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A special issue of Aerospace (ISSN 2226-4310). This special issue belongs to the section "Aeronautics".

Deadline for manuscript submissions: closed (31 July 2023) | Viewed by 20958

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Special Issue Editor

Prof. Dr. Zhenbing Luo

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Guest Editor

College of Aerospace Science and Technology, National University of Defense Technology, Deya Road, Kaifu District, Changsha 410073, China
Interests: aerodynamics; flight control; active flow control; synthetic jet; plasma synthetic jet; thermal management; icing and deicing control; air-breathing propulsion power

Special Issue Information

Dear Colleagues,

Drag reduction is an eternal and hot topic in the design of low- and high-speed aircraft as well as underwater vehicles in order to achieve the purpose of saving fuel, improving speed, and increasing range. The conventional method of reducing drag through shape optimization has met a development bottleneck, whereas the adoption of certain flow control measures to affect the flow around various shapes can improve its drag characteristics and even the stealthy performance of the aircraft. Flow control can be applied to delay/advance transition, inhibit/promote flow separation, enhance/weaken flow stability, increase shock wave control, etc., so as to achieve drag reduction, which has broad application prospects and research value. This Special Issue will include the following topics: flow control techniques, flow separation control, lift enhancement and drag reduction, flight control, laminar flow control, transition control, turbulence drag reduction, shock wave control, SWBLI control, and other applications to cause drag reduction.

Prof. Dr. Zhenbing Luo
Guest Editor

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Keywords

passive flow control
active flow control
flow separation control
lift enhancement and drag reduction
laminar flow control
transition control
turbulence drag reduction
shock wave control
SWBLI control

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Published Papers (10 papers)

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Research

16 pages, 1876 KiB

Open AccessArticle

Aircraft Wing Design for Extended Hybrid Laminar Flow Control

by Lennart Lobitz, Hendrik Traub, Mats Overbeck, Maximilian Bień, Sebastian Heimbs, Christian Hühne, Jens Friedrichs and Peter Horst

Aerospace 2023, 10(11), 938; https://doi.org/10.3390/aerospace10110938 - 2 Nov 2023

Cited by 3 | Viewed by 2255

Abstract

Laminar flow offers significant potential for increasing the energy-efficiency of future transport aircraft. The German Cluster of Excellence SE²A is developing a new approach for hybrid laminar flow control. The concept aims to maintain laminar flow up to 80% of the chord length by integrating suction panels at the rear part of the wing, which consist of a thin suction skin and a supporting core structure. This study examines effects of various suction panel configurations on wing mass and load transfer for an all-electric short-range aircraft. Suction panel material, as well as thickness and relative density of the suction panel core are modified in meaningful boundaries. Suction panels made from Ti6Al4V offer the most robust design resulting in a significant increase in wing mass. For the studied configurations, they represent up to 33.8% of the mass of the wingbox. In contrast, panels made from Nylon11CF or PU1000 do not significantly increase the wing mass. However, the use of these materials raises questions about their robustness under operational conditions. The results demonstrate that the choice of material strongly influences the load path within the wing structure. Ti6Al4V suction panels provide sufficient mechanical properties to significantly contribute to load transfer and buckling stiffness. Locally, the share of load transfer attributed to the suction panel exceeds 50%. In contrast, compliant materials such as Nylon11CF or PU1000 are inherently decoupled from load transfer. Unlike the thickness of the suction skin, the relative density of the core structure strongly affects the wrinkling stiffness. However, wrinkling failure did not appear critical for the examined suction panel configurations. In the present study, the mechanical properties of Ti6Al4V cannot fully be exploited. Therefore, compliant suction panels made from Nylon11CF are preferred in order to achieve a lightweight solution, provided that they meet operational requirements. Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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20 pages, 8118 KiB

Open AccessFeature PaperArticle

Numerical Investigation of Hypersonic Flat-Plate Boundary Layer Transition Subjected to Bi-Frequency Synthetic Jet

by Xinyi Liu, Zhenbing Luo, Qiang Liu, Pan Cheng and Yan Zhou

Aerospace 2023, 10(9), 766; https://doi.org/10.3390/aerospace10090766 - 29 Aug 2023

Viewed by 1428

Abstract

Transition delaying is of great importance for the drag and heat flux reduction of hypersonic flight vehicles. The first mode, with low frequency, and the second mode, with high frequency, exist simultaneously during the transition through the hypersonic boundary layer. This paper proposes a novel bi-frequency synthetic jet to suppress low- and high-frequency disturbances at the same time. Orthogonal table and variance analyses were used to compare the control effects of jets with different positions (USJ or DSJ), low frequencies (f₁), high frequencies (f₂), and amplitudes (a). Linear stability analysis results show that, in terms of the growth rate varying with the frequency of disturbance, an upstream synthetic jet (USJ) with a specific frequency and amplitude can hinder the growth of both the first and second modes, thereby delaying the transition. On the other hand, a downstream synthetic jet (DSJ), regardless of other parameters, increases flow instability and accelerates the transition, with higher frequencies and amplitudes resulting in greater growth rates for both modes. Low frequencies had a significant effect on the first mode, but a weak effect on the second mode, whereas high frequencies demonstrated a favorable impact on both the first and second modes. In terms of the growth rate varying with the spanwise wave number, the control rule of the same parameter under different spanwise wave numbers was different, resulting in a complex pattern. In order to obtain the optimal delay effect upon transition and improve the stability of the flow, the parameters of the bi-synthetic jet should be selected as follows: position it upstream, with f₁ = 3.56 kHz, f₂ = 89.9 kHz, a = 0.009, so that the maximum growth rate of the first mode is reduced by 9.06% and that of the second mode is reduced by 1.28% compared with the uncontrolled state, where flow field analysis revealed a weakening of the twin lattice structure of pressure pulsation. Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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12 pages, 3991 KiB

Open AccessArticle

Improving the Aerodynamic Performance of WIG Aircraft with a Micro-Vortex Generator (MVG) in Low-Speed Condition

by Zinnyrah Methal, Ahmad Syahin Abu Talib, Mohd Supian Abu Bakar, Mohd Rosdzimin Abdul Rahman, Mohamad Syafiq Sulaiman and Mohd Rashdan Saad

Aerospace 2023, 10(7), 617; https://doi.org/10.3390/aerospace10070617 - 5 Jul 2023

Cited by 1 | Viewed by 1783

Abstract

This present study investigated the potential of passive flow control to reduce induced drag by using a micro-vortex generator (MVG) at a backward-facing step (BFS) location. A wing-in-ground (WIG) craft is a fast watercraft that resembles a dynamically stabilised ship that can move or glide across the surface of water or land. Therefore, the wing of the WIG is designed to glide when in contact with water, which helps to decrease drag and enhance the lift of the overall vehicle. However, the existing design of the hull-fuselage of WIG tends to induce more drag during the flight, especially at a flow downstream of a BFS, which will cause inefficient fuel consumption over the distance travelled. MVG with the ramp type was chosen and tested at various angles (°) and heights (h). The angles (°) tested were 12°, 16°, and 24°, while the heights (h) tested were 0.4 δ, 0.6 δ, and 0.8 δ, where δ refers to the boundary layer height. The model was designed and fabricated using a 3D printer. The 3D model was tested in a subsonic wind tunnel at Re = 6.1 ×

10^{4}

m⁻¹ to 6.1 ×

10^{5}

m⁻¹ between 1 and 10 m/s. This study demonstrated that the most effective angle and height of MVG for reducing the drag coefficient were 16° and 0.6 δ, respectively. In comparison to an uncontrolled case, the drag coefficient decreased significantly by 38% compared to the baseline. Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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9 pages, 4160 KiB

Open AccessCommunication

Effect of Air Jet Vortex Generators on the Shock Wave Boundary Layer Interaction of Transonic Wing

by Tingkai Dai and Bo Zhang

Aerospace 2023, 10(6), 553; https://doi.org/10.3390/aerospace10060553 - 11 Jun 2023

Cited by 1 | Viewed by 1770

Abstract

The interaction between shock waves and turbulent boundary layers (SBLI) is a common phenomenon in transonic and supersonic aircraft wings. In this study, we simulated the SBLI of a classical NACA0012 wing at an angle of attack (AOA) of 1.4° and Mach number (Ma) of 0.78 using the open-source software OpenFOAM. Our results show that an air-jet vortex generator can effectively reduce the length of the separation zone and improve the lift coefficient of the airfoil. The vortex structure generated by the jet vortex generator significantly reduces the separation caused by SBLI. We conducted simulations with jet angles of 30°, 45°, and 60° and found that the larger the jet angle, the stronger the vortex and the greater the improvement in the lift coefficient. When the jet angle was 60°, the vortex structure generated by the jet vortex generator transformed the normal shock wave into a λ shock wave, resulting in a maximum increase in the lift coefficient of 2.35%. The simulations focused on exploring the effect of the jet angle and determined that that optimal jet parameters that effectively reduce SBLI damage and improve the lift coefficient of the airfoil. Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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29 pages, 24159 KiB

Open AccessArticle

Numerical Investigation of Asymmetric Mach 2.5 Turbulent Shock Wave Boundary Layer Interaction

by John-Paul Mosele, Andreas Gross and John Slater

Aerospace 2023, 10(5), 417; https://doi.org/10.3390/aerospace10050417 - 29 Apr 2023

Cited by 2 | Viewed by 1358

Abstract

Supersonic shock wave boundary layer interactions are common to inlet flows of supersonic and hypersonic vehicles. This paper reports on wall-resolved implicit large-eddy simulations of a canonical Mach 2.5 turbulent shock wave boundary layer interaction experiment at the NASA Glenn Research Center. The boundary layer upstream of the interaction was nominally axisymmetric and two-dimensional. A conical centerbody with a 16 deg half-angle and a maximum radius of

0.147 D

of the test section diameter was employed to generate a conical shock wave, where D is the test section diameter. Asymmetric (swept) interactions were obtained by displacing the shock generator away from the test section centerline. The present simulation is for a shock generator displacement of

D / 6

. Results from the asymmetric simulation are compared with results from an earlier simulation of a corresponding axisymmetric interaction. The experimental Reynolds number based on test section diameter was

R e_{D} = 4 \times 10^{6}

. For the simulations, the Reynolds number was lowered to

R e_{D} = 4 \times 10^{5}

to keep the computational expense of the simulations within limits. Compared to the axisymmetric interaction, the streamwise extent of the separation varies considerably in the azimuthal direction for the asymmetric interaction. The separation is strongest at the azimuthal location that is closest to the shock generator. The streamwise extent of the separated flow regions is noticeably reduced and substantial crossflow is observed between the locations that are closest and farthest from the shock generator. A Fourier analysis of the unsteady flow data indicates low-frequency content for the separated region that is closest to the shock generator. Away from this region, with increasing sweep angle and cross-flow, the low-frequency content is diminished. A proper orthogonal decomposition captures spanwise coherent structures for the more two-dimensional parts of the interaction. Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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12 pages, 4423 KiB

Open AccessArticle

Effective Distance for Vortex Generators in High Subsonic Flows

by Ping-Han Chung, Yi-Xuan Huang, Kung-Ming Chung, Chih-Yung Huang and Sergey Isaev

Aerospace 2023, 10(4), 369; https://doi.org/10.3390/aerospace10040369 - 12 Apr 2023

Cited by 2 | Viewed by 2411

Abstract

Vortex generators (VGs) are a passive method by which to alleviate boundary layer separation (BLS). The device-induced streamwise vortices propagate downstream. There is then lift-off from the surface and the vortex decays. The effectiveness of VGs depends on their geometrical configuration, spacing, and flow characteristics. In a high-speed flow regime, the VGs must be properly positioned upstream of the BLS region. Measurements using discrete pressure taps and pressure-sensitive paint (PSP) show that there is an increase in the upstream surface pressure and the downstream favorable pressure gradient. The effective distance for a flat plate in the presence of three VG configurations is determined, as is the height of the device (conventional and micro VGs). Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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18 pages, 11822 KiB

Open AccessArticle

Time-Series-Data Interpolation Applied to Boundary-Layer Profiles Measured on Different Flights

by Hidemi Takahashi, Mitsuru Kurita, Hidetoshi Iijima and Seigo Koga

Aerospace 2023, 10(4), 322; https://doi.org/10.3390/aerospace10040322 - 23 Mar 2023

Cited by 1 | Viewed by 1480

Abstract

Turbulent boundary-layer profiles on an aircraft surface were measured during flight by pitot rakes in an experiment at subsonic speeds. Because separate flights have different flight sequences in terms of time, it is not easy to compare boundary-layer profiles measured on different flights with the corresponding premised conditions directly. Using one flight as a reference, this paper proposes a method to find the closest flight condition for each time instance from data from other flights by calculating a residual norm in combinations of flight variables. The results show that the proposed method successfully finds the best matches of the time instances from the second flight with those of the first flight. In addition, applying the interpolation method using response surface methodology further improves the accuracy of evaluation in the flight range of Mach 0.4 to Mach 0.8. The total uncertainty level of the proposed interpolation method was found to be 5.7%. Although this level of uncertainty is expected to be reduced, the effectiveness of the proposed interpolation method was presented in conjunction with an evaluation of its applicability to determine the riblet effect in reducing skin-friction drag qualitatively. Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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17 pages, 7614 KiB

Open AccessArticle

Internal Characteristics of Air-Supplied Plasma Synthetic Jet Actuator

by Rubing Liu, Shenghui Xue, Wentao Wei, Qi Lin and Kun Tang

Aerospace 2023, 10(3), 223; https://doi.org/10.3390/aerospace10030223 - 25 Feb 2023

Cited by 1 | Viewed by 2112

Abstract

Conventional plasma synthetic jet actuators rely only on jet orifice for suction when functioning for long durations. A limited supplementary gas leads to jet velocity reduction and weakening of the flow control ability. Therefore, this study proposes an air-supplied actuator with a check valve externally connected to the cavity to improve its gas-supplying ability and jet performance. A quartz glass discharge chamber is developed to clarify the internal working mechanism of the air-supplied actuator. High-speed schlieren is employed to photograph the internal flow field of the discharge chamber. The results reveal that the inhalation airflow velocity of the jet orifice is doubled when the actuator is continuously working in the effective frequency band under the combined action of additional air supply from the check valve in the inhalation recovery stage. The gas pressure in the cavity is closer to the initial discharge state, discharge breakdown voltage is higher, discharge energy is stronger, and the process of gas expansion to generate a jet is less affected by the core defect of the heat source, thereby significantly increasing the jet velocity and saturation operating frequency of the actuator. The obtained results have important implications for the performance optimization of the air-supplied actuator. Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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26 pages, 11137 KiB

Open AccessArticle

Analytic Solution of Optimal Aspect Ratio of Bionic Transverse V-Groove for Drag Reduction Based on Vorticity Kinetics

by Zhiping Li, Long He, Yueren Zuo and Bo Meng

Aerospace 2022, 9(12), 749; https://doi.org/10.3390/aerospace9120749 - 24 Nov 2022

Cited by 7 | Viewed by 1964

Abstract

Previous studies have implied that the AR (aspect ratio) of the transverse groove significantly affects the stability of the boundary vortex within the groove and thus drives the variation in the drag-reduction rate. However, there is no theoretical model describing the relationship between the AR and the stability of the boundary vortex, resulting in difficulty in developing a forward method to obtain the optimum AR. In this paper, the velocity potential of the groove sidewalls to the boundary vortex is innovatively described by an image vortex model, thus establishing the relationship between the AR and the induced velocity. Secondly, the velocity profile of the migration flow is obtained by decomposing the total velocity inside the groove, by which the relationship between the AR and the migration velocity is established. Finally, the analytical solution of the optimal AR (

A R_{o p t} = 2.15

) is obtained based on the kinematic condition for boundary vortex stability, i.e., the induced velocity equals the migration velocity, and the forms of boundary vortex motion at other ARs are discussed. Furthermore, the stability of the boundary vortex at the optimal AR and the corresponding optimal drag-reduction rate are verified by the large eddy simulations method. At other ARs, the motion forms of the boundary vortex are characterized by “vortex shedding” and “vortex sloshing,” respectively, and the corresponding drag-reduction rates are smaller than those for vortex stability. Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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18 pages, 7281 KiB

Open AccessArticle

The Influence of Steady Air Jet on the Trailing-Edge Shock Loss in a Supersonic Compressor Cascade

by Yinxin Zhu, Zhenbing Luo, Wenqiang Peng, Qiang Liu, Yan Zhou, Wei Xie, Pan Cheng, Zhengxue Ma and Xuzhen Xie

Aerospace 2022, 9(11), 713; https://doi.org/10.3390/aerospace9110713 - 12 Nov 2022

Cited by 1 | Viewed by 1715

Abstract

To effectively reduce shock wave loss at the trailing edge of a supersonic cascade under high back-pressure, a shock wave control method based on air jets is proposed. The air jet was arranged on the pressure side of the blade in the upstream of the trailing-edge shock. The flow control mechanism and effects of parameters were analyzed by computational methods. The results show that the air jet formed an oblique shock wave in the cascade passage which decelerated and pressurized the airflow. The resulting expansion wave downstream of the jet slot weakened the strength of the trailing-edge shock. This could effectively change the normal shock into oblique shock and thus weaken the shock loss. Optimal control effect was achieved when the mass flow rate ratio of the jet to the passage airflow remained 0.35–1.11% and the distance from the jet slot to the shock foot of the trailing-edge shock was about five times the thickness of the boundary layer. The proposed method can reduce the total pressure loss of a supersonic cascade, with the maximum improvement effect reaching 7.29% compared to the no-control state. Full article

(This article belongs to the Special Issue Flow Control and Drag Reduction)

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