Hypersonic boundary layer transition control experiments should be sorted out from different perspectives. For example, from the perspective of measurement technology, it can be classified into qualitative measurement experimental research and quantitative experimental research. According to the path of transition development, it can be classified into natural transition experimental research and bypass transition experimental research. Based on the control intention, it can be divided into promoting transition experimental research and delaying transition experimental research [3, 50]. However, these methods are not conducive to explaining the advantages, energy consumption, cost and feasibility of control methods, and the selection of control methods for different transition problems.
As mentioned in Introduction, the transition control problem is a flow control problem. Since the 1990s, flow control has flourished and gradually developed into two branches, namely passive control and active control. By comparison, the research on the active flow control technology is more active and gradually becomes the forefront and mainstream in the past 10 years of the entire flow control field, especially plasma actuation, synthetic jet and other means [51,52,53].
On the contrary, due to the current development status of hypersonic vehicles and the technical complexity of active transition control, the research on hypersonic transition control mainly focuses on passive control methods [29]. Even since the twentieth century, the National Center for Hypersonic Laminar-Turbulent Transition Research in the United States has mainly investigated the passive transition control technology [5]. However, there are various active control methods emerging, while they are still in the embryonic stage.
This paper attempted to review the experimental studies on hypersonic boundary layer transition control, clarify the control mechanism of different transition control methods, guide the emerging methods and means with control advantages to optimize the layout of control devices (actuators), parameter selection, energy utilization and other issues, further develop hypersonic boundary layer transition control strategies, and provide technical support for aerodynamic optimal design.
3.1 Passive control
It is well known that there are lots of researches on passive control, and many experts and scholars have sorted out and analyzed them. Here, the typical results are just picked out to show the methods and the corresponding control mechanisms are pointed out.
Passive control methods mainly include rough element, cavity, corrugated wall, porous wall, micro channel, and so on [3]. In the following, four typical methods are introduced separately.
3.1.1 Rough element
Rough element has been investigated in depth and certain results were achieved in engineering application. There are many experiences that can be summarized.
The forced transition devices of X-43A, X-51A, Hyfly and other hypersonic vehicles were designed as rough element arrays, and corresponding experimental research has been carried out in the United States [54]. The forced transition devices of rough belt studied involve randomly distributed rough elements, discrete spheres, discrete diamond bodies and slope transition devices. The flight tests conducted by X-43A, X-51A and Hyfly are all ramp-type transition devices, mainly due to the high transition efficiency and low heat flow characteristics of the ramp-type transition devices. Figure 3(a) illustrates the rough belt installed on the X-43A. The slope rough belt successfully achieved a forced transition in the precursor during the flight test of the X-43A. Figure 3(b) demonstrates the forced transition device installed on the X-51A [54].
However, flight tests need the support of a lot of ground wind tunnel experimental research. In the wind tunnel experiments, Schneider et al. [30] took the lead in carrying out a large number of experimental studies and summarized three modes of rough element affecting transition, which provided an important reference for the development and analysis of the subsequent research on the influence law of parameters.
A great deal of subsequent research works [55,56,57,58,59] display that, among many parameters of rough element, the height of the rough element plays the most important role in the transition effects [60, 61], and evaluating whether the height can effectively promote transition (or lead to forced transition) by the local parameters of boundary layer has been recognized by academic circles. The corresponding empirical formula has also been summarized. The control effects of the rough element on the windward centerline can be quantified by Eq. (1), where k/δ represents the rough element height versus the local boundary layer thickness, and Reθ/Me refers to the boundary layer momentum thickness Reynolds number versus the outer edge Mach number. Later, a further research revealed that the empirical formula is also suitable for the rough elements in other positions on the windward side, and even suitable for cavities [62].
$$\left({Re}_\theta/M_e\right)\cdot\left(k/\delta\right)=\mathrm C$$
(1)
For most hypersonic vehicles (the first two types of vehicles mentioned in the second section of this article), the transition control measures are mainly to prevent the transition caused by rough elements. According to the wind tunnel data of the space shuttle, McGinley et al. [63] obtained that when k is the height of the surface roughness element, C = 27 is taken; if k is the depth of the cavity, C = 100 is taken; if the length of the cavity is taken, C = 900 is taken. These values indicate that the rough element cannot cause transition and does not affect the maximum size of the transition position. This provides a criterion for whether gaps, bulges and ablated walls can cause boundary layer transition.
From the demand of promoting transition, another problem that needs to be paid more attention to is the minimum size of transition caused by rough element and its influence on transition position.
In response, Van Driest [61] proposed the concepts of “critical” and “effective”, which solves the measurement problem for the boundary of the control effects of promoting transition demand, and provides a reference for parameter selection for rough element transition control.
In recent years, the relational mechanism was explained by combining control effects. Wheaton et al. [64] systematically studied the influence of different heights of cylindrical rough element in Boeing/AFOSR Mach 6 quiet wind tunnel of Purdue University. It was obvious that the change of the height of the rough element will inevitably lead to the change of the instability mechanism, but the root cause and influence law are still unclear.
Interestingly, even the effects of delaying transition appear with further research. In the Mach 6 quiet wind tunnel of Peking University, Tang et al. [65] captured the instantaneous velocity field and flow structure of the second mode unstable wave passing through the rough element for the first time. The results revealed that rough elements can not only be used to promote transition, but also play a role in delaying transition when the parameters are appropriate through refined flow field display. Here, it should be noted that, as early as the 1950s, Sterrett et al. [66] have discovered the phenomenon of rough elements delaying transition at Mach 6. Fujii et al. [67] and Fong et al. [68] also confirmed this point in subsequent studies, which attached the importance of parameter optimization and adjustment, and also showed the advantages and necessity of active control in hypersonic transition control research.
3.1.2 Cavity
After the crash of the U.S. space shuttle Columbia in 2003, the wind tunnel tests were carried out by Everhart J et al. [69, 70] on the space shuttle model and the flat plate model to study the influence of the cavity on the transition at Mach 6 and Mach 10.
Firstly, the flow physics of different cavity types was reviewed by Everhart J et al. The length-to-depth ratio L/H is typically used to distinguish and classify different cavity flow regimes, as depicted in Fig. 4. He specifically pointed out that transitional cavities (10 < L/H < 14) are avoided where possible in the present tests due to the complexity of the required instrumentation and test time necessary to address flow steadiness, which provides guidance for the subsequent experimental research.
The research work of NASA’s Langley Research Center revealed that the influence of the cavity on the transition depends on the length-depth ratio, the length-width ratio, and the ratio of depth to boundary layer thickness. The cavity with longer flow direction, wider span-wise and deeper depth is more likely to promote the downstream boundary layer transition.
Lawson and Barakos [31] also come to a similar conclusion with that of NASA’s Langley Research Center: In three-dimensional cavity flow, the flow characteristics depend not only on the length-depth ratio, but also the width-depth ratio of the cavity.
Ref. [71] pointed out that it is the downstream shear layer flow caused by the cavity that leads to boundary layer transition. Ohmichi and Suzuki [72] carried out wind tunnel experiments on the flow of a flat plate with a three-dimensional rectangular cavity under the condition of Ma = 7. It was discovered that the cavity induces flow vortex structure inside and outside the cavity, thus enhancing the heating rate of the inner wall and downstream region of the cavity and promoting the transition process. The conclusion is consistent with that of Ref. [71], and the internal physical mechanism is preliminarily revealed. The control mechanism can be described as that a shallow cavity generates longitudinal vortices inside and outside of the cavity and these vortices augment heating rates on the cavity floor, rear and side walls and the downstream region of the cavity. Figure 5 displays the experimental configuration and experimental results by Ohmichi and Suzuki. This figure first shows the schlieren results, and then the oil flow results of two configurations.
With the continuous development of flow field testing methods, the mechanism of using the cavity to control the transition of the hypersonic boundary layer has been further revealed. During the last 2 years, several hypersonic quiet wind tunnel tests were conducted on the flat-plate model with independent cavity on the surface under the condition of Ma = 6.5 by Shuo Chen [73]. Figure 6 illustrates the boundary layer structure in the presence or absence of a cavity.
In Fig. 6(a), when the airflow passes through the cavity, the flow begins to become unstable. It is determined to be the first mode wave by the wavelength of the disturbance wave. While, in Fig. 6(c), when the flow passes through 200 mm, the boundary layer still maintains a stable laminar flow state. The first mode wave structure begins to appear at a further downstream position.
Through PCB pressure measurement, Shuo Chen also found that when a cavity was present, the second mode wave did not appear, then the secondary mode appeared, and gradually the amplitude remained at the same level as the flat-plate model. This indicated that the cavity itself does not stimulate the secondary mode, which is consistent with Ref. [65]. Therefore, the cavity promotes the transition by increasing the growth of the first mode wave rather than the growth of the second mode wave [73].
3.1.3 Porous Wall
Porous wall is considered as a passive method with high practical prospect, and its main function is to delay transition, which is of great significance to drag reduction and thermal protection of hypersonic vehicles. The instability of the second mode disturbance wave is one of the main reasons and leads to the transition in hypersonic boundary layer. The second mode has acoustic mode characteristics, so the control mode capable of absorbing sound waves generally has the potential to inhibit the growth of the second mode. In 1998, Malmuth et al. [74] was the first to discover that a porous coating capable of absorbing high-frequency sound waves can effectively inhibit the growth of the second mode. Subsequently, as a potential passive control method to delay hypersonic boundary layer transition, this porous surface has attracted the attention of a large number of scholars.
Maslov [75] confirmed the attenuation effects of the porous wall on the disturbances with more in-depth analysis. In fact, the porous wall used by Maslov is an ultrasonically absorptive coating (UAC) with conventional porosity, and “UAC” reflects the nature of the acoustic disturbance generated by the porous wall. This method can delay the transition and has a certain relationship with the acoustic mode of the hypersonic boundary layer [10]. As shown in Fig. 7, the boundary layer can be regarded as an acoustic waveguide, and acoustic disturbances such as the second mode and its higher-order harmonics propagate forward between the wall and the sonic line.
Fedorov et al. [76,77,78,79,80] found that the use of ultrasonic absorbing materials can inhibit the development of unstable waves of the second mode and higher frequency, and the porous wall can inhibit the transition of hypersonic boundary layer, thus increasing the Reynolds number of boundary layer transition by 50%.
Chokani et al. [81] used bispectral analysis to compare the nonlinear interaction (subharmonic resonance) of the second mode in the boundary layer of conventional wall and porous surface, in which the harmonic resonance of the boundary layer of porous surface was almost completely eliminated.
Maslov et al. [82] also carried out comparative experiments by adding braided materials to the part of a sharp cone at Mach 6. The magnified sixty-fold image and distribution of porous materials on the model are shown in Fig. 8. The experimental results proved that this measure prolongs the laminar flow area by 15% – 66% or even longer. Based on their analysis, the porous surfaces absorb part of the energy increased by the second mode disturbance, thus weakening the development of the second mode disturbance. Therefore, the flow can still remain stable and the transition is delayed.
With the gradual deepening of the research on the influence of porous wall on boundary layer instability, the mechanism of delayed transition of porous wall also develops.
During the last 2 years, combining wind tunnel experiments and theoretical analysis, Zhu et al. [83,84,85] found that the permeable wall can suppress the near-wall disturbances by changing the spatial distribution of perturbations in the fundamental resonance, which disrupted the phase-locked relationship and prevented the growth of fundamental oblique waves. Meanwhile, the research also indicated that the permeable surface greatly reduced the aerodynamic heating and delayed the transition. According to the effects of permeable wall on aerodynamic heating, he inferred that the behavior of the two-dimensional structure on the permeable wall was correlated with the suppression of the growth of three-dimensional waves. The distribution and surface microstructure of the permeable material are displayed in Fig. 9, and Fig. 10 shows the effects of the permeable wall on aerodynamic heating.
The inherent control mechanism of delaying the transition can be summarized in this way. Acoustic disturbances in the hypersonic boundary layer cause violent movement of the internal air after entering the porous material. Under the action of viscous dissipation, part of the mechanical energy of the acoustic disturbance is converted into heat energy. In addition, when the acoustic disturbance in the flow passes, it will produce changes in compression and expansion. The temperature gradient between the adjacent compression zone and expansion zone could cause heat transfer from the high temperature part to the low temperature part. What’s more, part of the mechanical energy of the acoustic disturbance will also be converted into thermal energy. Therefore, under the combined action of viscous dissipation and heat conduction, the mechanical energy of the second mode wave is converted into heat energy, and the second mode unstable wave is suppressed, thereby delaying the transition of the boundary layer [33].
However, Fedorov A also discovered that, although the porous surface could suppress the instability of the second mode, it also brings the phenomenon of low-frequency disturbances [10], which may be a problem we expect to solve in active control.
3.1.4 Wavy wall
Recently, with the emergence of various new measurement techniques, the experimental research on the stability of hypersonic boundary layer is more in-depth. Fujii studied the control effects of the wavy wall surface on the hypersonic boundary layer transition of a sharp cone with a 5° half cone angle under the condition of incoming flow Ma = 7.1 [55]. By reasonably designing the wavelength of the corrugated rough band (about 2 times the height of the boundary layer), the boundary layer transition was effectively delayed. According to Fujii’s experience, the control of transition of wavy wall surface to hypersonic boundary layer depends on the total temperature of incoming flow and the wavelength of wavy rough band to a great extent. Bountin et al. [86] also conducted experimental research on the stability of wavy walls in Mach 6 boundary layer. The plate model with wavy surface and locations of ICP and ALTP sensors and its section are exhibited in Fig. 11. The experimental results showed that the channels between wavy walls are prone to slight flow separation and reattachment, with certain interference to the flow field outside the boundary layer. However, the second mode instability wave along the wavy wall is effectively suppressed, and the laminar flow in the boundary layer is prolonged. In terms of the control mechanism, Bountin [86] agreed that the control effects of delaying transition produced by the wavy wall mainly are corrected with the basic flow, rather than directly acting on the second mode. However, in the process of his experiments, new unstable waves are noticed on the wavy wall, and the specific reasons for them are still uncertain.
In 2019, Peking University made new progress in using wavy walls to delay the hypersonic transition. Si et al. [87] attempted to weaken the local heating spot of aerodynamic heating by controlling the strength of the second-mode instability. The results provided the information that the wavy-wall can suppress the second-mode instability to a certain degree and eliminate the local heating spot before the transition is completed. Figure 12 exhibits the model Si et al. used.
As shown in Fig. 13(a), the regular periodic structures are second-mode waves. In contrast, there is no significant second-mode structure in Fig. 13(b), which means that the second-mode wave is depressed by the wavy wall. For the wavy-wall case displayed in Fig. 14(b), the hot spots have been eliminated completely.
This simple passive control method has the potential to be used in hypersonic vehicles, but its effects at different Mach numbers and Reynolds numbers still need to be further verified.
To sum up, for the negative effects in the off-design state, almost all passive control methods cannot avoid it. Passive control methods have such limitations not only in the field of hypersonic boundary layer transition control, but also in the whole field of flow control, which also makes active transition control methods rise since the twenty-first century.
3.2 Active control
Although passive control can meet the demand to a certain extent and some achievements have been formed in engineering applications, it is difficult to achieve good results in wide speed range and working conditions due to the fixed installation position and geometric profile. In addition, it cannot be controlled in real time and most passive transition control devices often produce additional resistance and other problems.
Besides, researchers also realized that the energy injection cost paid by active flow control methods is worthwhile compared with the benefits obtained by the hypersonic transition control. Therefore, in recent years, some active flow control methods have begun to rise [88, 89].
At present, there are many methods that have achieved some exciting results or showed great potential for hypersonic flow control, including active blowing control, CO2 injection active control technology, wall cooling/heating, local plasma discharge, and so on [90,91,92].
3.2.1 Active blowing (suction) control
Since the beginning of the twenty-first century, the NASA Langley Research Center has used Hyper-X model to study the active blowing control method of hypersonic boundary layer control in Mach 6 and Mach 10 wind tunnels [93]. The research objects involve a variety of opening models and blowing and sucking layouts, and different hypersonic transition control results have been obtained. The Hyper-X precursor blowing (sucking) model is displayed in Fig. 15 [94].
In the study on the active boundary layer control devices, the rough elements for passive control of the boundary layer are replaced by various blowing module assemblies. However, the experience gained from a large number of passive transition device control experiments still needs to be followed. Berry et al. [94] designed 9 active control unit configurations, as shown in Fig. 15, corresponding to 4 blowing (suction) modes, namely straight groove blowing, straight groove belt suction blowing, sawtooth groove blowing and porous blowing. A total of 14 blowing configurations were screened, and the results revealed that all configurations were effective for generating transition starting position movement. When the blowing pressure ratio is 5, the sound velocity jet condition at the nozzle can just be ensured. In order to move the transition start position near the boundary layer control device, a pressure ratio of 40 or higher is required so as to provide effective control. The sawtooth configuration requires the minimum blowing pressure ratio to produce effective transition movement. The H4 configuration with a single row of larger holes is the best in the concept of circular holes. The experimental results suggested that the active boundary layer control method for hypersonic air-breathing vehicles is feasible. At the same time, a comparative study on the effects of active and passive control methods was carried out for the same controlled object under Ma10 condition. Figure 16 displays the comparison of typical results of shock wave system. Compared with the control results of passive forced transition device, the active blowing is accompanied by small jet shock wave and flow separation, and the single jet penetration height cannot be directly compared with the height of the passive control device. The research results suggested that magnifying the jet height by twice is more suitable to equate the active result with the passive result, which provides a better idea for the development of active control methods in the future.
In addition, the type of blowing gas will also affect the control effects. Pappas and Okuno [95] studied the wall blowing effects of air, helium and freon-12 on boundary layer transition on a cone model with a half cone angle of 7.5° at NASA’s Ames Research Center. The research results illustrated that, under the same mass flow condition, the low-density gas helium has the greatest influence on the transition, and the heavy-density gas freon-12 has the smallest influence. Increasing the mass flow makes the transition position move forward, but the transition does not develop to the jet region.
Demetriades et al. [96] performed a study on the influence of different gas blowing effects on the transition of a cone boundary layer with a half cone angle of 5° in the Ma6 wind tunnel, and the unstable second mode disturbance wave at the outer layer of the boundary layer was observed. With the increase of mass flow rate, the instability of boundary layer intensifies, which leads to the advance of boundary layer flow transition. In addition, they also found that even a small wall blowing flow can result in the amplification and development of unstable disturbance waves in the case of angle of attack, and finally induce boundary layer transition. Schneider [97, 98] investigated and summarized the research on the influence of wall blowing effects on boundary layer transition. It was found that the blowing effects usually induce the transition to advance. The larger the mass flow rate or the lighter the composition is, the greater the transition advance caused by the gas is, and the closer the blowing effect is to the front end of the model, the greater the impacts are, as shown in Fig. 17. Therefore, the jet that appears near the tip has a very large influence on the transition of the boundary layer, which also provides a reference for the arrangement of other active flow control actuators.
The above reviews mainly focus on wall blowing, while wall suction mainly has the effects of delaying transition. Some studies displayed that this method can enhance the stability of boundary layer mainly by reducing boundary layer thickness and eliminating or weakening boundary layer velocity [99].
For wall blowing and suction control, it is quite difficult to analyze the control mechanism of blowing and suction on the hypersonic boundary layer by experiment alone. Combined with the numerical simulation research, blowing and suction are mainly adjusted by modifying the velocity profile of the boundary layer, adjusting the shape of the boundary layer, and then adjusting the movement of the fluid in the boundary layer, which could achieve the effects of controlling the transition of the boundary layer [100].
Although this technology can achieve a relatively good control effect, it is difficult to control its impacts on the aircraft/model profile at high Mach numbers, and it is easy to produce shock waves and other disturbances that interfere with the control effects or affect the boundary layer flow.
3.2.2 CO2 injection active flow control technology
The results of Clarke and McChesney [101] illustrated that the relaxation effects of chemical reaction can damp the sound waves whose frequency is near the reciprocal of relaxation time, which opens a new door for suppressing/delaying hypersonic boundary layer transition.
In the wind tunnel test, Adam et al. [102] accidentally discovered that, when the test gas is CO2, the transition position of the boundary layer is greatly delayed. At the same time, as the concentration of incoming CO2 increases, the transition delay becomes more significant under certain high temperature/high enthalpy conditions.
The internal mechanism can be explained as follows. Since the transition of the hypersonic boundary layer is usually dominated by the second mode, and the second mode is an acoustic disturbance, the non-equilibrium effects on the flow can attenuate these acoustic disturbances [103].
Similarly, in the comparative study of blowing effects on different media gases in the boundary layer, researchers discovered the unique boundary layer transition control ability of CO2 gas based on the principle of chemical reaction.
Leyva et al. [104,105,106] used shock wave wind tunnel system to study the effects of injecting CO2 into hypersonic boundary layer on transition. The results demonstrated that when injecting CO2 into boundary layer, the Reynolds number of boundary layer transition is higher than that of other injected media. The mechanism is that the vibration mode excited by CO2 interacts with the acoustic mode in the boundary layer, and the carbon dioxide vibration relaxation absorbs the energy of the acoustic mode disturbance wave, which attenuates the amplitude of the second mode instability wave, thus delaying the boundary layer transition. Leyva et al. [105] also found that the greater the CO2 injection rate in the boundary layer is, the more obvious the delay effects are.
John D. Schmisseur [107] summarized and prospected the active control technology research for CO2 gas to suppress hypersonic boundary layer transition in 2015, and he pointed out that CO2 injection active flow control technology has the advantages of sufficient gas source, convenient access and adjustable flow rate.
Here, the novel findings of the numerical simulation work must be mentioned. When the temperature is high to a certain level, the three dissociation modes of CO2 can be excited, and the effects of the delay transition are better at this time [108]. Combined with the experimental research details, the mechanism of CO2 injection control can be further clarified.
The vibration and dissociation of CO2 molecules can absorb most of the energy near the second mode frequency, thus making the second mode wave grow slowly, and inhibiting the transition. Obviously, the higher the boundary layer temperature is, the better the control effects are. The threshold of temperature is the excitation temperature of CO2 vibration.
However, from an engineering point of view, it is difficult to apply this measure under hypersonic conditions. A series of issues such as the introduction of air source and the installation of injection devices need to be considered. Similarly, the injection of CO2 gas will also bring disturbances that we do not want to see.
3.2.3 Wall cooling/heating
In actual situations, the surface of the thermal protection system of hypersonic vehicle has uneven heat flow distribution or sometimes local ablation occurs, which has brought attention to the wall cooling/heating control methods.
Experimental studies in the last century have confirmed that by reducing the wall temperature, the heat dissipation rate of the boundary layer can be suppressed, thereby inhibiting the occurrence of transition [102].
Wall cooling [29] can stabilize T-S wave, but it will make the second mode tend to be unstable. Cooling leads to the thinning of boundary layer, which makes it more sensitive to roughness.
The thermal protection system of hypersonic aircraft is usually composed of materials with different properties, thus resulting in uneven heat flow distribution on the surface. Therefore, local wall heating or cooling may achieve better results.
Soudakov et al. [109] placed a local heating/cooling unit on the surface of a cone, and it was found that local cooling would delay the transition at Ma6, and local heating would make the transition happen earlier. Fedorov et al. [110] also maintained that local cooling can suppress the amplitude of the second mode and delay the occurrence of transition through wind tunnel test research, and local temperature distribution on the surface of the experimental model is shown in Fig. 18.
What’s more, the research made by Zhao et al. [111] revealed that wall cooling can delay boundary layer transition, but wall cooling has limited laminar flow control effects on the high-speed three-dimensional boundary layer, compared with two-dimensional boundary layer.
For the control mechanism, wall heating or cooling is to cause changes of the boundary layer’s thickness and the corresponding position of sound velocity line, which will lead to changes in velocity evolution of disturbance wave phase in boundary layer and migration of synchronous points, thus achieving the purpose of controlling transition.
However, this type of control method is difficult for experimental study, and requires high technical requirements for heat transfer and insulation in the control process, so it is also difficult for it to be used in engineering application.
3.2.4 Local plasma discharge
As early as the 1990s, the Institute of Theoretical and Applied Mechanics of the Siberian Branch of the Russian Academy of Sciences invented a method of glow discharge to control disturbance and study wave development in supersonic flow [112]. According to Kosinov et al. [113], the disturbance introduced by glow discharge into the mainstream has the essence of acoustic disturbance, which proved that glow discharge has the potential to induce transition.
In 2001, Maslov et al. [114] applied the same disturbance method to hypersonic speed flow, and it was discovered that Tollmien-Schlichting wave excited by sound wave at the leading edge of a flat plate can affect the receptivity of hypersonic boundary layer, which indicated that glow discharge is possible to be applied as an active control method in hypersonic boundary layer transition research.
Heitmann et al. [115] used a non-intrusive pulsed YAG double-pulse laser to generate glow discharge plasma on the surface of the model, and it was further converted into acoustic disturbance in the hypersonic boundary layer. The artificial disturbance further excites the second mode unstable wave in the hypersonic boundary layer. By comparing the disturbance growth amplitudes of natural transition, forced transition and linear theory, Heitmann et al. discovered that the disturbance growth rate of unstable waves obtained by experiments is not consistent with the theoretical prediction, which further reflected that plasma discharge can be used as an active controlled artificial disturbance to exert more flexible control on boundary layer transition.
Casper et al. [116] studied in detail the development process of pulsed glow discharge on the nozzle wall of Purdue University Ma6 quiet wind tunnel, and gave the change of pressure field. In Casper et al.’s experiments, controlled disturbances were created by pulsed-glow perturbations based on the electrical breakdown of air and a disturbance first grew into a second-mode instability wave packet that is concentrated near its own centerline. Weaker disturbances spread from the center. The waves grow and become nonlinear before breaking down to turbulence. The breakdown begins in the core of the packets where the wave amplitudes are the largest. The second-mode waves are still evident in front of and behind the breakdown point and propagate in the span-wise direction. The turbulent core grows downstream, thus resulting in a spot with a classical arrowhead shape. Behind the spot, a low-pressure calm region develops. These can lay a foundation for the subsequent exploration of internal mechanism.
Recently, in the experiments carried out by Lee Cunbiao et al. [117, 118] of Peking University, glow discharge was introduced as an artificial disturbance under the condition of Ma6. Firstly, an artificially introduced disturbance in the first-mode frequency range can excite a specific second-mode wave, one of the high-order harmonics of the added disturbance. A clear harmonic relationship between the first-mode and second-mode waves was found, and the phase lock phenomenon between them was found for the first time.
Secondly, 105 kHz glow discharge in the second-mode frequency range was introduced into the boundary layer of the flat plate, and the second mode wave was evidently excited. At the same time, the amplitude of the first mode has also been significantly enhanced, and the position of boundary layer transition has been significantly advanced, which proved that glow discharge has a broad prospect in the application of hypersonic boundary layer transition control, as a controllable active flow control method. Figure 19 displays the glow discharge image created and power supply and Fig. 20 displays the results of promoting transition.
Combined with the experimental results of various test methods, the internal mechanism can be explained. There are two reasons for making the transition happen earlier. On the one hand, the increase of the amplitude of the second mode significantly changes the effects of the average flow. On the other hand, the second mode instability interacts with the first mode via a phase-lock mechanism, which leads to the rapid amplification of the first mode [118].