Forebody Drag Research Articles

Uticaj veličine i lokacije srednjeg aerodiska postavljenog oštrog vrha na smanjenje otpora preko hemisferičnog tela pri brzini od 2.0 maha

To minimize forebody drag in high-speed flying vehicles such as missiles and rockets, contemporary research has focused on computational methods to analyze drag reduction strategies. This study investigates the efficacy of an intermediate aerodisk mounted on a sharp-tip spike at a Mach number of 2.0. Through a parametric analysis, variations in aerodisk size and location on the spike stem are explored. Results indicate that reducing the size of the intermediate aerodisk to 3 mm maintains identical reattachment shock strength but leads to higher pressure values at the transition from separation shock to reattachment shock. The model with an expanded 5 mm aerodisk size exhibits the second-lowest peak pressure coefficient for reattachment shock, suggesting improved flow recirculation and lower heating levels. Conversely, a 6 mm aerodisk size increases reattachment shock pressure but enhances flow recirculation, impacting total drag. Overall, the study concludes that an intermediate aerodisk, particularly with a 5 mm diameter, provides an optimal configuration for drag reduction before flow separation.

Numerical study of flow over blunt bodies moving at supersonic speed

In the present study, shock waves produced over different blunt bodies moving at Mach 2.0 at zero angle of attack have been captured and analysed using computational investigations. Forebodies of distinct bluntness (hemisphere, ellipse, and ogive) are adopted, and the effect of the forebody geometry on the flow parameters is investigated. The flow parameters such as shock stand-off distance and shock strength have been captured over various forebody geometries and correlated with the forebody drag coefficient. With the reduction in the forebody bluntness, the shock stand-off distance and the shock strength are found to be decreasing, resulting in a decrease in the forebody drag coefficient values. The effect of the free-stream supersonic Mach number over a typical blunt body (hemisphere) was also investigated. With increasing freestream Mach number, the shock wave generated ahead the blunt hemisphere moved closer to the forebody and an increase in the forebody drag was observed.

Coupled Fluid-Thermal Investigation on Drag and Heat Reduction of a Hypersonic Spiked Blunt Body with an Aerodisk

In the traditional investigations on the drag and heat reduction of hypersonic spiked models, only the aerodynamic calculation is performed, and the structural temperature cannot be obtained. This paper adopted the loosely coupled method to study its efficiency of drag and heat reduction, in which the feedback effect of wall temperature rise on aeroheating is considered. The aeroheating and structural temperature were obtained by the CFD and ABAQUS software respectively. The coupling analysis of the hypersonic circular tube was carried out to verify the accuracy of the fluid field, the structural temperature, and the coupled method. Compared with experimental results, the calculated results showed that the relative errors of stagnation heat flux and stagnation temperature were 1.34% and 4.95% respectively, and thus the effectiveness of the coupled method was verified. Installing a spike reduced the total drag of the forebody. The spiked model with an aerodisk reduced the aeroheating of the forebody, while the model without an aerodisk intensified the aeroheating. The spiked model with a planar aerodisk had the best performance on drag and heat reduction among all the models. In addition, increasing the length of the spike reduced the drag and temperature of the forebody. With the increase of the length, the change rates of drag, pressure, heat flux, and temperature decreased gradually. Increasing the diameter of the aerodisk also reduced the temperature of the forebody, while the efficiency of forebody drag reduction first increased and then decreased. In conclusion, the heat and drag reduction must be considered comprehensively for the optimal design of the spike.

Multidisciplinary research method for designing and selection of bio-inspired profiles in the conceptual designing stage

A multidisciplinary research method was employed with the intention to create a series of bio-inspired flattened airfoils, observe their aerodynamic characteristics, and analyse their applicability to small devices or to designs of high-speed trains, within the shortest period in the conceptual stage. A research specimen of a kingfisher, selected for biomimicry, was examined with the following methods: visual inspection, analysis of photographs, manufacturing quality control measurement with a 3D laser scanner, and microscopy. A basic multi-arc-line profile, re-engineered from the overlapped specimen shape data and based on the observations, was used for designing a series of seven derived airfoils. The aerodynamic characteristics of the bio-inspired airfoils were obtained with the panel methods at low and moderate subsonic speeds, while the small transonic difference method was used in the high-subsonic speed range. Basic and ellipse-like airfoils produce higher total drag at low and moderate velocities and higher forebody drag in the high-subsonic range when compared to derived and parabola-like airfoils. The obtained critical Mach numbers are in the range from 0.76 to 0.78, where three bionic airfoils show values equal to or smaller than the values of ellipse- and parabola-like airfoils. The profile with the shortest bio-inspired relative chord has a higher critical Mach number value than the parabola-like profile. The sonic lines above these profiles appear at close positions. The applied set of examination methods of the bio-inspired design is not time consuming and produces sufficiently good results in the conceptual stage. Therefore, a further development of unique and adjusted numerical methods and codes at pre-computational fluid dynamics run is encouraged, together with shape parameterization.

Uticaj međuaerodiska sa montiranim oštrim šiljastim vrhom na smanjenje otpora po površini poluloptastog tela pri Mahovom broju 2.0

Reduction of forebody drag in high speed flying vehicles such as rockets and missiles are of high research interest in the present time. In the present research, drag reduction obtained by using an intermediate aerodisk mounted sharp tip spike has been investigated using computational studies at Mach number of 2.0. The flowfield over a hemispherical blunt body with an intermediate aerodisk mounted sharp tip spike is investigated at zero degree angle of attack and the amount of drag reduction obtained is then compared with that of a conventional sharp tip spike mounted hemisphere. The presence of an intermediate aerodisk changes the flow physics and shock system over the blunt body. The change in the system of shock waves by mounting an intermediate aerodisk results in a higher percentage (20% higher) of drag reduction generated by the blunt spiked body moving at a supersonic speed of Mach 2.0. Use of intermediate aerodisk proves to be beneficial in terms of drag reduction for spike lengths ranging beyond the critical length.

A new concept and preliminary study of variable hypersonic inlet with fixed geometry based on shockwave control

The inlet plays a significant role in the hypersonic airbreathing propulsion. A fixed-geometry inlet is troubled by low air-capture ratio and large additional drag while operating below the design Mach number. Whereas a variable-geometry inlet can maximize performance, but adds weight and complexity to the propulsion system. Based on a fluidic shock shape control technique, this paper proposes a new concept of variable hypersonic inlet with fixed geometry, gives the realization scheme, and conducts a preliminary validation. The results show that the control of the external shock system and the effective throat area can be achieved by the self-provided high pressure fluid of the inlet. For an inlet with an operating Mach-number region of 4 to 6, the shock-on-lip condition can be maintained from Mach 5 to Mach 6 with the maximum expense of 1.8% secondary flow ratio, resulting in 20% extra captured mass flow and 8% less forebody drag at low Mach numbers compared with conventional fixed-geometry inlets. Thus, the performance enhancement by using the proposed variable inlet can substantially benefit the acceleration process of hypersonic vehicles at low Mach numbers.

Minimum forebody drag in hypersonic continuum and rarefied flows

AbstractMinimum drag shapes of given length and base area are investigated for hypersonic flow using both Newtonian impact theory and free molecular flow theory. The drag of Newton’s minimum drag body, which has previously been evaluated by numerical means, is derived as an analytic expression. The analytical results are applicable to a range of local pressure laws allowing minimum drag shapes obtained using impact theory to be directly compared with low density flow equivalents using free molecular flow. The low density shapes are found to have larger blunt regions at the nose and significantly larger drag coefficients. For free molecular flow the drag varies with the surface reflection characteristics. As the fraction of diffuse reflection at the surface increases, the drag increases and the sensitivity of the drag to changes in the minimum drag shape is reduced.

Base-Drag-Reduction Experiments on the X-33 Linear Aerospike SR-71 Flight Program

Drag-reduction tests were conducted on the Linear Aerospike SR-71 experiment. The Linear Aerospike SR-71 experiment e ight tested a 20% scale model of an X-33 forebody with a linear-aerospike engine mounted at the rear of the body. The entire apparatus was mounted on top of an SR-71 aircraft. This paper suggests a method for reducing base drag by adding surface roughness along the forebody. Calculations show a potential for basedrag reductions of 8 ‐14%. Flight results corroborate the base drag reduction, with actual reductions of 15% in the high-subsonic e ight regime. An unexpected result is that drag benee ts persist well into the supersonic e ight regime. Flight results show no overall net drag reduction. The applied roughness caused forebody pressures to rise and offset base-drag reductions. Apparently, the grit displaced streamlines outward, causing forebody e ow compression.Clearlythedrag optimizationmustbemodie edto includenotonly thebasepressuredragandviscous forebody drag coefe cients but must also include the forebody pressure distribution. Because of the mixed experimental results— there was no overall net drag reduction, the existence of an optimal forebody surface roughness must still be proven. Clearly, however, the forebody grit method has been proven as a viable drag-reduction tool.

Comparison of theory and experiment for propfan inlets

Studies have indicated that the use of propfans can result in fuel savings of 20-25% relative to turbofans. Therefore, a program was initiated to perform wind-tunnel tests on a series of inlet models for advanced tractor, turboprop installations. Single-scoop, twin-scoop, and annular inlets were tested to obtain inlet pressure recovery and distortion, external cowl surface pressures, external cowl forebody drag, and propeller blade stress data. An analysis was undertaken for each of the propfan inlet configurations tested using Lockheed's QUADPAN panel program. Calculated nacelle surface pressures correlate well with experiment, except where Mach number is high or viscous effects predominate. The effect of the inlet interference flowfield on the propeller was calculated using Lockheed's PROPVRTX propeller analysis program. The results showed that displacing the inlet in the aft direction reduced the adverse effects on the propeller cyclic stresses. Measured propeller stress data support these results. It is concluded that inlet/cowl designs can be assessed by computational methods.

Bus drag reduction by the trapped vortex concept for a single bus and two buses in tandem

A wind-tunnel investigation of a 1/20 scale model of a modified Denning Mono Mark II bus indicates that the drag coefficient can be reduced to 0.29 by using optimal radii on the forebody/roof junction, 8° tapers behind the forebody and in front of the tail and incorporating a vortex trapping cavity in the tail. Surface pressure distributions and surface-mounted tufts on two forebodies and two tails provide the detailed flow behaviour. For the vortex cavity tail a vortex is well-defined over the top half of the tail but the underbody upwash interferes with the vortex flow over the lower half. The flow field between two in-line buses is determined at various gaps up to 0.472 D ( D = rmhydraulic diameter of the base) for two tails and two forebody configurations. An forebdoy/tail combinations produce a thrust ttail / D, T =2.087 tail is mode iffective in increasing the tail thrust than a flat tail, particularly at larger gaps (⩾ 0.314 D). Either tail, used with a blunt forebody, is effective in reducing the single-bus forebody drag coefficient of C D,F = 0.172 to 0.084–0.095 for gaps ⩽ 0.157 D.

Afterbody Configuration Effects on Model Forebody and Afterbody Drag

An experimental investigation was conducted to obtain additional information necessary to complete a study of the influence of changes in afterbody configuration on the drag of models tested in transonic wind tunnels. The model was a sting-supported axisymmetric body configuration. The model geometry was the same as used in a similar investigation conducted in the 1 x 1-m transonic wind tunnel of DFVLR in Gottingen, Germany. The current investigation was conducted in the AEDC 16-ft Transonic Wind Tunnel at Mach numbers from 0.6 to 1.4 and over a Reynolds number range of (8.12-21.7)xl 06 based on model length. The results of this investigation, which show that forebody drag is not sensitive to changes in afterbody configuration, are in general agreement with previous results obtained at AEDC.

Reply by Authors to F. Aulehla and G. Besigk

T basic purpose of the investigations reported in Ref. 1 and our current paper appears to have been missed by Aulehla and Besigk. It was not to assess the accuracy of their model drag coefficients, since we are aware of the problem in using the trapezoidal method of integration with inadequate pressure orifices. The purpose was to determine if large incremental changes in afterbody drag due to afterbody configuration changes would be compensated by a change in forebody drag as has been stated. If so, the standard wind tunnel testing technique used during the configurational development phase of an aircraft development program would lead to erroneous conclusions. Again, the results presented in Figs. 11-13 of Ref. 1 and Figs. 16-18 of our paper provide a solid basis for the validity of the testing technique. However, as we have always noted, caution must be used in

Forebody Drag Reduction

It is shown that the mechanisms of forebody drag reduction by means of either a spike or a forward-facing jet are similar, with the maximum achievable drag reduction being of the same order. Because the jet may be a relatively cool gas, however, the forward facing jet has the additional capability of reducing the aerodynamic heating that is so severe at high Mach numbers. By means of the correlation presented, jet ejection parameters may be chosen to achieve maximum permissible forebody drag reduction. The correlation method uses a momentum coefficient that characterizes jet efflux and freestream conditions.

Resonant wave interactions on a swept wing

References 'Pugh, P.G. and Ward, L.C., Novel Method for the Estimation of Zero Lift Forebody Drag of Axisymmetric Non-Slender Shapes at Supersonic and Hypersonic Velocities, Aeronautical Research Council, CP 1142, 1971. Oswatitsch, K. and Kuerti, G., Gas Dynamics, Academic Press, New York, 1956, p. 203. Kaattari, G., Method of Predicting Shock Shapes and Pressure Distributions for a Wide Variety of Bodies at Zero Angles of Attack, NASA TN D-4539, 1968. Philpott, D.R., Simple Flow Model for the Calculation of the Flow Round a Axisymmetric Body at Supersonic Speeds, AIAA Paper No. 73-1356, 1978. Albone, C.M., Blunt Body Flows at Supersonic Speeds, City University Research Report 71 /1, 1971. Philpott, D.R., A Program for the Calculation of Axisymmetric Supersonic Flows around Bodies using a Reference Plane Method, The Hatfield Polytechnic, THP/4-73R, 1973. . 7 Engineering Sciences Data Unit (Royal Aeronautical Society), Engineering Sciences Data Item 68021, 1968. Philpott, D.R., An Investigation into the Flow around a Family of Elliptically Nosed Cylinders at Zero Incidence at M=2.50 and M=4.00, Aeronautical Quarterly, Vol. XXIII, Nov. 1972.

Theoretical lower limits of forebody drag

One of the more important objectives of research in aerodynamics is drag reduction. In this paper we consider one part of the drag problem, the forebody pressure drag, and enquire about the lower limit to which it may theoretically be reduced by elimination of flow separation. It is customary to decompose the total drag force into two parts: pressure drag and friction drag, represented by force coefficients CD and CF, respectively. The ratio of these two forces may be taken as a convenient criterion to decide whether a body is ‘streamlined’ or ‘bluff’:At high enough Reynolds numbers many bodies of interest are (according to this definition) bluff bodies, i.e. their pressure drag is much larger than the friction drag. This analysis is most relevant for bluff bodies where one needs to be concerned only with the pressure drag.

Minimum Drag Bodies of Given Length and Base Using Newtonian

T forebody of a high-speed aircraft or missile generates a substantial proportion of the overall drag, such that even a modest reduction in the forebody drag can produce a significant improvement in the performance. The requirement that the forebody blend into a fuselage which is nearly axisymmetric has resulted in an intuitive assumption that any low-drag forebody will also be nearly axisymmetric. Indeed, this assumption is built into most optimization studies by geometric constraints on the forebody. Even when the forebody cross-sections are left unspecified, it is usually assumed that these sections and those of the afterbody are similar, and this approach has led only to unrealistic star cross-section bodies which are discussed in Ref. 1. In this paper, although the base contour is fixed, the cross-sections upstream of the base are merely required to be convex. The minimum drag shapes which emerge using these constraints are termed spatular shapes, because of their flat region near the nose and near circular cross-section downstream. The shapes combine the prospects of significantly lower drag (1015% lower) and improved pilot visibility.

Comment on "Flutter of Flat Finite Element Panels in a Supersonic Potential Flow"

about Mach 3, cone angle and bluntness ratio are equally important; and above Mach 3, bluntness is the relatively stronger parameter in establishing the base pressure level. The validity of the correlation is demonstrated by comparison with specific 9 and 10 deg cone data for sharp and 0.3 bluntness ratio noses which are given in Fig. 4. Data for other cone angles and bluntness ratios exhibit an equally good agreement with the prediction as would be expected based on Fig. 2; however, all of these data are not given in Fig. 4 for clarity. Also, total drag measurements obtained in an aeroballistics range on sharp 10-deg semiangle cones during the course of the transition experiments reported in Ref. 4 were analyzed. A sting free, flight base pressure was extracted from these data by subtracting a calculated forebody drag (pressure and skin friction) from the measured total drag. This calculations can be made with great precision for the sharp slender cone in question since the location of transition was measured. The relatively precise free-flight base pressure results obtained in this manner are given in Fig. 4 and further confirm both the absence of sting effects and the validity of the correlation. Even at Mach numbers in excess of eight, base pressure predictions utilizing this technique are in good qualitative agreement with the available turbulent data for sharp or small bluntness cones (up to Mach ~ 20). Summary At supersonic speeds, the base drag of sharp or blunt conical bodies represents a significant fraction of the total