In the azimuth range of 72◦ ∼139.5◦ , the RCS value of M ∗ is significantly smaller than that of the initial model. The high temperature region of m0 nozzle face is concentrated in the middle and the tail part, and the maximum temperature could exceed 800 K and exist in a large area. The nozzle face of M ∗ does not have a large area of high temperature, and the original hot zone is dispersed by the rectangular lobe M r . Only a very small area of M ∗ nozzle face has a temperature of 800 K, most of the hot and concentrated places have a temperature of about 630 K and the average temperature of the nozzle face is just 522.5 K, showing that the optimized suppressor model has obvious effect on exhaust gas cooling at nozzle face. Fig. 27 suggests that the high temperature zone of M ∗ is much reduced compared to m0 , including hot gas flow in the pipe, the spout and the large area below the spout in the central flow field. Due to the addition of two secondary air inlets, the upper part of the high temperature gas in the M ∗ pipe is gradually cooled, and the mixing of the narrow rectangular lobes further reduces the temperature of the secondary air flow. On the section y = 0.45 m and x = −9.5 m, the length and area of the hot exhaust gas are greatly reduced after optimization, and the exhaust gas outside the spout of M ∗ is almost no longer extremely high, showing a good cooling performance of the engine intake and exhaust system. After optimization, the M c , W h and W t of the suppressor model changes greatly as shown in Table 4. M ∗ finally adopts the secondary ejector technology, the first-stage blender uses M c6 , and the second-stage blender adopts M r1 , while the value of W h is reduced from 800.142 mm to 487.142 mm and that of W t is reduced from 510.455 mm to 250.455 mm, where the other designs remain unchanged. Table 5 provides that the σ value of M ∗ is reduced by 13.0740 dB.m2 compared with that of m0 , because of the overall radar stealth design and the optimization of M f , W h and W t , indicating that the optimized system has a good radar stealth characteristics in the level plane. The decrease of I n value is 374.8010 W/sr after optimization, and the I n value of the secondary suppressor is reduced by 20.1777 W/sr, which indicates that the suppression effect of the suppressor itself is greatly improved. The I h value is reduced by 129.5646 W/sr, showing that the infrared stealth performance of the system in the level observation field are significantly improved. The T value is reduced
Fig. 26. Comparison before and after optimization.
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Z. Zhou et al. / Aerospace Science and Technology 95 (2019) 105483
Fig. 27. Performance comparison before and after optimization.
Acknowledgements
Table 4 Parameters or sub models comparison before and after optimization.
m0 M∗
Mb
Mf
Nl
Mc
W h /mm
W t /mm
Mr
D e /◦
M b0 M b0
Mf1 Mf1
1 2
M c1 M c6
800.142 487.142
510.455 250.455
M r0 M r1
−90 −90
This work was supported by the Excellence Foundation of BUAA for PhD and the National Natural Science Foundation of China (Grant No. 91641123). Appendix A
Table 5 Comparison of indicators before and after optimization.
m0 M∗
σ /dB.m2
I n /W/sr
I h /W/sr
T /K
7.6142 −5.4598
533.1166 158.3156
204.9310 75.3664
592.4057 522.5354
by 69.8703 K, implying that the average temperature of the nozzle face is well controlled. These results show that the proposed COM method could achieve better results for both the RCS reduction and the infrared radiation suppression of the engine intake and exhaust system.
Table A1 Grid size distribution of each part of the flow field. Area
Max size/mm
Area
Max size/mm
Part min size Second lobe Nozzle First lobe inlet First pipe Internal fluid field Rotor face Center fluid field Outer boundary surface
0.5 10 15 20 30 50 500 500 8000
First lobe First inlet Center body Second inlet Second pipe Fuselage Interface in Center fluid field Interface in External fluid field External fluid field
5 15 20 20 30 100 500 500 8000
5. Conclusions Aiming at the focus of radar/infrared integrated stealth of helicopter engine intake and exhaust system, a preliminary stealth suppressor model design is carried out, and a series of work of RCS reduction and infrared radiation suppression are carried out. The most important points are: (1) Compared with traditional non-radar stealth design and flow field construction, the initial model of the suppressor based on the secondary ejector and radar stealth is designed, and a reasonable flow field including internal, central, and external zone is established and a high-precision unstructured grid is provided. (2) According to the characteristics of the hovering state of the helicopter, the proposed CFD method is used to simulate the flow field characteristics of the system. The downwash flow is a good measure to suppress infrared radiation and the air inlets on the fuselage back are popular for radar stealth. (3) The radar scattering characteristics of the system are calculated by PO+PTD and the IR radiation intensity is determined by MC+RTM (ray tracking method, RTM). Reasonable M f , W h and W t are the main measures of low RCS design for the system, while the main factors affecting the IR characteristics are N l , M c , W h and W t parameters or sub models. (4) The optimization design scheme of the helicopter engine intake and exhaust system is obtained by the proposed approach and the radar/infrared integrated stealth performance of the system has been well improved. Declaration of competing interest The authors declare that there is no conflict of interest.
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