Cone Guration Research Articles

Attitude Estimation for a Flexible Spacecraft in an Unstable Spin

Anattitudereconstruction algorithm hasbeendeveloped fora e exiblesounding rocketwhosespin vectornutates unstablyaboutitsminorinertia axis.Theattitudeestimatesareneededforposte ightanalysisoftherocket’ sscience data. An additional goal of the work is to show that a e exible-body model can be used in a Kalman e lter or in a smoother. The attitude is estimated using a smoother whose dynamic model includes a main rigid body and a pair of elasticbooms. Boom e exure is the principal causeof nutation instability. Boom bending is modeled by a Markov process, but the laws of mechanics are used to determine its ine uence on the attitude dynamics. This smoother achieves a better attitude estimation accuracy than can be achieved using a rigid-body model. The peak attitude error is estimated to be 4 deg, and the principal error source seems to be the limited accuracy of the rocket’ s attitude sensors. HIS work deals with the poste ight attitude determination for a sounding rocket mission, the Cleft Accelerated Plasma Experimental Rocket (CAPER). CAPER was launched from the Andoya Rocket Range in Norway in January 1999. Its e ight lasted slightly longer than 1200 s and reached an apogee altitude of 1360 km. The goal of this mission was to study the behavior of the ionosphere during auroral events. Attitude information is needed to transform CAPER’ s measurements of electric e eld components into geodetic coordinates. An attitude accuracy of from 2 to 4 deg is considered acceptable for the mission. From an attitude determination standpoint, the important characteristics of the CAPER sounding rocket were as follows: Its attitude sensors were a vector sun sensor with a slit e eld of view, a e xed-head horizon crossing indicator (HCI), and a three-axis magnetometer that was mounted on a short boom. The rocket’ s attitude was passively spin stabilized with the nominal spin vector directed along its minor inertia axis. CAPER deployed several booms after exit from the atmosphere and after the e nal stage motor burn. The longest of these were two e exible 3-m booms that deployed perpendicular to the nominal spin axis and in opposite directions from each other. A schematic diagram of this cone guration appears in Fig. 1. Minor-axis spin stabilization is often used in sounding rocket experiments. Near the dawn of the space age, the Explorer-1 mission demonstrated that such a cone guration has an unstable nutation mode due to energy dissipation in the e exible booms. 1 In many sounding rocket experiments it is acceptable for the nutation amplitude to grow as long as the resultant coning half-angle does not become too large by the end of the e ight. There are two major challenges in doing poste ight attitude determination for CAPER. The e rst is the lack of rate-gyro data. This challenge dictates the use of an Eulerian dynamics model to propagate the attitude and rates between measurement sample times. The model becomes especially important toward the end of the e ight, when only magnetometer data are available. The attitude estimate in rotation about the magnetic e eld vector is based solely on dynamic model propagation during this phase, and model inaccuracy

Tail Buffet Alleviation Through the Use of Wing-Strake Fillet Shapes

Various e llet shapes at the strake ‐wing juncture of a 76/40 deg double-delta wing model were experimentally investigated to gauge the effect of the e llet in alleviating adverse vortex/tail interaction. The vertical tails were instrumented with 28 fast-response pressure transducers. Pressure time histories and frequency power spectral densities were analyzed for four wing ‐strake e llet cone gurations and at three spanwise tail locations. Angle of attack was varied from i 2 to 40 deg, tunnel dynamic pressurewas 26.74 psf (1280 pascals ), and Reynolds number was 1.3 ££ 10 6 . The results show that the e llet geometry can increase or decrease tail buffet in comparison to the baseline, no e llet case. Linear and parabolic e llet shapes were found to increase tail buffet above the baseline case. A diamond e llet shape was found to alleviate tail buffet and shift the tail buffet to a higher frequency. The use of such a shape on e ight vehicles may result in a longer fatigue life of the vertical tails. Nomenclature A = constant,b=V1, 0.0091/s AR = aspect ratio, 2.41 a1 = speed of sound, 1155 ft/s, 352 m/s b = span, 1.36 ft, 0.414 m C = coherence CP = pressure coefe cient, p i p1=q1 c = chord, ft, m cR = root chord, 1.33 ft, 0.405 m cRtail = tail root chord, 0.38 ft, 0.116 m ct = tip chord, 0.21 ft, 0.064 m cttail = tail tip chord, 0.13 ft, 0.040 m f = frequency, Hz G = amplie er gain Iraw = raw integer value (digital) M1 = freestream Mach number, V1=a1, 0.13 P = pressure, psf Pref = reference pressure, psf q1 = tunnel dynamic pressure, 26.74 psf, 1280 pascals Re = Reynolds number, (Ω1V1CR/=π1, 1.3E6 S = wing area, 0.768 ft 2 , 0.071 m 2 Se = transducer sensitivity, V/psf Sr = Strouhal number, (bf /=V1

Low-Order Method for Predicting Aerodynamic Performance Degradation Due to Ground Icing

Alow-ordermethod to computethe degradation in the maximum lift coefe cient of an aircraft due to distributed surfaceroughnesshasbeendeveloped.ThealgorithmisappliedtotheFokkerF28Mk1000,anaircraftinvolvedinan accidentinwhichgroundicingwasdeterminedtoplayasignie cantrole.Themethodconsistsoftwocomplementary portions. The e rst uses a low-orderpanel method to computethequasi-steady aerodynamiccharacteristics, in and out of groundeffect,ofthecompleteaircraftgeometryincluding one- ortwo-elemente aps,a boundary-layerfence, nacelles, and a horizontal tail. The numerical model then, in the second portion, generates an engineering estimate of the maximum lift coefe cient for the aircraft cone guration. The method is based on the assumptions that the pressure difference between the suction peak and the trailing edge of an airfoil is a maximum at the maximum lift coefe cient and that the condition in which one spanwise station of a wing in a stripwise analysis violates this maximum pressure difference rule is the point at which the maximum lift of the entire aircraft is generated. The application of the pressure difference rule is modie ed to account for the presence of the boundary-layer fence and can estimate (on the conservative side ) the effects of distributed surface roughness on the maximum lift coefe cient using an appropriate maximum allowable pressure difference.

Synthesis of a FESTIP Airbreathing TSTO Space Transportation System

The recent Future European Space Transportation Investigation Program (FESTIP)system study has included onetwo-stage-to-orbit (TSTO)conceptwithanairbreathinge rststageandarocket-propelledsecondstage.Because ofthepureaccelerationmission from Kourou/CSGandthelimitedtechnologylevelthatshouldbeapplied,turbojet engines for staging at Mach 4 were assumed in the e rst stage. Two different TSTO cone gurations were studied in parallel: the fully integrated delta-wing cone guration and the less integrated trapezoidal-wing cone guration. The predesign activities were completed in July 1998, and a better data basis is now available for comparing the two design options in terms of launcher size, performance, complexity, operations, and cost. In spite of the high aerodynamic integration of the delta-wing cone guration, its overall performance is rather limited because of weak turbojet performance at high Mach numbers and altitudes. Inclusion of a ramjet for operation beyond Mach 3 would increase the optimum staging Mach number and overall complexity of this TSTO concept. The trapezoidal-wing cone guration makesbest useofa boost to beyond Mach 1.3 by rocketenginesof the second stage, which are fueled during the combined ascent by cross feeding from the e rst stage. The limited analyses performed on the double-wing aerodynamic interaction between the mated stages has not revealed any severe objections. This TSTO concept has exceeded the current FESTIP performance requirements, delivering 9 Mg into equatorial 250-km-reference orbit and 3 Mg into polar 250-km-reference orbit. With respect to cost, this launcher could be optimized by avoiding any stage integration. Then the internal tank arrangement of the e rst stage is not penalized by additional geometric constraints, and stage separation becomes simpler. Takeoff mass of the trapezoidal-wing cone guration is 470 Mg compared to 430 Mg for the delta-wing cone guration. Nevertheless, the vehicle dry mass, which drives development cost, is 40 Mg heavier for the delta-wing cone guration. The absolute cost e gures of both TSTO options are signie cantly higher than those of the single-stage-to-orbit vehicles investigated in the FESTIP program.

Effect of Flame-Holding Cavities on Supersonic-Combustion Performance

An experimental study was performed to evaluate the e ame-holding and mixing enhancement characteristics of supersonic reacting e ow over acoustically open cavities. Several cone gurations of acoustically open cavities were placed inside a supersonic-combustion duct just downstream of the fuel injection ports. The resulting changes in e ame behavior and combustion characteristics were assessed using schlieren visualization of the uncone ned e ow and wall pressure measurements of the cone ned e ow along the duct. The results were then compared with the baselinecase, which used no cavity. Although thecavitiesimproved thecombustion performancefrom thebaseline, the amount of enhancement was dependent on the particular shape of the cavity as well as the e ow conditions. Certain cavity cone gurations that were strategically placed inside the combustion duct led to a faster increase in the axial pressure force. The data showed that the recovery temperature was higher and the total pressure proe le was more uniform at the exit plane, suggesting enhanced volumetric heat release and faster mixing associated with the cavity- ine uenced e owe eld.

Shock-Wave/Turbulent Boundary-Layer Interactions in Nonequilibrium Flows

Shock-wave/turbulent boundary-layer interactions in the presence of thermal and chemical nonequilibrium phenomena are analyzed. The approach relies on a linear eddy viscosity two-equation turbulence modeling that accounts for the coupling of turbulence with chemistry and vibration, and it employs a total variation diminishing e nite volume numerical methodology. The capability of the model has e rst been assessed for a cylinder e are cone guration, and results have been compared with experiments. The model has then been applied to assess the aerodynamic performance of control surfaces of a reusable launch vehicle. In particular, the effects of wall temperature and e ap dee ection on the separation, aerothermal loads, and e ap efe ciency have been studied. The analysisshowsthatturbulencebecomesimportantfore apdee ectionanglesgreaterthanacriticalvalue[ O (15deg)], thus avoiding thecrisis of thee ap efe ciency that is observed under laminarconditions and extending the operating capabilitiesofthecontrolsurface.Thestudy also showsthatthewalltemperatureaffectssignie cantly theefe ciency and the operating envelope of the e ap primarily under turbulent conditions.

Multidisciplinary Optimization of Airbreathing Hypersonic Vehicles

Airbreathing hypersonic aircraft and missiles are characterized by a high degree of interdependence between airframe and engine. For nonaxisymmetric vehicles the propulsion system exerts a major ine uence on vehicle lift and pitching moment; this in turn ine uences vehicle stability, control, and overall mission performance. Because of strong interactions between the airframe and engine, conceptual design of this class of vehicle requires a multidisciplinary design optimization (MDO) process that can simultaneously account for the impact of selected geometric variables on all vehicle subsystems. This paper describes the development and implementation of an MDO design system that combines propulsion and external aerodynamic forces, mass properties and internal volumetric modeling, and performs geometric optimization of a hypersonic cruise missile to maximize overall mission range. The result is a cone guration with range 46% greater than the initial baseline. Such a dramatic performance increase is indicative not only of the power of optimization, but of the dife culty in cone guring hypersonic vehicles to synergize the interaction of all vehicle components without MDO methods.

Significance of Fuel Selection for Hypersonic Vehicle Range

Optimization studies of engine-integrated hypersonic aircraft for both cruising and accelerating missions tend to demonstrate noticeably lower lift-to-drag ratios than those of pure aerodynamic forms. One explanation of this is that, with low-density fuels such as hydrogen, matching lift to weight results in cone gurations that cannot take advantage of high-lift aerodynamics, especially when they operate at high dynamic pressures for airbreathing propulsion. Hydrogen has been the fuel of choice for hypersonic e ight because of its rapid burning rate, high specie c energy content, and good heat transfer properties for active cooling and recuperation. In contrast, hydrocarbon fuels have much longerburn times and substantially lower specie c energies, although they havesubstantial packaging and handling advantages. To study this, range performance is evaluated in terms of fuel density for hypersonic craft with time-varying lift-over-drag ratio. Historical data and geometric scaling are used to show that with hydrocarbon fuels, which have an order of magnitude greater density than hydrogen, hypersonic cruiser designs can take greater advantage of optimal aerodynamics. As such, hydrocarbon engine-integrated vehicles may have comparable or superior cruise range performance in comparison to cryogenically fueled designs.

New Tools for Computing Tight Bounds on the Real Structured Singular Value

New tools are presented for the computation of tight lower bounds on the structured singular value π, for highorder plants subject to purely real parametric uncertainty. The e rst approach uses the π-sensitivity function to systematically reduce the order of the real uncertainty matrix, so that exponential time lower bound algorithms can be applied. The second approach formulates the search for a worst-case real destabilizing perturbation as a constrained nonlinearoptimization problem.Both approachesareapplied to theproblem ofanalyzing thestability robustness properties of an integrated e ight and propulsion control system for an experimental vertical/short takeoff and landing aircraft cone guration. Currently available software tools for calculating lower bounds on real π fail for this problem, whereas both new approaches deliver tight bounds over the frequency range of interest.

Input/Output Conditioning of Robust Integrated Flight and Propulsion Controller

We describe the application of a conditioning scheme to a linear integrated e ight and propulsion control system for an experimental vertical/short takeoff and landing aircraft cone guration in the transition (approach to hover) region of the e ight envelope is presented. The proposed scheme guarantees satisfaction of limits on safety critical engine variables by using output conditioning to back off pilot demands in the event of a limit being encountered. Bounded-input/bounded-output stability results are presented for the resulting closed-loop system. At the plant input, an actuator linearization scheme is designed to minimize performance degradation due to backlashdynamicsintheaircraft’ sthrustvectoringnozzles.Theconditioningschemealsoprovidesamechanismfor explicitly prioritizing requirements on e ight-path angle overvelocity in the case of engine performance limitations being reached. Performance and safety improvements due to the addition of the controller conditioning scheme are demonstrated in nonlinear simulation.

Interactions Between a Canard and Thick Bodies: Analysis and Applications

A modular model, consisting of cruciform canard controls mounted on a thin forebody and e ve interchangeable thick (larger-diameter ) main bodies, was tested at a Mach number of 0.8. The canard was installed at the + + and x positions and at dee ection angles of 0, 6, and 12 deg. The model was equipped with two sting balances in a setup that enabled to obtain theloads acting on thecanard unitand those acting on themain bodies. Fourinteractionson mainbody normalforceandcenterofpressureareidentie edand quantie ed:effectsofthecanard on thethickening of the main bodies and on the boattails, as a function of angle of attack and canard dee ection. The interactions are analyzed using an approximate vortex tracker, which is based on slender body theory. Then, the pertinent part of the Pitts, Nielsen, and Kaattari component buildup method is used to calculate the interactions. The results of the analysis are in good agreement with the present test data, except for the effect of canard dee ection on the boattail, where the analysis leaves a large gap. Using the present method, to add the subject effects to component buildup analysis of variety of cone gurations improves the agreement with test data.

Single-Point and Multipoint Aerodynamic Shape Optimization of High-Speed Civil Transport

Single-point and multipoint aerodynamicshapeoptimization methods weredeveloped and demonstrated forthe designofanadvancedsupersonictransportsubjecttomanygeometricconstraints.Thestartingpointcone guration baseline was developed in support of the NASA High-Speed Research program using linear-theory-based methods and multidisciplinary system analyses. The single-point design method simulated the presence of nacelles and diverters at supersonic cruise by superimposing nacelles-on/nacelles-off pressure differences, from complete cone gurationanalyses,ontosingle-block-gridwing/bodycalculations.Themultipointdesignmethodusedamultiblock gridto treatthecompletecone guration,including nacelles/diverters, canard,empennage,and wing e aps/slats.Two forms of multipoint optimization were performed at Mach 2.4, 1.1, and 0.9: sequential (design at cruise followed by e ap and canard/tail incidence angle optimization at the two transonic conditions ) and multipoint (simultaneous design at the three e ight conditions via a composite objective function ). Euler-based optimization using a combination of the two methods achieved signie cant performance gains derived from the nonlinear effects. The single-point approach produced much of the improvement, lowering the appropriately weighted thrust coefe cient by 4.28 counts after trimming the full cone guration at the three design points. (A seven count drag reduction was achieved at cruise for the untrimmed vehicle. ) The sequential and multipoint methods achieved 6.03 and 7.55 counts of composite thrust reduction, respectively.

Dispersion Reduction of a Direct Fire Rocket Using Lateral Pulse Jets

Impact point dispersion of a direct e re rocket can be drastically reduced with a ring of appropriately sized lateral pulse jets coupled to a trajectory tracking e ight control system. The system is shown to work well against uncertainty in the form of initial off-axis angular velocity perturbations as well as atmospheric winds. For an example case examined, dispersion was reduced by a factor of 100. Dispersion reduction is a strong function of the number of individual pulse jets, the pulse jet impulse, and the trajectory tracking window size. Proper selection of these parameters for a particular rocket and launcher combination is required to achieve optimum dispersion reduction. Forrelatively lowpulsejetimpulse, dispersion steadily decreasesasthenumberofpulsejetsisincreased or as the pulse jet impulse is increased. For a e xed total pulse jet ring impulse, a single pulse is the optimum pulse jet cone guration when the pulse jet ring impulse is small because the effect of a pulse on the trajectory of a rocket decreases as the round e ies downrange. Nomenclature CDD = e n cant roll moment aerodynamic coefe cient CLP = roll damping aerodynamic coefe cient CMQ = pitch damping aerodynamic coefe cient CNA = normal force aerodynamic coefe cient CX0 = zero yaw axial force aerodynamic coefe cient CX2 = yaw axial force aerodynamic coefe cient D = rocket reference diameter ethres = trajectory tracking window size L; M; N = total applied moments about rocket mass center expressed in the aft body reference frame nJ = number of individual lateral pulse jets nRXi ;nRYi , = ith main rocket motor direction cosines in the nRZi body frame p;q;r = components of the angular velocity vector of the projectile in the body reference frame T = P ° time constant TJi = ith lateral pulse jet thrust TRi = ith main rocket motor thrust t ¤ = time of the most recent pulse jet e ring u;v;w = components of the velocity vector of the mass center of the composite body in the body reference frame uA;vA;wA = components of the velocity of the mass center of the projectile with mean wind expressed in the body reference frame VA = magnitude of the velocity vector of the mass center of the projectile experienced with mean wind expressed in the body reference frame VMW;aeMW = magnitude and wind factor of the mean atmospheric wind expressed in the initial reference frame X;Y; Z = total applied force components in the aft body reference frame

Design-Oriented High-Lift Methodology for General Aviation and Civil Transport Aircraft

A high-lift system design methodology that can be incorporated during the early stages of aircraft development is presented and thus has the potential to provide a superior and more cost-effective vehicle than one developed utilizing traditional linear design methods. The present methodology offers two different levels of e delity: one applicable to the conceptual design stage and the other to the preliminary design stage. The underlying e ow solver couples a three-dimensional nonlinear Weissinger method with two-dimensional viscous data to provide fast and accurate aerodynamic predictions for high-lift cone gurations. Several test cases that illustrate the capabilities of this hybrid e ow solver are presented.

Novel Design of a Bonded Lap Joint

A new design of a single-lap joint was proposed and investigated experimentally. In this new design, load eccentricity as well as singular peel stresses in the joint interface were avoided. In fact, numerical calculations show that, in the new design, the peel stress becomes compressive in the joint end region, and the shear load is more evenly transferred over the length of the joint. Two adherend layups, that is, (0/90/0/90) 2s and (90/0/90/0) 2s, were considered. Experimental results show that the strength of the new joint is signie cantly higher than that of the conventional single-lap joint. It is believed that even higher strength can be obtained by optimizing the new design cone guration. ANY structures consist of a number of basic parts that are connected together to take up the load. The connections or joints are usually the weakest locations in the structure. Among the many technical challenges facing engineers using advanced com- posite materials is joining composite parts or composite to metallic components.Eventhoughtheuseofcompositematerialscangreatly reduce the number of parts, joining parts is still unavoidable. Tradi- tional joining methods for metals such as welding and soldering are apparently unsuitable for e ber-reinforced composites. Compared to mechanical fastening, adhesive bonding is attractive for joining e ber-reinforced composite structures because it reduces the local- ized stresses encountered when using bolts and rivets. Moreover, bolts and rivets cut the continuous reinforcing e bers, and, as a re- sult, may greatly reduce the overall load carrying capacity of the structure. Adhesively bonded joints are used in many different cone gura- tions, among which the most commonly used are single-lap joints, double-lap joints, scarf joints, and step-lap joints. 1 3 The single-lap jointisgenerallythesimplestandleastexpensiveofalljointstoman- ufacture. However, because of its intrinsic load eccentricity and the abruptchangeofload path,boththe normal (peel)and shearstresses are highly localized at the two joint ends; thus, the conventional lap jointsare not efe cient in load transfer. 4 Moreover,because theinter- facial normal stress is tensile and theoretically singular at the joint ends, it can initiate failure at these locations. Although tapering the adherend or adding adhesive e llets at joint ends can lead to some reduction of peel stresses in the lap joint, these improvements are, however, quite modest. 5;6 Structural adhesives have relatively poor resistance to peel or cleavage stresses. Therefore, an obvious step to improve joint strength is to reduce the magnitude of peel stress or to make it compressive. To obtain maximum joint efe ciency, the interfacial shear stress distributions should also be made more uniform. In this paper, a newwavy bonded lap joint is proposed and inves- tigated. With this new joint design, the interfacial normal stress is compressiveatthetwojointends,andtheinterfacialshearstressdis- tribution is less localized. Experimental results show that the wavy joint has much greater strength than the conventional lap joint. Conventional Single-Lap and New Wavy-Lap Designs For comparison, a two-dimensional sketch of the conventional single-lap joint is shown in Fig. 1. Only the case of identical ad- herends is considered in this study. The length of both outer ad- herends is L, the overlap (joint) length is l, and the thickness of adherend and adhesive are ta and tg, respectively. the waviness angles (Fig. 2) ® DD tan 1 .2ta=l 0 ) and round out sharp corners of the joint with an adequate radius R. The e nal joint shape selected is shown in Fig. 3. The shape chosen here is mainly for simplicity; other possible shapes may be as effective. Strictly speaking, the two adherends in this wavy joint are not in a collinear cone guration. The dashed lines in Fig. 2 show this. If a collinear cone guration is desired, then the thickness of the adhesive must be taken into account during the e rst design step. Because the thickness of the adhesive is very small compared with that of the adherends in this study, it is neglected during the design process. The values of the geometrical parameters for both conventional and wavy joints are given in Table 1. Even though l 0 <l, the overall bonding areas are the same for the two joints. This ensures that analysis and test results of these joints can be directly compared. TheonlydimensionthatisnotshownintheFigs.1and2isthewidth of specimen, which is d D25:4 mm. The material for adherends is AS4/3501-6 carbon/epoxy. Two different layups, (90/0/90/0) 2s and (0/90/0/90) 2s, are used. The 0-deg direction is parallel to the longitudinal (loading)directionofthejoint.Thereasonforchoosing these two layups is to study the difference between 90 -90 deg and 0-0 deg bonding interfaces. The material constants of AS4/3501-6 are listed in Table 2. The two adherends are bonded together with a layer of e lm adhesive FM73M, which is treated as an isotropic material.Theelastic constants forFM73Mare also listed in Table2.

Spanwise Variation in the Unsteady Stalling Flowfields of Two-Dimensional Airfoil Models

Recent investigations of two-dimensional airfoil stalling characteristics have revealed low-frequency and highly unsteady e owin somecases and large-scalethree-dimensionalstructuresin other cases.Thelatterwerereferred to as “ stall cells” and can form on two-dimensional cone gurations where the ends of the airfoil model are e ush with tunnelsidewallsorend plates.Thispaperpresentsresultsofdetailedinvestigationsofthestalling characteristicsof several airfoils that exhibited both low-frequency unsteadiness and large-scale three-dimensional structures. The airfoils were wind-tunnel tested in a two-dimensional cone guration. The primary measurements were spanwise wakevelocityand mini-tufte owvisualization. Theresultsshowedthatairfoilswithtrailing-edgeseparationsatand above maximum lift (static stall )exhibited stall-cell patterns. Conversely, airfoils that had leading-edge separation bubblesthatgrewinsizeastheangleofattackwasincreasedintostalldevelopedthelow-frequency,highlyunsteady e ow. This unsteadiness was found to be essentially two dimensional. Therefore, the development of either of these phenomenaappearstobedeterminedbythecharacteristicsoftheboundary-layerseparationleadinguptothestall.

Three-Dimensional Navier-Stokes Simulations for Transport Aircraft High-Lift Configurations

Computations of lift and drag polars for a transport aircraft wing/fuselage high-lift cone guration using the MEGAFLOWcodesystemarecarriedoutandcomparedtowind-tunnelexperiments.Themainemphasisislaidon acomparisonoftheblock-structuredandtheunstructuredcodemodulesforsuchtypeofapplication.FortheblockstructuredFLOWercodeincombinationwith a k‐! turbulencemodel,thenumericalresultsarein good agreement with the available experimental data in the linear CL range. Beyond 15-deg incidence, a strong separation near the e ap cut-out is simulated, leading to an underprediction of total lift near CL; max compared to the experimental data. In contrast to this, the results of the unstructured TAU code utilizing the Spalart ‐Allmaras turbulence model are characterized by a nearly constant lift overestimation up to maximum lift without the aforementioned separation tendency at moderate incidences. The lift overprediction in the unstructured results is attributed to the main wing and the slat upperside suction peaks, which are higher resolved by the unstructured grid. Neither code reproduces the lift breakdown beyond CL; max according to the experiments. The use of preconditioning in conjunction with theFLOWercodeshowsonly minorimprovement of theaccuracy,but considerabledeterioration of the convergence properties, requiring improvements for routine use. Further studies will focus on the ine uence of geometry simplie cations at the wing root in the theoretical models and its impact on the experimental evidence.

Comparison of Glideback and Flyback Boosters

The aim of the present study is to investigate alternative reusable launch vehicle concepts and to provide a comparison of two different vehicle cone gurations. The e rst concept vehicle uses airbreathing engines to perform a powered booster e yback mission, whereas the second concept employs only aerodynamic lift to achieve booster e yback. The optimization software employs a system parameter optimization, in conjunction with a trajectory parameter optimization, to calculate the optimal mass split and corresponding optimal trajectory. A reference mission is used to comparelaunch vehicle payloads fora e xed grossliftoffweight, assuming launch from Woomera Prohibited Area, South Australia. The payload for a vehicle using a powered e yback strategy is found to be considerably higher than that obtained by a vehicle employing an aerodynamic glide return e ight.

Magnus Effect over Finned Projectiles

Computations of steady e ows over yawing and spinning axisymmetric projectiles are largely carried out by numerical algorithms using steady methods. Of particular interest is the prediction of the Magnus force and moment. However this axisymmetric characteristic is lost with e ns addition, and the e ow becomes unsteady whatever the framework is. ONERA and GIAT Industries have developed a new unsteady scheme, based on grid movement, that allows such a turbulent unsteady e ow to be solved. This scheme has been used successfully over a spinning and yawed body-tail cone guration. The Magnus effect is generated on the body by the spin-induced boundary-layer distortion at moderate incidences, whereas asymmetric vortices tend to invert this effect at upper incidences. Fins contribute to an opposite and greater lateral force. The total Magnus force appeared to be linear with respect to angle of attack and spin rate, but the range of linearity of angle of attack is much smaller than for a none nned body.

Cavity Flame-Holders for Ignition and Flame Stabilization in Scramjets: An Overview

This paper describes ongoing research efforts in the scramjet community on cavity e ame holders, a concept for e ame holding and stabilization in supersonic combustors. During the last few years, cavities have gained the attention of the scramjet community as a promising e ame-holding device, owing to results obtained in e ight tests and to feasibility demonstrations in laboratory-scale supersonic combustors. However, comprehensive studies are needed to determine the optimal cone guration that will yield the most effective e ame-holding capability with minimum losses. The e owe eld characteristics of cavities and research efforts related to cavities employed in lowand high-speed e ows are summarized. Open questions impacting the effectiveness of the cavities as e ame holders in supersonic combustors are discussed.