Afterbody Drag Research Articles

Mitigation of Antisymmetric Mode and Drag With Base Cavities

Abstract Experiments were carried out to examine the impact of base cavities on the base pressure fluctuations and total drag of a cylindrical afterbody for freestream Mach numbers 0.6–1.5. Significant improvement in the base pressure and a substantial reduction in the afterbody drag was noticed in the presence of a base cavity at subsonic Mach numbers. However, on increasing the cavity length beyond a certain value, its performance deteriorates. At supersonic Mach numbers, their effectiveness drops drastically. Tones in the spectra can be classified into two types depending on the dominant azimuthal mode, which is either 0 or 1 and are referred to as symmetric and an antisymmetric mode, respectively. Spectra at subsonic Mach numbers exhibit tones, which are related either to mode 0 or 1. However, at supersonic Mach numbers, only tones related to mode 0 exist. The base cavity either effectively suppresses the antisymmetric mode or modifies it into a symmetric mode resulting in mitigation of the tones related to antisymmetric mode.

Effect of Reynolds number on flow behavior and pressure drag of axisymmetric conical boattails at low speeds

The effect of Reynolds number on flow behaviors and pressure drag around axisymmetric conical boattails was experimentally investigated at low-speed conditions. Four conical boattails with slant angles of 12°, 16°, 20°, and 22° were studied. The Reynolds number ranged from 4.34 × 104 to 8.89 × 104 based on the model diameter. The global-luminescent-oil-film skin-friction measurement was employed to analyze the surface skin-friction topology. Quantitative skin-friction values at the centerline were obtained in this study. The results show that a separation bubble can be formed on boattail surfaces at angles from 12° to 20°. However, at a boattail angle of 22°, flow is fully separated near the boattail shoulder. The integrated afterbody pressure drag indicated that, at angles of 12°, 16°, and 22°, the Reynolds number has very small effect on the afterbody drag, while, at 20° the drag coefficient decrease was relatively large with increasing Reynolds number. We believe that this study provided the first results for a boattail angle of 20° and we observed that the size of the separation bubble decreased as the Reynolds number increased. The effect of the separation bubble on the pressure distribution was also examined in detail.

Aerodynamic assessment of nozzle area variation by core fairing modulation

This paper describes the aerodynamic assessment of a particular variable area fan nozzle (VAFN) concept. This concept offers a variation in the fan nozzle throat area by displacing the inner fairing structure (IFS), which is also known as the core fairing structure. A few different design configurations were developed through the variation in the design parameters. The aim of the assessment was to perform a comparative aerodynamic study by evaluating the main performance parameters such as the nozzle thrust coefficient, nozzle discharge coefficient and the after-body drag at different flight conditions. The required boundary conditions and the magnitudes of the variation in the nozzle throat area were defined following the performance calculations of the whole engine.

Correlation for the Estimation of Afterbody Drag with Hot Jet Exhaust

MODERN turbojet and turbofan engines of combat aircraft operating over a wide range of power settings experience jet exhaust temperature typically varying from 1000-2000 K, whereas much of afterbody-nozzle testing is conducted with a cold jet near 300 K.1-5 Thus there remains a problem to determine the extent to which jet total temperature (and its associated gas constants) affects the afterbody drag of a combat aircraft under various operating conditions of its nozzle during the flight operations-$ Physical modeling of jet freestream interactions with temperature effects is quite difficult_ and, calculations of afterbody drag with hot jet exhaust are computationally intensive. Efforts made earlier for the estimation of afterbody drag with hot jet exhaust had limited success. In the present analysis, a simple correlation is proposed that can be used in the subsonic and transonic Mach number range for the estimation of afterbody pressure drag with jet temperature effects.

Zero-Lift Drag Characteristics of Afterbodies with a Square Base

Results of zero-lift drag characteristics of afterbodies with a square base relevant to missile and projectile applications at several transonic Mach numbers and Mach 2 are presented. These afterbodies, at zero incidence, generate vortical flows upstream of the base like that on the lee side of a delta wing at incidence. The measurements made consisted primarily of afterbody drag using a balance and base pressure and extensive surface flow visualization studies were carried out to infer features associated with vortex flows. Results of base pressure, boat-tail profile drag, and total afterbody drag are presented and compared with results from the axisymmetric counterpart involving circular arc and conical boat tailing having the same base area. Some aspects of the flow features on square-base afterbodies are discussed as well.

Overexpanded Two-Dimensional-Convergent-Divergent Nozzle Flow Simulations, Assessment of Turbulence Models

HERE is renewed interest in two-dimensional‐ convergent-divergent (2D‐CD) nozzles because of the advantages they offer over axisymmetric cone gurations for supersonic transport. These include higher performance, reduced afterbody drag, easier integration with airframes, and large mechanical area excursion capabilities. The performance of these nozzles can suffer at off-design conditions when the large nozzle area ratio reductions required during subsonic and transonic acceleration cannot be achieved under the mechanical and control system constraints. Because this can adversely affect the acceleration time to cruise, the fuel burnt, and the range, it is desirable to predict off-design performance with a high degree of accuracy. The purpose of the present investigation is to assess the computational results obtained for overexpanded 2D‐ CD nozzles using e ve different turbulence models. Overexpanded nozzle e ow predictions are challenging because of the large shock-induced separated e ow regions in the divergent nozzle section. The 2D‐ CD nozzle tested under static conditions by Hunter 1 was selected for the numerical assessment because the large number of pressure taps gave a better dee nition of the shock location. The nozzle has a design pressure ratio of 8.8 for an exit Mach number of 2.1. Approach The implicit numerical solution of the compressible twodimensional Navier‐ Stokes equations was obtained using the NPARC code.

Supersonic near-wake afterbody boattailing effects on axisymmetric bodies

An experimental investigation of the near-wake flowfield downstream of a conical boattailed afterbody in supersonic flow is presented. The afterbody investigated is typical of those for conventional boattailed missiles and projectiles in unpowered flight. Flow visualization, mean static pressure measurements, and three-component laser Doppler velocimeter data have been obtained throughout the near wake of the body. The effects of afterbody boattailing on the physics of the near-wake flow are determined by comparing the present data with similar data obtained on a cylindrical afterbody. Results indicate that a net afterbody drag reduction of 21 % is achieved with the current boattailed afterbody for a freestream Mach number of 2.46. The shear-layer growth rate, and therefore mass entrainment from the recirculation region behind the base, is shown to be significantly reduced by afterbody boattailing due to the reduction in turbulence levels throughout the near wake as compared to the cylindrical afterbody.

Necessity of an end face for the optimum two-dimensional afterbody with a boundary layer

UDC 533.6.011.5:532.526 Many engineering components (struts, airfoils, turbomachinery blades, jet engine exhausts, etc.) have aftersections (tail sections) passed over by viscous flow. Of course, as far as possible they should have minimum drag. Usually it is assumed that if the aspect ratio is high enough these optimum aftersections should not have end faces. Theoretical research [1--3] has shown that for relatively small aspect ratios the optimum aftersection of a two-dimensional body in a supersonic inviscid gas flow may include an end face, behind which the flow separates. With increase in the length of the afterbody the height of the end face decreases and when a certain length is attained it becomes equal to zero, i.e., there is no flow separation. On the other hand, there are experimental data which show that even for relatively large aspect ratios the optimum afterbody has an end face. As far as the author is aware, this decrease in afterbody drag following the introduction of an end face was first demonstrated by Zhdanov in 1959 on the basis of an experimental investigation of an axisymmetric model of a jet engine exhaust. For a given exhaust length the afterbody contour was varied by introducing an end face. A parametric investigation made it possible to find the optimum height of the annular end face ensuring minimum afterbody drag and, consequently, maximum thrust. This effect was obtained for both supersonic and subsonic external-flow velocity. Similar data for an axisymmetric exhaust at supersonic external-flow velocities were obtained in [4], where the total afterbody drag Cxr " = Cxp + Cxb was found as a function of the angle of inclination of the wall. Here, Cxp is the pressure drag coefficient of the lateral surface of the afterbody, and Cx2 , is the base drag coefficient. It was established that for each value of the internal- and external-flow pressure ratio (pc/pn = 1--15) there is an optimum height of the annular end face for which the total afterbody drag reaches a minimum. A recently published book [5] provides the, at ftrst glance, unusual information that shortening the aftersections of the wing and fen profiles of the Concorde by introducing an end face reduced the drag. This would appear to contradict the conclusion to be drawn from the theory of base flows: since when the wing profile is shortened the angle of inclination of the wall ahead of the end face and the ratio of the boundary layer thickness to the height of the end face ~/h decrease, it follows from the theory that the base pressure should decrease and the drag increase. Similar results were obtained in [6] on the basis of an experimental investigation of subsonic and transonic flow past an axisymmetric body with a pointed rear end. The investigations were carried out on bodies with various degrees of shortening resulting from the introduction of an end face. The results also show that up to a certain size the introduction of an end face does not increase the total drag. These experimental data indicate the existence of an effect which was not taken into account in [2, 3]. In those studies the base pressure behind an end face of height h was determined without taking into account the effect of the initial boundary layer of thickness/~, which is valid only for 5/h 1.

Study of External Dynamic Flap Loads on a 6 Percent B-1B Model

The origin of dynamic pressure loads on external divergent engine nozzle flaps of the B-1B aircraft was investigated in the NASA/LaRC 16-ft transonic tunnel using a 6 percent full-span model with powered engine nacelles. External flap dynamic loads and afterbody drag associated with flap removal were measured using this model. Both dry and maximum A/B power nozzles were evaluated in this study. As a result of this study, the principle mechanisms responsible for high dynamic external flap loads were determined along with performance penalty associated with flap removal.

Sting corrections to zero-lift drag of axisymmetric bodies in transonic flow

AbstractExperiments at transonic speeds have been performed on several boat-tailed afterbodies and sting combinations with a view to assessing sting corrections to the measured afterbody drag at transonic speeds. Measurements made included afterbody total drag and base pressure in the Mach number range of 0.6-1.0 and Reynolds number range of 8-9.5 x 106. Correlations of base pressure and boat-tail pressure drag for the sting diameter and flare effects have been proposed using dimensional arguments. The correlations provide quick and reliable estimates for corrections that can be applied to the measured zero-lift drag of axisymmetric bodies with either contoured or conical boat-tailing.

Effectiveness of passive devices for axisymmetric base drag reduction at Mach 2

Experiments have been made to assess the effectiveness of passive devices for reducing base and total afterbody drag at Mach 2. The devices examined include primarily base cavities and ventilated cavities. Results show that ventilated cavities offer significant base-drag reduction and net-drag reduction of engineering value. A correlation of base- and net-drag reduction for ventilated cavities has been suggested. Typical effects of these devices on boat-tailed and flared bases have also been assessed. ^DB -pb D h L M #00 t Nomenclature = forebody (max) cross-sectional area, 490.625 mm2 = total area of ventilation = total afterbody drag coefficient, net-drag force/to^ A) = base-drag coefficient = base-pressure coefficient = forebody (max) diam, 25 mm = cavity depth = afterbody length = freestream Mach number = freestream dynamic pressure = cavity lip thickness

Underexpanded jet-freestream interactions on an axisymmetric afterbody configuration

Experimental investigations are carried out in the 0.3 m trisonic wind tunnel at NAL to study exhaust jet-freestream interactions on an axisymmetric afterbody configuration. The model chosen for these studies is an equivalent body of revolution of a representative combat aircraft configuration. The effects of jet pressure ratio and freestream Mach number on boattail, base, and afterbody pressure drag are studied along with color Schlieren flow visualization of the afterbody flowfield. Experiments are carried out in the freestream Mach number range of 0.6 to 1.1 at Reynolds number from 20 amp;times; 10lt;supgt;6lt;/supgt;/m to 35 amp;times; 10lt;supgt;6lt;/supgt;/m keeping model at zero degree of incidence. Jet pressure ratio is varied from 1 to 6 during this investigation. The jet at nozzle exit is sonic throughout these tests.

Methodology for analysis of afterbodies for three-dimensional aircraft configurations

Today's modern fighter design, with multifunction nozzles, is placing an increased emphasis on nozzle/airframe integration. The current tools available to the aircraft designer for aft-end design and evaluation are model test reports, being disseminated mainly by government laboratories, and three-dimensional numerical computation codes. Test data utilization usually is limited by the suitability of the area that has been tested. The second approach, analysis, usually requires time-consuming three-dimensional configuration data input. Recognizing the need for a quicker means of solution, useful in a preliminary design environment, a semiempirical computer methodology for determining three-dimensional aircraft afterbody performance has been developed. The essence of the approach is to construct equivalent bodies of revolution of three-dimensi onal bodies and then to utilize a straight or hybrid axisymmetric analysis. This approach has been developed for single- and twin-engine axisymmetric and two-dimensional afterbodies. The methodology has been verified by comparisons of afterbody drag and axial and longitudinal pressure distributions.

Passive devices for axisymmetric base drag reduction at transonic speeds

Experiments have been made to assess the effectiveness of several base modifications or passive devices for reducing base drag and total afterbody drag at transonic speeds. The modifications tested include base cavities, ventilated cavities, and two vortex suppression devices. Results show that, while appreciable base drag reductions are possible with many of the devices examined, the net total drag reductions are relatively lower, presumably because of the additional losses associated with the devices. (Author)

Comparison of computational and experimental jet effects

Computations were made for transonic flow over a nozzle afterbody with a real plume. The effects of parameters such as nozzle pressure ratio, plume stagnation temperature, and interior nozzle shape on afterbody drag were quantified and compared with available experimental data. Remarkable agreement between the computational and experimental results is demonstrated, including the drag minimum associated with low nozzle pressure ratios and the absolute dependence of drag on temperature. The computer program used for the calculations is an axisymmetric modification of Deiwert's planar Navier-Stokes program and utilizes a mixed explicit-implicit MacCormack algorithm with an algebraic turbulence model.

Comment on 'Afterbody configuration effects on model forebody and afterbody drag'

I the above paper Thompson and Smith show an appreciable difference between pressure drag values obtained from AEDC and pressure measurements (see Fig. 1). As the basic data, i.e., the measurements of the pressure distributions, were obtained at ten years ago in a small wind tunnel and also with much smaller effort in instrumentation, the unprejudiced reader will certainly assume the data to be incorrect. However, what Thompson and Smith present as drag data is in fact a not sufficiently accurate integration of the original pressures. From their reported need to use for the integration also the stagnation pressure at the nose of the body and from the discrepancy between the two drag distribution curves it is clear that they have conducted a trapezoidal integration of Cp=f(R) instead of RCP=f(R), with R being the radius of the body varying with x. Particularly, if there are so few pressure orifices in the nose region, as was the case with our model, the first method used by Thompson and Smith will result in a too large forebody pressure drag. This is because the errors due to the trapezoidal approximation in the regions of overand underpressure, respectively, add to each other, while with the second integration method used by us they are partly compensated, as shown in Fig. 2. Note that these two methods of approximating one single set of tabulated data (CP and R) result in very different forebody pressure drag coefficients, the magnitude of which differs by more than a factor of 5 if only the actually measured pressures are used. The drag value according to the less accurate method No. 1 yields a forebody pressure drag coefficient of +0.0162, which agrees perfectly with the DFVLR forebody pressure drag given by Thompson and Smith (Fig. 1). The above two integrations were performed in November 1981, after we had seen the original version of the Thompson and Smith paper. The purpose was to understand their integration of the data. In our earlier data reduction in 1972 and 1974 we had further improved method No. 2 by obtaining additional CP values for the polynomial approximation through interpolation of the CP=f(x) values. Depending on the number of interpolated CP's, the above forebody pressure drag coefficient is raised from -0.003 to 0.001 or 0.004. Our 1972 data, which have been added in Fig. 1 to the results presented by Thompson and Smith, are found to agree very well with their data, both forebody and afterbody drag. Similar statements hold also for afterbody No. 5.

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

GAU-8 Projectile Afterbody Drag Reduction by Boattailing and Heated Gas Injection

An experimental investigation was performed on a GAU-8 projectile model in order to determine whether base injection of different gases at a temperature of up to 900°R, which simulates the effects of fumers, would produce a reduction in the afterbody drag. Wind tunnel tests were performed at Mach 3, using different boattail angles and two different gases—air and helium. Pressure data were taken together with schlieren photographs illustrating the flowfield. Results show that afterbody drag reduction of 30-40% is obtained by small gas injection. Afterbody boattail angles of 4-7 deg additionally reduce the afterbody drag by about 15%. Heating the injected gas to 700°R does not further reduce the afterbody drag, but increasing its temperature up to 900°R produces an afterbody drag reduction of approximately 15%. For the same mass injection rates, helium injection is more efficient than air.

Mass Injection and Jet Flow Simulation Effects on Transonic Afterbody Drag

Abstract : An experimental investigation has been performed to determine the effects of boattail injection and jet flow simulation on the afterbody drag of a slender body of revolution in the transonic regime at zero angle of attack, such as engine nacelles and boattailed afterbodies with isolated engines. A correlation between sting and jet diameter has been established. The jet plume and the nozzle pressure ratio simulators have been found appropriate and useful as a testing technique. Boattail mass injection usually produces a drag coefficient reduction and is more effective at high nozzle pressure ratios. Boattail injection is more effective if used in regions of separated flow. (Author)

Prediction and Measurement of Propulsion System Performance

The problem of predicting installed propulsion system performance for a specific configuration from isolated inlet, nozzle and airframe data, and from data for similar configurations is discussed. The degree to which element performance may be isolated from measurements made on an integrated propulsion system will be discussed. An approach to evaluation of external performance of inlets and nozzles is presented. The design of parametric tests and some correlations for afterbody pressure drag from such tests are described. Improvement of data quality from model tests with sophisticated simulation of propulsive flows by pretest studies and in-test quality control are proposed.