Final Report Summary - STARLET (Basic Wind Tunnel Investigation to Explore the Use of Active Flow Control Technology for Aerodynamic Load Control) Executive Summary:The STARLET project „Basic wind tunnel investigation to explore the use of Active Flow Control technology for aerodynamic load control” was realised by Institute of Aviation, Warsaw, Poland, based on Clean Sky call for proposals no. JTI-CS-2011-01-SFWA-01-037.The duration of the project was 36 months, budget was 250 000 EURO with Clean Sky JU financing 75% of the costs. The main requirements of the Call for Proposal were to investigate the capabilities of flow control solutions, developed in the past for active load control, particularly for reduction of high off-design aerodynamic loads. It was envisaged that modifications of the past solutions would be necessary to adapt them for this particular purpose. Two locations on the wing chord were indicated for the flow control devices: the first one in the region where traditional spoiler is located and the second one in the vicinity of the trailing edge. The project goal was to investigate the effectiveness of the proposed by the applicant fluidic load control concept in low-speed wind tunnel investigations as well as in numerical flow control simulations. It was suggested also that a wing model with sweep would be beneficial for the investigations. The Institute of Aviation proposed conducting the investigations in the 5m-diameter low speed wind tunnel on a 2.4m half-span wing model with flow control devices located as recommended by the Topic manager. The project included literature study of the load control capabilities of the existing flow control solutions, design of the flow control devices to be implemented on the wing model, numerical flow simulations of the flow control concepts, fabrication of the flow control system and modification of the model and, finally, wind tunnel investigations and analysis of the results.Project Context and Objectives:Publishable summaryThe objective of the STARLET project was to check the usability of flow control actuation systems as developed in the past for load control purposes. The aim of the load control was reduction of high, off-design wing loads. Two locations of the actuation on wing profile were chosen: the first one was the region of spoiler, where significant values of lift occur in off-design conditions and the second one was the trailing-edge region strongly influencing the velocity circulation around the airfoil, where flow actuation could also reduce wing loads in off-design conditions. The project conditions required the development and low-speed wind tunnel testing of the fluidic actuators performing the load-alleviation function.The project started with a literature study in order to determine current state-of-the-art solutions and directions of investigations, to gather reference data for evaluation of the effectiveness of the concepts developed in the STARLET project and to define practical requirements and guidance for designing effective load control solutions. The main conclusions of the literature study were that a) producing flow separation on the suction side of the wing by blowing air through a system of nozzles is a viable strategy for significant reduction of lift in off-design load conditions; it is even likely that the natural pressure difference between the pressure and suction sides of the wing could be exploited for this goal, b) that despite the small thickness of the trailing edge successful concepts producing the Coanda effect in this region were in the past exploited for load control, and c) that for fluidic load control the very important guiding principle, and often hard to achieve is obtaining gradual aerodynamic effect in response to gradual control impulse. Apart from this, practical parameters for assessing the effectiveness and reference data was gathered.In order to meet the project objectives two concepts of fluidic load control were developed: the first one, designated as Fluidic Spoiler consisted of an array of nozzles in the central part of wing section emanating streams of compressed air and generating this way a large area of separated flow on the suction side of the wing. The nozzles were arranged in a chequered fashion: each nozzle was followed by a space in the span-wise as well as in the chordwise directions. The second concept, designated Double-Trailing-Edge Nozzle (DTEN) consisted of two, different-length, upward-directed nozzles in the modified trailing-edge that generated the Coanda effect and negative pressure on the lower side of the trailing edge and at the same time diverted the flow upwards. Both concepts exploited pressure chamber inside the wing model to feed the systems of nozzles with compressed air. The concepts were implemented on a moderate-sweep, 4.6-aspect-ratio 2.4 m. half-span wing model designed for investigations in the 5m-diameter low-speed wind-tunnel of Institute of Aviation.The design of the fluidic load control concepts was conducted with significant support from CFD flow simulations. In the initial phase two-dimensional (2D) simulations and two-and-a-half simulations (narrow wing strips of constant chord) were conducted in order to determine the most promising locations of the nozzles and blowing directions of the Fluidic Spoiler. The design and optimisation of the shape of the DTEN nozzles was also conducted at this stage of the project. The flow simulations indicated that the expected load alleviation could be obtained by blowing air in the direction normal to wing surface or in upstream direction, at an non-zero angle between the jet stream and wing surface. For practical reasons concerning the design of nozzles the blowing angles were investigated in the range from 30 from surface upstream to 30 from surface downstream. It was also concluded that the chequered arrangement of the nozzles achieves the expected design goals of producing flow separation while being at the same time technologically feasible. The results of flow simulations indicated at the same time that similar maximum values of load alleviation could be obtained from the Fluidic Spoiler and from the DTEN actuator. Based on the results of flow simulations the details of the Fluidic Spoiler and DTEN concepts were designed. The active part of the Fluidic Spoiler was the nozzle system. This system was designed as integral element, with the nozzles being the voids in the material, to be produced by three-dimensional printing technology, being at the same time the upper wall of the pressure chamber and fragment of upper (suction) surface of the wing model. Due to cost constraints this element was produced in two versions: one with nozzles normal to the surface and the second one with the nozzles deflected 45 from normal towards the flow. Both versions had nine rows of 60 nozzles, each nozzle 2,5mm by 5mm, wider side spanwise. As planned, the nozzles were arranged in a chequered fashion. The DTEN actuator was designed as a system of three-dimensional nozzles, provided with air from the same pressure chamber as the nozzles of the Fluidic Spoiler by pipes reaching the vicinity of the trailing edge. In this configuration the upper wall of the pressure chamber was solid. The nozzles of the DTEN actuator changed shape gradually from circular pipe to narrow, upward bent channels, acommodated in the trailing-edge region, separated by vertical and spanwise-oriented walls responsible for obtaining the Coanda effect.In order to complement the wind tunnel investigations of the fluidic load control concepts more detailed information about the effectiveness of the designed actuators and flow details was gathered from fully three-dimensional flow simulations.. The unsteady, Reynolds-averaged Navier-Stokes equations with k- turbulence model were solved in order to get knowledge about the effectiveness of the devices as well as of the steadiness of the flow pattern and aerodynamic loads. Due to constraints of computing time and large size of the computational problem only one blowing direction of the Fluidic Spoiler was simulated, but different numbers of active rows were also investigated. The results indicated that in blowing normal to wing surface the maximum effectiveness of load alleviation, measured by the reduction of wing-root bending moment was approximately 15% at mass flow of 0.270 kg/s and was obtained with two rows active. However, the flow pattern was strongly fluctuating and the noticeable effects of load alleviation occured at nozzle mass flow rate higher than 0.1 kg/s, whereas the configurations with larger number of active rows produced more steady flow pattern and became effective at very low non-zero values of nozzle flow. For all rows active the alleviation effect was lower (12%) at the same mass flow rate but the dependence of load alleviation level on nozzle mass flow rate was almost linear. For intermediate configuration (five active rows) the wing-root-moment alleviation level was 13% at the same mass flow rate but unsteady flow effects produced fluctuation of aerodynamic loads while increasing the nozzle mass flow rate. The results of flow simulations revealed also that the DTEN actuator was working in two modes: for lower values of nozzle mass flow rate, up to 0.1 kg/s the nozzles were acting similar to a Gurney tab, deflecting the flow but without producing the Coanda effect. Only after reaching this value of nozzle mass flow rate the Coanda effect appeared and upward deflection of flow behind the trailing edge was accompanied by appearance of a negative pressure region in the lower part of the trailing edge, supporting the load alleviation effect. The maximum wing-root alleviation level was 32% and this was achieved at nozzle mass flow rate of 0.15 kg/s. The side-effect of this concept was a nose-up increase of the pitching moment which was equal to Cm=0.068 at the maximum load alleviation effect. For the Fluidic spoiler the pitching moment changes were lower; depending on the location of the number and location of the active rows the pitching moment changes did not exceeded the value of Cm=0.001 positive or negative, depending on the chord position of the active nozzles. It must be noted, however, that changes of pitching moment occur also with the traditional load-alleviation systems, such as symmetrical deflections of ailerons on the Lockheed C-5 Galaxy aircraft. They have to be countered by compensatory deflection of elevators.Wind-tunnel tests of the designed fluidic load control systems were conducted in the 5m-diameter wind tunnel of the Institute of Aviation. The Mach number was M=0.1 and Reynolds number, based on the mean geometric chord was Re=2.4*106. For the Fluidic Spoiler the results of the wind tunnel tests qualitatively confirmed the results of numerical flow simulations conducted for nozzles directed normal to wing surface. The alleviation effect for all active nozzle rows active was more linearly dependent on the mass flow rate than was for lower number of active rows and for two rows active the noticeable effects of load alleviation occurred at nozzle mass flow rate higher than 0.1 kg/s, as in the flow simulations. Maximum values of wing-root bending moment reduction were, however, 30% higher than predicted by the flow simulations. A likely reason for this effect was absence in real flow, or lower intensity of small vortices that appeared near the nozzles in the numerical solution and produced local suction. The values of static pressure measured in the vicinity of the nozzles were higher than predicted in flow simulations. More information could be obtained from PIV scans, however, the PIV technique was not applied in these investigations. In contrast to numerical flow simulations that required large amount of computational time, in wind tunnel experiment larger number of nozzle configurations could be investigated. The experiment confirmed the conclusions of early two-dimensional and two-and-a-half dimensional flow simulations that deflecting blowing direction 45 upwind increases the load alleviation effect. Of the total number of 20 configurations differing in the number of active rows and blowing directions the most effective one was the configuration with first two nozzle rows active, deflected 45 toward the flow from direction normal to surface. The maximum wing-root alleviation level was 28% at 0.135 kg/s nozzle mass flow and at that blowing rate the saturation effect did not yet occur. This was, however, the maximum applied value of nozzle mass flow rate for this configuration due to concerns about the pressure rise in the pressure chamber inside the model and about the strength of the nozzle plate produced by the 3D-printing technology. The noticeable effects of this configuration on wing-root bending moment appeared at nozzle mass flow of 0.4 kg/s and after exceeding this value the dependence of wing-root bending moment alleviation on the nozzle mass flow was almost linear. For the DTEN actuator the wind-tunnel investigations confirmed the predictions of numerical simulations that at low values of the nozzle mass flow rate the Coanda effect does not appear or has low intensity, but the strong Coanda effect appeared at the nozzle mass flow rate of 0.055 kg/s which was similar mass flow rate to the rate necessary for activation of the most effective variant of the Fluidic Spoiler (generating flow separation on suction side of a wing). The maximum achieved wing-root bending moment alleviation effectiveness of the DTEN actuator was 31% which was slightly higher than for the most effective variant of the Fluidic Spoiler, but at this maximum effectiveness level the saturation effect was noticeable.It must be noted also, that both investigated concepts had higher load-alleviation capability than traditional spoiler of 10% wing section chord. For the classic spoiler located in place of the Fluidic Spoiler the wing-root bending moment alleviation level was 21% and this was achieved at a high deflection of 35 degrees. In practical situations achieving such high deflection requires large actuation moments produced by the hydraulic system and generates large increase of drag. For fluidic Spoiler the drag increase is moderate and results from the decreased suction of the leading edge due to decreased velociity circulation. In case of the DTEN actuator there occurs another component of drag – the suction on the rounded trailing edge which acts in the rearward direction. For this reason the drag increase due to activation of the DTEN device is higher than due to deflection of aileron, but lower than due to deflection of spoiler.Complementary actionsIn order to provide more possibilities of further development of the proposed concepts some complementary actions were taken, extending beyond the original scope of the project. Positive verification of the load-alleviation capability of the Fluidic Spoiler in wind tunnel tests led to the development of its simplified version, named the “Leaky Wing” concept, and to simulations of the load-alleviation effectiveness of the developed concepts in gust conditions. The “Leaky Wing” concept doesn’t rely on an independent source of high-pressure air but exploits the difference between the static pressure on the pressure and suction sides of the wing. This concept, investigated in numerical flow simulations, consists of slots connecting the pressure and suction sides of the wing. The shape and number of the channels require further optimization, but the results obtained for the initial version consisting of five blocks of slots located within external 40% of wing span, each block consisting of five slots in the central 10% of wing cross section proved capable of producing similar effects in terms of the area of separated flow to the effects of the Fluidic Spoiler fed with independent source of compressed air. Simulations of load-alleviation capability in gust conditions were conducted also for the DTEN and Fluidic Spoiler concept. The common advantage of the investigated fluidic load control concepts over the traditional spoiler in load alleviation is their rapid activation. Significant load-alleviation effects appear in the moment of activation of the nozzle flow whereas similar level of load-alleviation achieved using deflection of classical spoiler or symmetric deflection of ailerons requires large values of the deflection of these surfaces against the dynamic pressure which may require more time.ConclusionsAs a conclusion it may be stated that the results of the conducted works have proven the load alleviation effectiveness of the developed concepts. The load alleviation effectiveness of Fluidic Spoiler was higher than the effectiveness of a classic spoiler of the size of the actuating plate of Fluidic Spoiler. It must be noted, however, that the presented solutions of fluidic load control are still at very low technology readiness level (TRL2 or TRL3). More research is needed, concentrated on shape, dimensions, positions and characteristics at higher Mach numbers including possibilities of integratiion with wing structure and aeroelastic properties of wing equipped with such systems. Continuation of this research may lead to achieving technologically and economically effective solutions.Project Results:Main Science/Technology results and foregroundsThe main scientific and technological results of the STARLET project consist in the conducting of a proof of concept of fluidic load control approach for reduction or redistribution of high, off-design wing loads. In effect two alternative variants of fluidic load control systems were designed, investigated by numerical flow simulations and wind-tunnel tests (with a third sub-variant investigated solely by flow simulations) that have proven the load-alleviation effectiveness of the fluidic load control approach. The alternative variants of the investigated fluidic load control solutions were designated as Fluidic Spoiler, Double-Trailing Edge Nozzle (DTEN) actuator and a Leaky Wing which is a sub-variant of Fluidic Spoiler.The concepts were applied on the external part of the half-span wing model in order to obtain high effectiveness of reduction of the wing-root bending moment. The half-span wing model dimensions were: span: 2400mm, root chord: 1450 mm, tip chord: 665mm, reference area: 2.49m2,aspect ratio 4.6. Both actuators were located at the same spanwise location, between 59% and 92% wing half-span. Both actuators had also the same span, equal to 725mm. The detailed description of the actuators is provided in the following paragraphs. The concepts are shown in detail in drawings in the deliverable D4.2 “Final Technical Report” of the project.1. Research methodologyThe research methodology consisted in designing the fluidic load control actuators, implementing them on a half-span wind-tynnel wing model and conducting numerical flow simulations and wind tunnel investigations to assess the wing load alleviation effectiveness of the concepts. In numerical flow simulations the unsteady, Reynolds-averaged Navier-Stokes equations with k- turbulence model were solved in order to get knowledge about the effectiveness of the devices as well as of the steadiness of the flow pattern and aerodynamic loads with fluidic alleviation system active. In order to increase the fidelity of the solution internal flow in the nozzles of the actuators was also resolved using mass flow boundary condition at the internal end of a nozzle (inlet from pressure chamber inside the model). The boundary conditions used in the external boundaries of the domain were pressure far-field, pressure outlet and symmetry. Most of the simulations were three-dimensional flow simulations around the actual wing model. In initial stages 2.5D (constant-chord wing strip) and 2D flow simulations were also conducted in support of the design of the wing model. This way the most promising values of blowing angle of the Fluidic Spoiler with respect to airfoil chord were determined in the early stage of the project. The load-alleviation effectiveness of fluidic devices was determined in numerical and in experimental investigations by the values of the two parameters, namely: wing-root-bending-moment-alleviation coefficient (CBMA = [MB0-MB]/ MB0), that is the final effect of action of the fluidic devices and the value of mass flow rate of the air required to achieve this effect. More details of the redistribution of aerodynamic load on the wing were also obtained from the spanwise distributions of the bending moment, measured by tensometer system on six spanwise stations and from the local values CBMA-local for which the local values of bending moment and reference bending moment at given spanwise station were used. The distributions of local values of CBMA parameter obtained from wind-tunnel measurements and from numerical flow simulations were later compared in order to validate the computational models of the developed flow control actuators. Another parameter used in the literature for comparison of the effectiveness of fluidic devices is blowing momentum coefficient C_mi, C_mi=mfr*V_jet/(q_inf*S), where mfr is nozzle mass flow rate, V_jet is the velocity of the jet leaving the nozzle, q_inf is free-stream dynamic pressure and S is reference surface. This parameter is proportional to power needed to supply the system with mass flow. In numerical flow simulations both parameters, mass flow rate and C_mi were computed and the achieved results were presented against both of these parameters. In wind tunnel investigations only mass flow rate was measured. The C_mi used for comparisons of the efficiency of the devices was computed by flow simulations. This way both directions of investigations – numerical and experimental were complementary.Large proportion of the effort was devoted to the modification of the existing half-span wing model in order to accommodate inside it the elements of fluidic load control devices. The internal wing-box structure was dismantled and made more elastic by removing some material from it. More space was obtained in the half-chord region in order to accommodate the pressure chamber of the Fluidic Spoiler. The inside of the frontal part was occupied by pipes providing air to the Fluidic Spoiler. The aileron was dismantled and replaced with the DTEN actuator. A tensometer system was applied on wing upper side in order to measure the distribution of the bending moment. In addition, the wing-root bending moment was measured by two balances located at the wing spars. As a reference solution to compare the effects of fluidic devices a classic spoiler of the size of the Fluidic Spoiler actuating plate was used. In the case of classic spoiler its effectiveness can be determined only by the values of the root-bending-moment-alleviation coefficient. Both numerical and wind-tunnel investigations were conducted for the same conditions of angleof attack of 10 degrees, Mach number of 0.1 Reynolds number of 2.4 million. Numerical investigations of the Fluidic Spoiler were restricted to one blowing direction due to constrains in computational resources, but several variants of the device were analyzed, differing in the number of active nozzles, with the same total mass flow in the nozzle system. Wind tunnel investigations of the concepts were conducted in the 5m-diameter wind tunnel of Institute of Aviation, Warsaw. Wind tunnel investigations of Fluidic Spoiler were conducted for two angles of blowing with respect to surface: normal direction and 45 degrees upwind. The DTEN actuator had only one possible blowing direction.2. Description of Fluidic Spoiler and achieved results for this conceptThe Fluidic Spoiler actuator consists of an array of mini-nozzles that blow air from a pressure chamber inside the wing model in the direction normal to the suction side of the wing or in direction inclined at a non-zero angle from the normal to the wing surface in the upstream direction. The actuators investigated in wind-tunnel experiments had two variants of inclination: 0 (blowing in the direction normal to the surface) and 45 in the upstream direction. These inclination angles were chosen based on numerical simulations of the effectiveness of the blowing nozzles. The most important elements of the Fluidic Spoiler are pressure chamber inside the wing and the exchangeable pressure-chamber roof plates in which the nozzles are implemented. The nozzles are arranged in 9 spanwise rows in a chequered fashion so as to avoid leaking of air in the spaces between the nozzles. Each of the nozzles had 5.6mm span and 1.0mm width. Each nozzle row consists of 60 nozzles. The three-dimensional view on the fluidic spoiler generated by the PARADES optimisation system developed by Instytut Lotnictwa. The nozzles were implemented in the roof plate using three-dimensional printing technology which allowed designing the external side of the roof plate as a part of the suction side of the wing. In the experiments conducted in the Institute of Aviation the nozzle system was provided with high-pressure air from the high-pressure tanks used as second-flow system in the low-speed wind tunnel and the high-pressure source for another, transonic wind tunnel. If such a fluidic flow control system is to be implemented on a real-scale aircraft a suitable source of high-pressure air has to be implemented onboard.The results of numerical flow simulations indicated that in blowing normal to wing surface the maximum effectiveness of load alleviation, measured by the reduction of wing-root bending moment was approximately 15% at mass flow of 0.270 kg/s and was obtained with two rows active. However, the flow pattern was strongly fluctuating and the noticeable effects of load alleviation occured at nozzle mass flow rate higher than 0.1 kg/s, whereas the configurations with larger number of active rows produced more steady flow pattern and became effective at very low non-zero values of nozzle flow. For all rows active the alleviation effect was lower (12%) at the same mass flow rate but the dependence of load alleviation level on nozzle mass flow rate was almost linear. For intermediate configuration (five active rows) the wing-root-moment alleviation level was 13% at the same mass flow rate but unsteady flow effects produced fluctuation of aerodynamic loads while increasing the nozzle mass flow rate.For the Fluidic Spoiler the results of the wind tunnel tests qualitatively confirmed the results of numerical flow simulations conducted for nozzles directed normal to wing surface. The alleviation effect for all active nozzle rows active was more linearly dependent on the mass flow rate than was for lower number of active rows and for two rows active the noticeable effects of load alleviation occurred at nozzle mass flow rate higher than 0.1 kg/s, as in the flow simulations. Maximum values of wing-root bending moment reduction were, however, 30% higher than predicted by the flow simulations. A likely reason for this effect was absence in real flow, or lower intensity of small vortices that appeared near the nozzles in the numerical solution and produced local suction. The values of static pressure measured in the vicinity of the nozzles were higher than predicted in flow simulations. More information could be obtained from PIV scans, however, the PIV technique was not applied in these investigations. In contrast to numerical flow simulations that required large amount of computational time, in wind tunnel experiment larger number of nozzle configurations could be investigated. The experiment confirmed the conclusions of early two-dimensional and two-and-a-half dimensional flow simulations that deflecting blowing direction 45 upwind increases the load alleviation effect. Of the total number of 20 configurations differing in the number of active rows and blowing directions the most effective one was the configuration with first two nozzle rows active, deflected 45 toward the flow from direction normal to surface. The maximum wing-root alleviation level was 28% at 0.135 kg/s nozzle mass flow and at that blowing rate the saturation effect did not yet occur. This was, however, the maximum applied value of nozzle mass flow rate for this configuration due to concerns about the pressure rise in the pressure chamber inside the model and about the strength of the nozzle plate produced by the 3D-printing technology. The noticeable effects of this configuration on wing-root bending moment appeared at nozzle mass flow of 0.4 kg/s and after exceeding this value the dependence of wing-root bending moment alleviation on the nozzle mass flow was almost linear.3. The DTEN conceptThe DTEN concept consist of an array of nozzles located in the trailing edge of the external part of the wing at the same span location where the Fluidic Spoiler is located. The DTEN actuator is supplied with air from the same pressure chamber from which the Fluidic Spoiler is supplied. Each nozzle section is connected by a pipe with the pressure chamber, which for the investigations of the DTEN concept is covered with a solid roof plate. The three-dimensional view of the concept is shown in Figure 2 6 and in Figure 2 7. In Figure 2 8 a detail of the trailing edge region is shown demonstrating the outlet of the deflected stream of air. The air leaving the pressure chamber flows on either side of the purple surface and the Coanda effect is generated on the lower side of the purple wall and on the lower side of the wing as a result of interaction between the nozzle stream passing on either side the narrow purple surface and the flow passing the wing lower surface. This way the air on the lower surface of the wing is accelerated in the trailing edge region producing area of low pressure there. The rapidly diverted upwards air stream generates stagnation region near the trailing edge on the upper side of the wing. Both phenomena combine to modify lift in the wing segment where the actuator is located. This actuator is one-directional – in the current configuration it may only be used for reduction of lift.The results of flow simulations revealed that the DTEN actuator was working in two modes: for lower values of nozzle mass flow rate, up to 0.1 kg/s the nozzles were acting similar to a Gurney tab, deflecting the flow but without producing the Coanda effect. Only after reaching this value of nozzle mass flow rate the Coanda effect appeared and upward deflection of flow behind the trailing edge was accompanied by appearance of a negative pressure region in the lower part of the trailing edge, supporting the load alleviation effect. The maximum wing-root alleviation level was 32% and this was achieved at nozzle mass flow rate of 0.15 kg/s. The side-effect of this concept was a nose-up increase of the pitching moment which was equal to Cm=0.068 at the maximum load alleviation effect. For the Fluidic spoiler the pitching moment changes were lower; depending on the location of the number and location of the active rows the pitching moment changes did not exceeded the value of Cm=0.001 positive or negative, depending on the chord position of the active nozzles. It must be noted, however, that changes of pitching moment occur also with the traditional load-alleviation systems, such as symmetrical deflections of ailerons on the Lockheed C-5 Galaxy aircraft. They have to be countered by compensatory deflection of elevators.For the DTEN actuator the wind-tunnel investigations confirmed the predictions of numerical simulations that at low values of the nozzle mass flow rate the Coanda effect does not appear or has low intensity, but the strong Coanda effect appeared at the nozzle mass flow rate of 0.055 kg/s which was similar mass flow rate to the rate necessary for activation of the most effective variant of the Fluidic Spoiler (generating flow separation on suction side of a wing). The maximum achieved wing-root bending moment alleviation effectiveness of the DTEN actuator was 31% which was slightly higher than for the most effective variant of the Fluidic Spoiler, but at this maximum effectiveness level the saturation effect was noticeable.It must be noted also, that both investigated concepts had higher load-alleviation capability than traditional spoiler of 10% wing section chord. For the classic spoiler located in place of the Fluidic Spoiler the wing-root bending moment alleviation level was 21% and this was achieved at a high deflection of 35 degrees. In practical situations achieving such high deflection requires large actuation moments produced by the hydraulic system and generates large increase of drag. For fluidic Spoiler the drag increase is moderate and results from the decreased suction of the leading edge due to decreased velocity circulation. In case of the DTEN actuator there occurs another component of drag – the suction on the rounded trailing edge which acts in the rearward direction. For this reason the drag increase due to activation of the DTEN device is higher than due to deflection of aileron, but lower than due to deflection of spoiler.4. The Leaky Wing conceptAs part of complementary actions, not included in the original scope of the project the "Leaky Wing" concept has been developed and investigated by numerical flow simulations. The complementary actions were aimed at increasing chances of further development of the proposed concepts in future projects. They included the development of the Leaky Wing concept and conducting numerical simulations of dynamic changes of aerodynamic loads and load alleviation effects in simulated gust conditions.The concept of the "Leaky Wing" has been developed as a particular case of Fluidic Spoiler, utilising only natural pressure differences between wing pressure and suction surfaces. During the design process of the concept, a number of requirements and constraints were taken into consideration, including the simplicity and feasibility of the concept, minimal interference with the wing structure (which could weaken the structure) and high efficiency, reliability and response rate of the proposed system in mitigating excessive aerodynamic loads. The proposed concept consists of a matrix of transverse ducts connecting upper and lower surface of the wing. The ducts should be placed rather in outer part of the wing, because this part is a source of largest bending moments acting on the wing structure. On the other hand, such placement of ducts may be risky because of possible unfavourable influence of "fluidic spoiler" on the aileron effectiveness. However this aspect needs further investigations.In normal flight conditions, the ducts are closed, preferably in such a way as not to interfere with the flow around the wing. In extraordinary flight conditions, when excessive bending loads of the wing may occur (e.g. during sudden gusts, accelerated manoeuvres, etc.) the transverse ducts are opened, which should initialise intensive air flow through the ducts, being the effect of considerable difference of static pressure between the suction and pressure sides of the wing. Such phenomenon is typical for the flight conditions with occurrence of high aerodynamic loads of the wing. After opening the air ducts, the flow through them is expected to initiate the separation of the main flow on the wing suction side. In situations when main-flow separation already exists before opening the ducts, it is expected that flow through the ducts may enhance effects of existing separation. In both cases, the opening of the ducts in high-wing-load flight conditions, should decrease difference in static pressures on suction and pressure side of the wing, this way decreasing aerodynamic loads acting on the wing structure.The effectiveness of the "Leaky Wing" concept and its response rate in conditions of sudden gust, were investigated by series of URANS simulations. The simulations were conducted for gust conditions at two flight speeds: M=0.20 and M=0.45. A simple closed-loop control system of the load alleviation was simulated, based on the value of Load Factor (LF=Lift/Weight). The load alleviation system was activated when Load Factor exceeded threshold value of 1.3 and deactivated when it fell below 1.2. In the simulated gust conditions the measure of load-alleviation effectiveness was the relative difference between the wing-root bending moments measured for inactive and active wing-load-alleviation system. Based on conducted flight simulation, the effectiveness was estimated to be around 17-18%. Results of the simulations confirmed that the proposed "Leaky Wing" system quickly responds to increasing aerodynamic loads and its load-alleviation response effect should be quick enough to protect wing structure against fatigue damage. 5. ConclusionsAs a conclusion it may be stated that the results of the conducted works have proven the load alleviation effectiveness of the developed concepts. The load alleviation effectiveness of Fluidic Spoiler was higher than the effectiveness of a classic spoiler of the size of the actuating plate of Fluidic Spoiler. It must be noted, however, that the presented solutions of fluidic load control are still at very low technology readiness level (TRL2 or TRL3). More research is needed, concentrated on shape, dimensions, positions and characteristics at higher Mach numbers including possibilities of integratiion with wing structure and aeroelastic properties of wing equipped with such systems.Potential Impact:Description of project potential socio-economic impact including wider societal implicationsInvestigations conducted in the STARLET project concern active control of aircraft aerodynamic loads by fluidic flow control actuators. They are focused on alleviation of excessive aerodynamic loads in off-design conditions which occur in gusts or rapid manoeuvres by flow control techniques as an alternative to using proven mechanical solutions such as symmetrical deflections of ailerons or deflections of spoilers.. This is area of research is relatively new, particularly as far as the fluidic load control technology is applied for alleviation of excessive loads on large transport aircraft. Importance of the completed investigations may be considered as their potential impact in the following fields:- aeronautical research and technology development,- economical aspects of the developed technology,- societal aspects of the research and development of technology.In the field of aeronautical research the results of the STARLET project prove the viability of the application of fluidic flow control to the control or redistribution of aerodynamic loads on aircraft wing. Significant load alleviation effects have been obtained by the developed flow control actuators. The technology is, however, on a low level of technology readiness. For successful application on a transport-class aircraft it needs work on reliable and compact sources of air mass flow and air momentum actuating the load control system and it needs also work on integrating of the flow control actuators with wing structure. Some of the manufacturing technologies applied in the STARLET project for the manufacturing of flow control actuators, such as three-dimensional printing can probably be applied for the development of new generation of actuators, more integrated with wing structure. Assuming that work in the directions mentioned above is continued, on can expect that as a result of such work, and the work conducted already in the STARLET project a new generation of efficient and effective load control solution will be developed which has shorter reaction time than the classical mechanical devices and produces more effective alleviation of aerodynamic loads. Such achievements may lead to the possibility of design of lighter aircraft structures and improving passenger comfort.The economic aspects of the developed technology depend on the possibility of the development of systems integrating the proposed flow control actuators with existing aircraft systems. Particularly, the efficient source of air mass flow and air momentum which is the critical factor for the effectiveness of the developed solutions is needed in order to achieve solutions on high technology readiness level, competitive with exixisting, mechanical solutions. If the issue of integration of the developed concepts with aircraft structure and systems is settled, the economic benefits may result from lighter wing structure. designed for decreased maximum loads or maximum deflections – whichever is more critical given the materials used in the structural design. One of the solutions developed in the STARLET project, the Leaky Wing concept has the important advantage of using the natural difference between static pressure on lower and upper wing surfaces to actuate the operation of the device. In this respect it may be most likely candidate for further development of efficient, integrated with aircraft wing structure fluidic load control system. Additionally, considering the economic aspects of development of an innovative technology it may be observed that a new technology, if more effective than an old one, doesn't have to be always cheaper than its predecessor. A good example for comparison is the retractable undercarriage which is more expensive and heavier than the fixed one. However, the aerodynamic benefits of using the retractable undercarriage justify its adoption on transport-class aircraft and make the final product - the aircraft an economically efficient means of transport. Another socio-economic aspect of the realisation of the STARLET project is increased professional experience and design capabilities of the project crew and subcontractor providing high-precision elements of the wind-tunnel wing model. The experiences gained in the project will allow to conduct future design works of similar systems faster and with greater confidence in the properties of the final systems.The societal aspects of the work conducted in the STARLET project may be considered in terms of the potential of the developed concepts for development to a higher technology readiness level in the main field of application – the aeronautical industry – and in other potential fields of development and application, such as wind turbines. In this respect the Leaky Wing concepts may have the potential for being applied as a solution for controlling the torque generated by a wind-turbine blade and the total aerodynamic force generated by the blade. In this case the potential societal aspects of the development of this solution may include further research in larger number of research facilities and the addition of a new technology to the set of technologies applied by the aeronautics and the wind energy sector which could result in a number of additional jobs in these sectors.The societal aspects of the STARLET project include also the potential effects of dissemination of the knowledge gained in the project. The dissemination of new knowledge achieved as an effect of realisation of a research project increases the public awareness of the state of the art of aeronautical research and increases confidence in air transport. Also increased is the awareness of the capabilities of European researchers, designers and manufacturers in the field ofaeronautical research. The results have been already presented in five scientific conferences, with two of them very important for presenting results of aeronautical research to the scientific community: the Inernational Congress of Aeronautical Sciences congress in 2014 in St. Petersburg, Russia, and the European Aeronautics Science Neteork Workshop in 2014 in Aachen, Germany. In both cases the presentations included the results of investigations of the possibility of alleviating rapidly rising aerodynamic loads in gusts. The results of the unsteady flow simulations have indicated rapid reaction of the flow and of aerodynamic loads to the aerodynamic actuation. The results presented on the EASN Workshop included also the results of wind tunnel tests that have proven, by the measurments of pressure distributions and wing bending loads the alleviation effectiveness of the developed solutions. Apart from these, the obtained results have been presented also on two other scientific conferences of more general scope, addressing specialists from more fields of research. The first one of these is the 40th International Scientific Congress on Powertrain and Transport Means, European KONES 2014 held in Jastrzębia Góra, Poland which is attended by specialists of engine design and of means of transport of different categories, including land, sea, and air transport. Two papers presenting results of the project have been presented on this conference. The second conference was WAMS - Workshop on Applied Modelling and Simulation, co-located with NATO CAX Forum, held in 2014 in Istanbul, Turkey. The presentation of the results of the STARLET project took place on the joint session of the two conferences and atracted questions regarding the aerodynamic effectiveness of the proposed flow control solutions. In addition to this a web page dedicated to the STARLET project has been created with link to it from the official web page of the Instytut Lotnictwa: http://ilot.edu.pl/startlet This page will be maintained after completing the work in the project. The list of publications includes:1. Wieńczysław Stalewski "Wing-Load Control Based on The "Leaky Wing" Concept " - proceedings of the 29th Congress of the International Council of Aeronautical Sciences ICAS2014. St. Petersburg, Russia, 2014.2. Wieńczysław Stalewski, Janusz Sznajder, "Wing Load Control via Fluidic Devices" - proceedings of the 4-th EASN Association International Workshop on Flight Physics &Aircraft Design, Aachen, 2014 (ISSN number will be assigned for the proceedings).3. Andrzej Krzysiak, " Wind tunnel investigation of the wing load control using self-supplying fluidic devices", Journal of KONES Powertrain and Transport, Vol.21 No.2 2014, pp.152-160 ISSN 1231-4005.4. Wieńczysław Stalewski, Janusz Sznajder, "Modification of Aerodynamic Wing Loads by Fluidic Devices", Journal of KONES Powertrain and Transport, Vol.21 No.3 2014, pp.271-278 ISSN 1231-4005 5. Wieńczysław Stalewski, Janusz Sznajder, "Computational Simulations Of Smart Aircraft-Wing-Load-Control Systems Based On Innovative Fluidic Devices", proceedings of Workshop on Applied Modelling and Simulation, WAMS 2014, pp.21-26 ISBN 978-88-97999-46-1.List of Websites:http://ilot.edu.pl/starlet/ Powiązane dokumenty final1-publishable-summary.pdf
