cfd analysis of complete helicopter configurations – lessons learnt from the goahead project

Aerospace Science and Technology 19 (2012) 58–71

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CFD analysis of complete helicopter configurations – lessons learnt from theGOAHEAD project

R. Steijl, G.N. Barakos ∗

CFD Laboratory, University of Liverpool, Liverpool L63 3GH, United Kingdom

a r t i c l e i n f o a b s t r a c t

Article history:Received 19 July 2010Received in revised form 2 January 2011Accepted 30 January 2011Available online 15 February 2011

Keywords:GOAHEADHelicopter rotorRotor–fuselage interactionCFD

The GOAHEAD project, funded under the 6th European Funding framework, provided valuable measure-ments of flow parameters for a realistic helicopter configuration. The wind tunnel investigations includedan extensive set of conditions from cruise at high speed, to very high speed flight as well as high diskloading cases. Several GOAHEAD partners, including the University of Liverpool, contributed CFD sim-ulations for the full helicopter in the blind test phase (prior to the wind tunnel test), as well as, thepost-wind tunnel test phase. The Helicopter Multi Block solver (HMB) of Liverpool University was theonly in-house code from the UK to be used in this project. To account for the relative motion of rotor(s)and fuselage, the sliding-plane approach was used. As a first step of the project, a family of multi-blockCFD meshes was developed at Liverpool designed to work with the sliding-plane method. For the blindtest phase, the tail rotor was omitted. Using the lessons learnt from this first phase, a more advancedmulti-block topology was developed for the post-wind tunnel phase of the project, which allowed mainand tail rotors to be included. The pre-test computations for the economic cruise condition were foundto be in good agreement with the experiments when comparing surface pressure at various places on thefuselage considering the relative coarseness of the employed grids. Also, the CFD results of the variouspartners agreed reasonably well. As expected, the main discrepancies were in the separated-flow regionsat the back of the helicopter. The improved meshes used in the post-test phase resulted in better spatialresolution of the flow in addition to having the added complexity of the tail rotor. These new sets ofresults were in better agreement with measurements and were also performed on finer meshes. Clearly,the quality of the CFD mesh is key for accurate predictions and an educated guess of the flow regionswhere severe interactions of flow structures will occur is of importance for such complex CFD compu-tations. Remarkably, the efficiency of the CFD solver was high, and CFD analyses on meshes of up to 30million cells were performed during this project.

© 2011 Elsevier Masson SAS. All rights reserved.

1. Introduction

In recent years, Computational Fluid Dynamics methods havebeen increasingly used in the design and analysis of rotorcraft.This trend was made possible by the progress in CFD algorithmsand the availability of ever more powerful affordable computers.For hovering rotors, often simulated as a steady-state problem fora single blade, the computational overhead has been reduced suffi-ciently to enable the routine use in the design process of helicoptermain and tail rotors. For isolated rotors in forward flight, CFD anal-ysis has become feasible as well. However, the aero-mechanicsof an isolated rotor is still a very challenging area, since it con-stitutes a complex multi-disciplinary problem involving complexvortical wake flows, transonic flow regions, rotor blade dynamics

* Corresponding author.E-mail addresses: [email protected] (R. Steijl), [email protected]

(G.N. Barakos).

1270-9638/$ – see front matter © 2011 Elsevier Masson SAS. All rights reserved.doi:10.1016/j.ast.2011.01.007

and blade elastic deformation. In the aero-mechanics of a full heli-copter configuration, the additional flow physics introduced by theaerodynamic interactions between the main rotor, tail rotor andfuselage have to be considered. Regardless of its importance, in-teractional helicopter aerodynamics has so far been considered byvery few researchers, mainly as a result of the complexities men-tioned above [15,14,13,17,1,19,20]. In addition to the complex flowphysics, the geometric complexity of a full helicopter configurationintroduces significant challenges to wind tunnel experiments, aswell as, CFD investigations. Therefore, most of the published worksconcern wind tunnel experiments with generic helicopter rotorsmounted on idealised fuselages, e.g. the rotor-cylinder test case atGeorgiaTech [4,11] and the ROBIN test case at NASA [9,7,8]. In thefirst example, the airframe was represented by a circular cylinderwith hemispherical nose, while the ROBIN fuselage was closer toa real helicopter fuselage shape, but tail planes, engine inlets, ex-hausts, etc. were ignored.

Despite the recent progress in the numerical simulation of he-licopter flows and the demonstrated capability by several research

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R. Steijl, G.N. Barakos / Aerospace Science and Technology 19 (2012) 58–71 59

Fig. 1. Blind test phase. Main rotor–fuselage modelled using a single sliding plane separating main rotor and fuselage meshes. Block edges for meshes along sliding plane areshown with solid lines. Tail rotor excluded from geometry.

Fig. 2. Post-WT test phase. Full geometry modelled using a total of 6 sliding planes separating main rotor, tail rotor and fuselage meshes. For both rotors, 3 sliding planesform a ‘drum’ in which rotating rotor mesh is embedded in back ground fuselage mesh.

groups in the CFD analysis of complete helicopters [3,6,12,5,15–17], the present state-of-the-art in CFD investigations of full heli-copter configuration has not yet reached the maturity of numericalinvestigations of hovering rotors or isolated rotors in forward flight.A major factor has been the lack of adequate wind tunnel flighttest data for validation purposes. Therefore, an urgent need existsfor a database of high quality experimental data, which can actas validation for the state-of-the-art CFD methods. To address thisneed, the European Commission funded the Framework 6 ProgramGOAHEAD, with the aim to create an experimental data base andto validate state-of-the-art CFD methods.

The GOAHEAD experiment was built around a developmentwind tunnel model of the NH90 aircraft [2]. Main and tail rotorswere present during the tests and a variety of data was gathered.Surface pressure data (steady and unsteady) were combined withpressure and deformation measurements on the rotors, PIV investi-gations at several locations, as well as, hot film surveys and forcesfrom a balance. The wind tunnel investigations included an exten-sive set of conditions from cruise at high speed, to very high speedflight as well as high disk loading cases.

Regardless of the challenging geometry to model with CFD (ashighlighted in Figs. 1 and 2) and the necessity to resolve severalinteractions, calculations for this case were performed by severalpartners across Europe. The Helicopter Multi Block solver (HMB) ofLiverpool University [1,21,19] was the only in-house code from theUK to be used in this project. To account for the relative motion ofrotor(s) and fuselage, the sliding-plane approach was used, as de-tailed previously in Refs. [19,20], and explained briefly in Section 2.The CFD grids available to the GOAHEAD partners were designedto work with an over-set grid or CHIMERA approach, since this wasthe approach most commonly used by the CFD research groups in-volved. Therefore, the CFD Laboratory at the University of Liverpoolalso designed and generated in-house multi-block meshes for theGOAHEAD test cases suitable to work with the sliding-plane ap-proach. This mesh generation process and the lessons learnt aredescribed in Section 3.

The results of the pre- and post-test computations for the Eco-nomic Cruise test case of the full helicopter model are described

in Section 4, while Section 5 presents results for the Dynamic Stalltest case. Finally, conclusions are drawn in Section 6.

2. The HMB CFD solver

The Helicopter Multi-Block (HMB) CFD code [1,21,19] was em-ployed for this work. HMB solves the unsteady Reynolds-averagedNavier–Stokes equations on block-structured grids using a cell-centred finite-volume method for spatial discretisation. Implicittime integration is employed, and the resulting linear systems ofequations are solved using a pre-conditioned Generalised Conju-gate Gradient method. For unsteady simulations, an implicit dual-time stepping method is used, based on Jameson’s pseudo-timeintegration approach [10]. The method has been validated for awide range of aerospace applications and has demonstrated goodaccuracy and efficiency for very demanding flows. A detailed ac-count of application to dynamic stall problems can be found inRef. [18]. Several rotor trimming methods are available in HMBalong with a blade-actuation algorithm that allows for the near-blade grid quality to be maintained on deforming meshes [21].

The HMB solver has a library of turbulence closures which in-cludes several one- and two-equation turbulence models and evennon-Boussinesq versions of the k–ω model. Turbulence simulationis also possible using either the Large-Eddy or the Detached-Eddyapproach. The solver was designed with parallel execution in mindand the MPI library along with a load-balancing algorithm are usedto this end. For multi-block grid generation, the ICEM-CFD Hexacommercial meshing tool is used and CFD grids with 10–30 mil-lion points and thousands of blocks are commonly used with theHMB solver.

The underlying idea behind the sliding-mesh approach, as wellas, details of the implementation in HMB were previously de-scribed in Refs. [19,20]. The method can deal with an arbitrarynumber of sliding planes between meshes in relative motion.The main requirement is that the grid boundary surfaces of twomeshes on either side of a sliding plane match exactly, whilethe mesh topology and meshes can be, and in general are, non-matching.

60 R. Steijl, G.N. Barakos / Aerospace Science and Technology 19 (2012) 58–71

Fig. 3. Comparison of geometry used in the CFD simulations for the blind-phase and post-WT phase of the GOAHEAD project. Main differences are in horizontal tail planeand wind tunnel support.

3. Mesh generation

The GOAHEAD geometry comprises a wind tunnel model of theNH90 with the 4-bladed ONERA 7AD main rotor, equipped withanhedral tips and parabolic taper, and the BO105 2-bladed tail ro-tor.

The pre-test phase CFD geometry was based on the CAD modeloriginally used to produce the wind tunnel model. The wind tun-nel support was an approximation of the support planned for thetest. The model actually tested had a more streamlined wind tun-nel support. Also, the fuselage geometry was different from theoriginal CAD model in a number of ways. In our computationsfor the post-test phase, the actual wind tunnel support was rep-resented correctly, while the rest of the fuselage geometry was leftunchanged, with the exception of the correction of the horizontaltail plane anhedral. The pre-test and post-test geometries are com-pared in Fig. 3. The solid black lines denote the block edges of themulti-block mesh on the surface. The mesh parameters for the dif-ferent grids used in the project are listed in Table 1. The grids

generated for the project were designed for Reynolds-averagedNavier–Stokes simulations without the use of wall-functions, i.e.sufficient near-wall resolution was required to ensure an adequateresolution of the boundary layers.

The rotor meshes employed during the GOAHEAD project werebuilt on a C-H type multi-block topology, where the H-type topol-ogy in the span-wise direction takes into account the blade rootand tip by incorporating 4 prism-shaped blocks emanating fromboth ends of the blade. The meshes have a C-type topology inthe chord-wise direction, with a good spatial resolution of bothleading-edge and trailing-edge of the blades. More details of thistype of multi-block meshes for rotors can be found in Ref. [21].

For the pre-test phase, the tail rotor was omitted from the ge-ometry. This allowed the use of a single sliding-plane surface sepa-rating the fuselage mesh and the main rotor mesh. This plane wasconstructed to be normal to the rotor shaft. Naturally, this singlesliding plane required the use of a cylindrical ‘far-field’ boundarycondition. Fig. 1 shows the block edges of both rotor and fuselagemeshes along the sliding plane in the vicinity of the fuselage. The


clearance of the sliding plane with respect to the tail rotor gearboxfairing is apparent from these figures. However, the single slidingplane would intersect the tail rotor disk in case the tail rotor wouldhave been included.

For the post-test phase a more advanced multi-block topologywas therefore developed which enabled both main and tail rotorsto be included. A significant change relative to the pre-test fuselagetopology was the use of the concept of embedding local O-type

Table 1Meshes used in GOAHEAD project.

Phase Blind Interm. Post-WT

Fuselage: Blocks 1624 2298 2308Cells 6.5 · 106 13.9 · 106 14.0 · 106

Main rotor: Blocks 856 1112 1112Cells 4.1 · 106 11.1 · 106 11.1 · 106

Tail rotor: Blocks – 376 376Cells – 2.8 · 106 2.8 · 106

Total: Blocks 2480 3786 3796Cells 10.6 · 106 27.8 · 106 27.9 · 106

sub-topologies into the topology for the post-test phase. As be-fore, the topology has a 1-to-1 block face connectivity throughout,which naturally leads to the large number of blocks in the em-ployed topologies. An interesting observation when comparing thetopology of the surface meshes for pre-test and post-test phases inFig. 3 is that the use of the embedded local O-type sub-topologiesresults in a reduced complexity of the surface topology and, forthe particular grids compared here, a more even distribution of themesh points with less pronounced localised refinements, as com-pared to the pre-test topology.

The multi-block topology for the post-WT phase of the projectwas developed before the actual wind-tunnel model geometry wasobtained. Therefore, an intermediate family of meshes was devel-oped, i.e. using the multi-block topology for the post-test phaseand the CFD geometry from the pre-test phase, excluding the windtunnel support. For the post-WT meshes, the only change was theaddition of the new wind tunnel support as well as the modifica-tion of the horizontal tail plane anhedral angle. Figs. 2, 4 and 5show the geometry and the multi-block structured mesh for the

Fig. 4. Comparison of meshes used in the CFD simulations for the blind-phase and post-WT phase of the GOAHEAD project.


Fig. 5. GOAHEAD full helicopter geometry. The mesh in the y = 0 plane is shown, which does not constitute a symmetry plane. The rotor meshes are not shown for clarity.The mesh shown is the ‘intermediate’ mesh with WT support and has 3786 blocks and 27 · 106 cells. (a) Global view of mesh, (b) detail of mesh in nose region, (c) close-upof mesh in tail region.

intermediate and post-test phases of the project. For this full he-licopter geometry, both main and tail rotors are placed within adrum-shaped sliding-plane interface, as shown in Fig. 2. The closeproximity of the main and tail rotor planes are notable in the fig-ure, which leads to an additional challenge in the generation ofthe multi-block structured meshes used here. The main rotor drumhas the 5◦ forward tilt of the main rotor shaft, while the tail rotordrum is tilted about the x-axis as well as the z-axis (in the tail ro-tor hub-centred coordinate system) to provide a small forward andupward tail rotor thrust component.

4. GOAHEAD economic cruise case

The case considered corresponds to an economic cruise condi-tion, for which the free-stream Mach number is 0.204 and the tipMach number of the rotor 0.62. A representative rotor trim sched-ule is used in the simulation, i.e. the rotor has cyclic pitch changeas well as a harmonic blade flapping. The multi-block topology ofthe rotors is designed to handle the grid deformation as discussedin Ref. [21].

4.1. Interactional aerodynamics

Fig. 6(a) shows the instantaneous surface pressure distributionat main rotor azimuth 90◦ of the third revolution for the eco-nomic cruise condition at μ = 0.3. The effect of the blade pass-ing on the surface pressure distribution of the front part of thefuselage is shown in detail in Fig. 6(b), where x = 0.75 plane isshown. The main rotor blade passing through the front of the ro-tor disk clearly induces a (delayed) pressure rise on the forwardfuselage, as discussed previously in Ref. [19]. The interaction of

the tail rotor with the fin is shown in Fig. 6(c), showing the cpcontours in the z = 0.775 crosssection. The tail rotor blade is atψ = 0◦ , which corresponds to the downward vertical position. Forthe rotation of the tail rotor used here, this position is in the re-treating side of the tail rotor disk. The blade stagnation pressurein the selected cross-section was found to be only around twicethe fin stagnation pressure. In addition to the direct impulsive ef-fect, the tail rotor–fin interaction also includes the effect of the tailrotor induced velocity on the flow around the side-force generat-ing fin, by effectively changing the flow angle in a time-periodicfashion. This effect is more difficult to analyse than the pressureimpulse effect shown in the figure. A comparison of simulationresults with and without tail rotor would clearly show this contri-bution.

The main rotor–fuselage interactional effect on the rotor loadsfor the GOAHEAD model was investigated in detail in Ref. [20].As a first step, the flow around the full helicopter was computedusing the ‘intermediate’ mesh, i.e. the WT-support and the an-hedral of the horizontal tail plane were omitted. Then, the mainrotor mesh of this grid was embedded in a new background meshfor which the far-field boundaries coincided with the wind tun-nel walls, but did not contain the fuselage. Using the sliding-planeapproach, the rotor–wind tunnel system was computed with iden-tical computational parameters such as time steps and dual-timestep truncation criteria as the previously computed full helicoptercase. For both cases, the blade sectional loading as function ofthe sectional span-wise position and blade azimuth was extracted.Comparison of the two results therefore highlighted the effect ofthe presence of the fuselage on this blade sectional loading. Fig. 7shows an example of the comparison for the blade at the fore andaft positions. Sectional blade loads are compared for the full he-


Fig. 6. GOAHEAD full helicopter geometry. Economic cruise condition. Instantaneous pressure coefficients are shown. (a) Instantaneous surface pressure coefficient at mainrotor azimuth 90◦ , (b) main rotor–fuselage interaction, x = 0.75 crosssection (approx. mid span of blade), (c) tail rotor–fin interaction, z = 0.775 cross-section, at base of fin.

licopter and the equivalent isolated rotor case. The results werediscussed in more detail in Ref. [20], highlighting the capabilityof the present sliding-plane approach to quantify the effect of therotor–fuselage interaction. The results clearly showed the extent ofthe blade inboard stations affected by the presence of the fuse-lage. It was found that the interactional effect is mostly restrictedto the front and rear of the disk, i.e. the advancing side as well asthe retreating side loading do not change significantly due to thepresence of the helicopter fuselage. Secondly, the fuselage inducesan up-wash, which effectively increases the blade angle of attackthrough a significant area at the front of the rotor disk. In a simi-lar vein, the fuselage creates downwash behind the fairing, whichleads to a reduction of the blade angle of attack for parts of therear of the rotor disk. Compared to the up-wash area at the frontof the disk, the extend of the downwash area at the rear of thedisk appears to be smaller. In particular, at the front of the rotordisk, the interactional effect extends to blade stations further out-board, as compared to the interaction at the rear of the rotor disk.For all computations shown here, the trim state of the completeconfiguration was applied to the isolated rotor computations andfor this reason, the effect of the fuselage on the rotor trim is notinvestigated.

4.2. Comparison with experiments

Figs. 8 and 9 present the comparison between the unsteadypressure transducers on the GOAHEAD model fuselage and theHMB results for the economic cruise case. In the following, thistest case will sometimes be referred to as TC3-4. On the modelsurface, coloured spheres are placed near each transducer, with thecolour representing the value of the surface pressure coefficient.With minor exceptions of a couple of transducers on the frontof the fuselage and on the side of the engine housing, all trans-ducers agree remarkably well with the HMB predictions. Resultsare shown for four azimuth angles suggesting that the methodis capable of capturing at least the main pressure transients. Itis important to highlight here the key role of the experimentsand the need to process the experimental data the same way asthe CFD results. For Figs. 8 and 9, the GOAHEAD data were aver-aged in phase and with the same resolution as the obtained CFDresults. This allowed for adequate resolution of the pressure varia-tion without biasing the comparison. Fig. 10 presents a closer lookof the obtained surface pressures and compares the HMB solutionwith results from other partners as well. As can be seen, most CFDcurves are within close proximity to the experimental data. Better


Fig. 7. Effect of rotor–fuselage interaction on the blade loading for the GOAHEAD economic cruise test case. Sectional blade loading are compared for the full helicopter andan equivalent isolated rotor case.

agreement is obtained at the front of the fuselage and the agree-ment is still fair near the empennage and the fin of the model. TheHMB solution is shown for the V1 station and it appears to cap-

ture all features shown by the experiments. This is not the casefor station S4 that is located near the rotor hub and just upstreamof the exhausts. As can be seen, all CFD methods tend to predict


Fig. 8. Comparison between CFD and experiments for the Economic Cruise (TC3-4) case. The spots on the fuselage correspond to the unsteady pressure transducers. Economiccruise conditions for the full helicopter configurations.

surface pressure coefficients above the values indicated by the ex-periments. The situation improves substantially for the station S7that is located downstream in the rear fuselage. The discrepanciesat station S4 for a particular set of azimuth angles (near 0 de-grees) suggests that some vortex shedding from the hub may not

be adequately resolved by the CFD. This is apparently due to theapproximate hub geometry employed for computations and sug-gests that further work is needed in this area. Hub drag and theexact representation of hub geometry is perhaps an area of re-search to be looked at in the near future. Fig. 11 presents results


Fig. 9. Comparison between CFD and experiments for the Economic Cruise (TC3-4) case. The spots on the fuselage correspond to the unsteady pressure transducers. Economiccruise conditions for the full helicopter configurations.

for the blade loads for the economic cruise case. The first compar-isons shown in Fig. 11(a) show fair agreement between the HMBresults and experiments with some marked discrepancies near thefront of the blade for azimuth angles of 90 and 270 degrees. Forthat figure the raw experimental data were phase averaged and

used for comparisons. The discrepancies appear not to be presentin Fig. 11(b) where the experiments were processed removing faultor intermittently-working transducers and re-constricting the loadsof one blade using the good, working transducers of all otherblades. This was necessary in the GOAHEAD data due to a num-


Fig. 10. Comparison between experiments and CFD for three stations along the fuselage for the Economic Cruise (TC3-4) case.

ber of transducers failing or working intermittently during the test.The agreement between HMB and the processed data is now muchbetter and this suggests that careful consideration of the outcomeof the experiment is needed before comparing with CFD results.It is remarkable that overall, results contributed by several GOA-HEAD partners show good agreement between codes and goodagreement with experiments. The coarse resolution of the trans-ducers doesn’t allow for accurate chord-wise integrations thoughat all available stations along the chord, the agreement with theexperiments is more than encouraging, given the complexity anddifficulty of this flow.

The investigation of the main rotor–fuselage interactional aero-dynamics for the ‘intermediate’ mesh showed that with thepresent method, this important aerodynamic effect was cap-tured using the mesh with approximately 28 · 106 mesh points.As can be seen in Fig. 7 the effect of the fuselage on the ro-tor loads is strong even at stations as outboards ad 70% ofthe blade radius. The strong interaction leads to higher loadsfor the isolated rotor at the back of the disk and the oppositeis the case for the front. Although this phenomenon is well-understood by design engineers, it is very rare to see quanti-tative results for such cases, due to the difficulty of computingthe flow and the lack of appropriate test cases. The fidelity ofthe computations were ascertained by comparisons between theHMB results and the GOAHEAD experiments for the economiccruise.

5. GOAHEAD Dynamic Stall case

The GOAHEAD project included a test case at high rotor thrust.In this Dynamic Stall test case (denoted alternatively as TC5),a trim-state was selected for which the occurrence of dynamic stallwas predicted by simulations using the HOST comprehensive ro-torcraft method of ONERA [2]. This served as the trim state forthe pre-test phase. Preliminary CFD data based on that trim statewas then used to guide the selection of the trim state used in thewind tunnel test, i.e. a test case with structural rotor loads withinacceptable limits.

Fig. 12 shows the results for the blade loads for the DynamicStall test case. This test case has been computed by Liverpool asan isolated rotor both during the blind and the post-test phasesof GOAHEAD. Different trim states were used for the two sets ofcomputations, as mentioned before. Fig. 12(a) shows that like theeconomic cruise case, the HMB results are in fair agreement withthe raw data of GOAHEAD. Again, several marked discrepancies arepresent especially near 270 degrees of azimuth. Processing the ex-perimental data (shown in Fig. 12(b)) leads to a much better com-parison between HMB and the experiments. It is interesting to notethat along the blade chord and for all azimuth angles compared forthis case, there is no clear evidence of the presence of a dynamicstall vortex. This suggests that the selected trim state did not allowfor a fully-developed dynamic stall vortex. Fig. 13 presents resultsobtained during the blind and the post-test phase for the DynamicStall case. The trim state employed for the blind test phase al-


Fig. 11. Comparison of sectional rotor loads between experiments and CFD for the Economic Cruise (TC3-4) case.

lowed for a dynamic stall vortex to be developed outboards. Theobtained tornado-like structure is similar to results obtained by theauthors for 3D dynamic stall [18]. The post-test computation sug-gests high blade loading, without the evidence of a dynamic stallvortex. The good agreement with the pressure transducers suggestthat the CFD solution is not wrong but clearly the dynamic stallphenomenon is either not present or it is rather weak in this testcase. Dynamic stall presents a challenge for CFD predictions and

for this reason, further work is needed in this direction using thecomplete database of GOAHEAD and by looking at the raw PIV datafor further insights.

6. Conclusions

The HMB results contributed to the GOAHEAD project wereall computed on an in-house Linux clusters. Both the blind-test


Fig. 12. Comparison of sectional rotor loads between experiments and CFD for the Dynamic Stall test case.

rotor-body test case and the post-WT full-helicopter test case tookapproximately one month to compute. This reflects the fact thatthe improvements in the CFD method coupled with performanceupdates of the used Linux cluster introduced in the mean time,compensated for the increased mesh sizes in the second phase.A comparison of the results obtained using the HMB method with

experimental data for the GOAHEAD economic cruise test caseshows that the method is capable of resolving the main aerody-namic flow features for the coarse blind test phase mesh, as wellas, for the improved and refined post-WT mesh. The improvementsin mesh topology and the increase in mesh density clearly im-proved the spatial resolution of the flow. However, comparisons


Fig. 13. Results from the pre-test and post-test simulations for the Dynamic Stall test case. Surface pressure contours indicate a dynamic stall vortex in the pre-test simulation,while for the post-test simulations this type of evidence is absent.

of the blind test phase and post-WT data with the experimentaldata, as well as, CFD results of the other GOAHEAD partners showthat the effect of the differences in employed rotor trim state anddifferences in the used geometry seems to be at least as signif-icant as the improvements in the mesh topologies and sizes. Itappears that, overall, the computational framework in HMB is ad-equate for the estimation of the loads on the components of thehelicopter configuration. On the other hand, the need for trim data,blade structural properties or direct measurements of the bladedeformation shows that these areas require further investigation.In particular, the obtained results were sensitive to the employedtrim state and although the model blades employed for the testwere relatively stiff, knowledge of their exact shape during flightis important for accurate predictions. Furthermore, the effect of thecoupling with the structural response, despite the relative stiffnessof the wind tunnel model, is very significant, partly because ‘weak-coupling’ simulations will drive the rotor trim state away from theinitial trim state estimate from comprehensive methods.

Acknowledgements

During this work, R. Steijl was supported by the EuropeanUnion under the Integrating and Strengthening the European Re-search Area Programme of the 6th Framework, Contract No. 516074(GOAHEAD project). Some of the computations presented in thispaper were carried out using the Hector super-computer of the UKunder the EPSRC grant EP/F005954/1.

References

[1] G. Barakos, R. Steijl, K. Badcock, A. Brocklehurst, Development of CFD capabilityfor full helicopter engineering analysis, in: 31st European Rotorcraft Forum,Florence, Italy, 13–15 September 2005.

[2] O. Boelens, G. Barakos, M. Biava, A. Brocklehurst, M. Costes, A. D’Alascio, M. Di-etz, D. Drikakis, J. Ekatarinaris, I. Humby, W. Khier, B. Knutzen, F. Le Chuiton,K. Pahlke, T. Renaud, T. Schwarz, R. Steijl, L. Sudre, L. Vigevano, B. Zhong, Theblind-test activity of the GOAHEAD project, in: 33rd European Rotorcraft Fo-rum, Kazan, Russia, September 2007.

[3] O. Boelens, H. van der Ven, B. Oskam, A. Hassan, Boundary conforming dis-continuous Galerkin finite element approach for rotorcraft simulations, J. Air-craft 39 (5) (2002) 776–785.


[4] A. Brand, H. McMahon, N. Komerath, Surface pressure measurements on a bodysubject to vortex wake interaction, AIAA J. 27 (5) (1989) 569–574.

[5] A. Datta, I. Chopra, Validation of structural and aerodynamic modeling usinguh-60a flight test data, J. American Helicopter Society 49 (3) (2004) 271–287.

[6] M. Dindar, M. Shephard, J. Flaherty, K. Jansen, Adaptive CFD analysis for ro-torcraft aerodynamics, Comput. Methods Appl. Mech. Engrg. 89 (2000) 1055–1076.

[7] J. Elliott, S. Althoff, R. Sailey, Inflow measurements made with a laser velocime-ter on a helicopter model in forward flight—Volume I: Rectangular planformblades at an advance ratio of 0.15, Tech. Rep. TM-100541, NASA, 1988.

[8] J. Elliott, S. Althoff, R. Sailey, Inflow measurements made with a laser velocime-ter on a helicopter model in forward flight—Volume II: Rectangular planformblades at an advance ratio of 0.23, Tech. Rep. TM-100541, NASA, 1988.

[9] C. Freeman, R. Mineck, Fuselage surface pressure measurements of a helicopterwind-tunnel model with a 3.15-meter diameter single rotor, Tech. Rep. TM-80051, NASA, 1979.

[10] A. Jameson, Time dependent calculations using multigrid, with applications tounsteady flows past airfoils and wings, in: 10th Computational Fluid DynamicsConference, Honolulu, Hawai, June 24–26, 1991, AIAA Paper 1991-1596.

[11] J. Kim, N. Komerath, Summary of the interaction of a rotor wake with a circularcylinder, AIAA J. 33 (3) (1995) 470–478.

[12] N. Komerath, D. Mavris, S. Liou, Prediction of unsteady pressure and velocityover a rotorcraft in forward flight, J. Aircraft 28 (8) (1991) 509–516.

[13] H. Nam, Y. Park, O. Kwon, Simulation of unsteady rotor–fuselage aerodynamicinteraction using unstructured adaptive meshes, J. American Helicopter Soci-ety 51 (2) (2006) 141–148.

[14] Y. Park, O. Kwon, Simulation of unsteady rotor flow field using unstructuredadaptive sliding meshes, J. American Helicopter Society 49 (4) (2004) 391–400.

[15] Y. Park, H. Nam, O. Kwon, Simulation of unsteady rotor–fuselage aerodynamicinteraction using unstructured adaptive meshes, in: American Helicopter Soci-ety 59th Annual Forum, Phoenix, AZ, May 6–8, 2003.

[16] H. Pomin, S. Wagner, Navier–Stokes analysis of helicopter rotor aerodynamicsin hover and forward flight, J. Aircraft 39 (5) (2002) 813–821.

[17] T. Renaud, C. Benoit, J.-C. Boniface, P. Gardarein, Navier–Stokes computationsof a complete helicopter configuration accounting for main and tail rotor ef-fects, in: 29th European Rotorcraft Forum, Friedrichshafen, Germany, Septem-ber 2003.

[18] A. Spentzos, G. Barakos, K. Badcock, B. Richards, P. Wernert, S. Schreck, M. Raf-fel, CFD Investigation of 2D and 3D dynamic stall, AIAA J. 34 (5) (2005) 1023–1033.

[19] R. Steijl, G. Barakos, Sliding mesh algorithm for CFD analysis of helicopterrotor–fuselage aerodynamics, Int. J. Numer. Meth. Fluids 58 (2008) 527–549.

[20] R. Steijl, G. Barakos, Computational study of helicopter rotor–fuselage aerody-namic interactions, AIAA J. 47 (9) (2009) 2143–2157.

[21] R. Steijl, G. Barakos, K. Badcock, A framework for CFD analysis of helicopterrotors in hover and forward flight, Int. J. Numer. Meth. Fluids 51 (2006) 819–847.

cfd analysis of complete helicopter configurations – lessons learnt from the goahead project

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