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Special Issue "Flow Control and Drag Reduction"

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

Deadline for manuscript submissions: 31 July 2023 | Viewed by 5962

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Prof. Dr. Zhenbing Luo

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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

Published Papers (6 papers)

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Research

Open AccessArticle

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

John-Paul Mosele

Andreas Gross

and

John Slater

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

Viewed by 346

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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Open AccessArticle

Effective Distance for Vortex Generators in High Subsonic Flows

and

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

Viewed by 743

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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Open AccessArticle

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

and

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

Viewed by 564

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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Open AccessArticle

Internal Characteristics of Air-Supplied Plasma Synthetic Jet Actuator

and

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

Viewed by 546

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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Open AccessArticle

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

and

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

Viewed by 809

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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Open AccessArticle

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

and

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

Viewed by 801

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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Planned Papers

The below list represents only planned manuscripts. Some of these manuscripts have not been received by the Editorial Office yet. Papers submitted to MDPI journals are subject to peer-review.

Title: Improvement of Aerodynamic Performance of Wing-in-Ground Effect Vehicle Fuselage Using Blowing Flow Control Technique
Authors: Noor Azman Dollah; Mohd Supian Abu Bakar; Rosdzimin Abdul Rahman; Muhammad Zuhairi Mohd Aliashak; Gunasilan Manar; Azam Che Idris; Mohd Rashdan Saad
Affiliation: National Defence University of Malaysia
Abstract: Wing-in-Ground Effect (WIG) vehicle is a type of vehicle that flies in the vicinity of the ground to take advantage of the ground effect phenomenon. Typically, it is designed to glide over a level surface (usually over the sea) by making use of ground effect, the aerodynamic interaction between the moving wing and the surface below. Predominantly, most of the drag in flight is induced by the wings, however, the cross-section profile of the WIG fuselage has a greater impact on the fuselage pressure drag and longitudinal moment due to the presence of a stepped hull. This research advocated using active flow control in the form of the blowing method in a wind tunnel experiment to determine the aerodynamic coefficient of the WIG fuselage. The small scale of WIG fuselage specimen (ratio 1:100) was conducted at six ground clearances from 0.05 to 0.30 meter and three blowing velocity coefficients, ranging from 0.98 to 2.36. The results show that the use of blowing at all ground clearance and blowing velocity coefficients led to the reduction of drag compared to the baseline experiment. In term of the best overall performance, the best combination is achieved at ground clearance of 0.05 and blowing velocity coefficient of 2.29. The data from this study proven that the addition of flow control gives a major improvement of about 6% in drag reduction and 5% of lift enhancement to the aerodynamic performance of WIG vehicle.

Title: A Comparison of Passive and Passive-Active Flow Control in Reducing Shock Wave Boundary Layer Interactions On Hypersonic Flow
Authors: Ahmad Syahin Abu Talib; Zinnyrah Methal; Mohd Supian Abu Bakar; Rosdzimin Abdul Rahman; Gunasilan Manar; Mark Kenneth Quinn; Shan Zhong; Benzi John; Mohd Rashdan Rashdan Saad
Affiliation: National Defence University of Malaysia
Abstract: This study investigates the effect of hybrid flow control on shock-boundary layer interaction (SBLI) in high-speed flow aerodynamics. Shock waves occur when a scramjet propulsion system reaches hypersonic, or higher than the supersonic speed, resulting in discontinuities and high gradient regions due to the interaction with the boundary layers on the vehicle surface. In general, exact prediction of shock wave structure and interaction with the boundary layer in operating conditions plays an important role in the design of the protective structures. Consequently, untreated SBLI could result in disastrous aerodynamic high-speed event. Therefore, a hybrid flow control effect that combines micro-ramps as passive flow control and blowing micro-jets as active flow control has been developed to investigate the SBLI in high-speed flow aerodynamics. Thus, the design for passive flow control will consist of a micro-ramp height of 60% and 80% of the boundary layer thickness (δ) in connection with a blowing micro-jet active flow control output of 0.3 δ, 0.9 δ, 1.6 δ and 2.2 δ coefficient momentum. Experiments have been conducted to investigate hybrid flow control's ability to delay separation towards an impinged SBLI. It has been shown that using a micro-ramp with a blowing micro-jet in between and upstream of the MR60 at 1.6δ coefficient momentum can reduce negative pressure gradients and slightly slow down the separation flow, resulting in a maximum performance of 20% separation delay.

Title: Improving the Aerodynamic Performance of WIG Aircraft with a Micro-Vortex Generator (MVG) in Low-Speed Condition
Authors: Zinnyrah Methal; Ahmad Syahin Abu Talib; Mohd Supian Abu Bakar; Mohd Rosdzimin Abdul Rahman; Mohamad Syafiq Sulaiman; Mohd Rashdan Saad
Affiliation: National Defence University of Malaysia
Abstract: This present study has investigated the potential of passive flow control towards induced drag by using a micro-vortex generator (MVG) at a backward facing step (BFS) location. A WIG craft is a fast watercraft that resembles a dynamically stabilised ship and 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 will stabilise the WIG structure and significantly decrease drag on the wing surface. However, the existing design of WIG hull fuselage tends to induce more drag during flight, especially at the flow downstream of a BFS, which will cause inefficient fuel consumption over the distance travelled. MVG with ramp type was chosen and tested at various angles () and heights (h). The values of θ are 12°, 16°, and 24°, and h are 0.4δ, 0.6δ and 0.7δ where δ refers to the boundary layer height. The model was designed using CAD software and fabricated using a 3D printer. The 3D model was tested in a subsonic wind tunnel at 2 x Re 2 x within 1 to 10 m/s. According to this study, the optimum angle and height of MVG for reducing drag coefficient were 16° at 0.6 height. In comparison to an uncontrolled case, the drag coefficient decreases significantly with the presence of MVG.

Title: Numerical investigation on hypersonic flat-plate boundary layer transition subjected to bi-frequency synthetic jet
Authors: Xinyi Liu; Zhenbing Luo; Qiang Liu; Pan Cheng; Yan Zhou
Affiliation: College of Aerospace Science and Engineering, National University of Defense Technology
Abstract: Transition delaying is of great importance for the drag and heat flux reduction of hypersonic flight vehicles. The first mode within low frequency and the second mode within high frequency exist simultaneously during the transition of 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 analysis are used to compare the control effects of jet with different frequencies, amplitudes and positions. Linear stability analysis results show that, low frequency synthetic jet can suppress the first mode when it is arranged upstream of synchronization point, while the second mode control effect is relatively weak. The higher the high frequency is, the stronger the suppression effect is on the first mode. For the second mode, the suppression effect is only at f2=89.09kHz. The larger the amplitude, the weaker the promoting effect for the first mode and the second mode, and the more obvious the suppressing effect. For the cases with synthetic jet downstream of synchronization point, all levels of the three parameters promote the unstable mode. In terms of the growth rate with the spanwise wave number, the control effect of the same factor and level under different spanwise wave number is different. In order to obtain the optimal control effect on transition, the three factors and the arrangement position of the synthetic jet should be selected as follows: the position is arranged in the upstream, with f1 = 17.82kHz, f2 = 89.9kHz, a =0.007, 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.

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