Reusable Launch of Small Satellites Using Scramjets

Preller, Dawid; Smart, Michael K.

doi:10.2514/1.a33610

Cited by 37 publications

(77 citation statements)

References 22 publications

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“…An example of the type of reduced modelling applied in MDO is the use of panel methods that rely on analytic solutions 2.3. EXISTING DESIGN METHODOLOGY to the shock-expansion equations and boundary layer correction techniques to evaluate aerodynamic forces [10,15,16,42]. Figure 2.2(a) presents a flow diagram showing the interactions between a partial set of analysis codes utilised in Boeing's hypersonic vehicle MDO.…”

Section: Existing Design Methodologymentioning

confidence: 99%

“…Other hypersonic vehicle MDO applications published in the literature have employed low-fidelity physics modelling coupled with a Simplex method [10,15,16,42]. In these cited studies, the vehicle geometries were also parameterised using on the order of 10 design parameters.…”

Section: Existing Design Methodologymentioning

confidence: 99%

“…In an effort to make MDO more practical, it is common to apply computationally inexpensive, low-fidelity methods to model the complex physics. For example, the use of panel methods based on shock-expansion theory to evaluate vehicle aerodynamic forces have been applied in several recent MDO of hypersonic vehicles [10,[15][16][17][18]. There is some concern that the reduced fidelity modelling does not adequately capture the complex flow phenomena that form around a hypersonic vehicle.…”

Section: Hypersonic Vehicle Design Challengesmentioning

confidence: 99%

“…This inviscid result took approximately 5 hours to converge on 24 compute cores. If this vehicle was parameterised by 10 design variables, as done in the work by Jazra et al [16,17] and Preller et al [10,18], then one design iteration would take approximately 50 hours using finite differences. This is too prohibitive for vehicle design.…”

Section: Modern Cfd-based Design Optimisationmentioning

confidence: 99%

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Adjoint-based aerodynamic design optimisation in hypersonic flow

Damm¹

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The drive for improved flexibility and reusability in satellite launch systems has brought a resurgence in the popularity of scramjet-powered access-to-space launch vehicle concepts. A challenge in scramjet-powered accelerator vehicles is achieving positive thrust margins at the high-speed end of their flight envelopes. As a consequence, scramjet-powered vehicles typically rely on highly integrated airframe-engine configurations. However, due to the tight integration, the geometries and the flow physics are complex. An automated optimisation method seems an excellent candidate approach to explore the complex design space of hypersonic vehicles. This thesis focuses on the development and application of a particular optimisation method, adjoint-based optimisation, to aid in efficient aerodynamic design in hypersonic flows. Shape optimisation of scramjet-powered vehicles requires many design parameters to capture the geometric detail, and, since the flow physics is complex, the fidelity of Reynolds-Averaged Navier-Stokes (RANS) analyses is desirable. The combination of many design parameters and an expensive objective function evaluation drives the need for gradient evaluation methods that scale well with the number of design variables. An advantage of the adjoint method is that all shape sensitivities for an objective function are evaluated at the cost of only one flow solution and one adjoint solution. As a result of its efficiency, the adjoint method has become widely used for aircraft design, evolving to the design optimisation of full configurations. Despite the wide use of adjoint methods in aircraft optimisation, there has been very little application to hypersonic vehicle design. Several high-speed adjoint solvers have been reported in the literature, however, a majority of these works have only verified the adjoint sensitivities, few have followed on to demonstrate the method for use in design optimisation. The contribution of this work is the description and application of a discrete adjoint solver in high-speed compressible flow optimisation. What is unique in this work is a demonstration that complex-step differentiation works well to linearise a second-order spatially accurate unstructured RANS solver. In particular, the approach presented in this work utilises the k − ω turbulence model in high-speed ducted flow configurations. As part of this work, to provide flow analysis with a rapid turnaround , an unstructured steadystate RANS solver driven by a Jacobian-Free Newton-Krylov method was developed. Turbulence is modelled using the two-equation k − ω turbulence model. The Newton method is globalised by using the pseudo-transient approach. A restarted GMRES method is used to solve the system of linear equations arising when solving for the Newton steps. Evaluation of the matrix-vector products required in the GMRES algorithm is accomplished by Fréchet derivatives using imaginary perturbations in the complex plane. This is necessary to achieve robust convergence of the types of turbulent hypersonic flows...

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Section: Existing Design Methodologymentioning

confidence: 99%

Section: Existing Design Methodologymentioning

confidence: 99%

Section: Hypersonic Vehicle Design Challengesmentioning

confidence: 99%

Section: Modern Cfd-based Design Optimisationmentioning

confidence: 99%

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Adjoint-based aerodynamic design optimisation in hypersonic flow

Damm¹

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“…The REST inlet design methodology of Smart (1999) was tailored to a Mach 12 engine, with this variant's design documented by , before being refined through a series of studies (Barth et al, 2015a;Barth, 2014;Doherty et al, 2015;Wise and Smart, 2015;Landsberg et al, 2018a). The engine is a candidate for an access-to-space, accelerating engine when integrated to a hybrid rocket-scramjet-rocket launch vehicle (Preller and Smarr, 2017). However, only the Mach 8 version of this engine has been the subject of proposed flight experiments , and no flight data for the high Mach number engine exists, nor have any numerical studies examined its steady-state flight operation.…”

Section: Introductionmentioning

confidence: 99%

Improving performance of high Mach number scramjets: fuelling strategies and combustor design

Landsberg¹

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With prospective flight speeds exceeding Mach 12, airbreathing scramjets, incorporated in hybrid satellite launch vehicles, offer efficiency and operability benefits when compared to rocket-based systems. By capturing atmospheric air, the scramjet second stage negates the requirement for onboard oxidisers. Addressing the small satellite niche market, these hybrid systems provide a cost-effective solution to launching such payloads to orbit. At hypersonic speeds, however, conventional wing-mounted engine nacelles are not viable; the scramjet must be integrated to the airframe. Hence, the flight-candidate, Mach 12 Rectangular-to-Elliptical Shape-Transitioning (M12REST) scramjet blends a rectangular capture area (permitting parallel mounting of engine modules to the planar vehicle underside), with a structurally efficient elliptical combustor. As access-to-space scramjets suffer immense heating loads, and engine airflow residence times approach air-fuel reaction timescales, hydrogen fuel is utilised for its exceptional cooling capacity and rapid ignition characteristics. However, efficient fuelling techniques are required to avoid crippling the engine's performance through conservative combustor lengths. This thesis numerically and experimentally investigated whether the performance of such an airframe-integrated scramjet can be sufficiently enhanced through customisation of fuel injection and combustor geometry to achieve net thrust at Mach 12. Three-dimensional, chemically reacting Reynolds-averaged Navier-Stokes solutions enabled numerical analysis, while experimental validation was performed within The University of Queensland's, T4 Stalker Tube.Preliminary work investigated cascaded fuel injectors, targeting mixing and penetration improvements. Two streamwise-aligned jets were employed, with the upstream injector half the diameter of the rear, while the distance between each was varied. The upstream jet induced a low dynamic pressure region in its wake, shielding the downstream jet from the hypersonic crossflow. The leeward jet benefits from an increased local jet-to-freestream momentum ratio, while its larger diameter increases absolute penetration. Examining performance at M12REST mean combustor entrance conditions, unique optimal injector spacings existed for each performance metric, displaying improved penetration, spread, and mixing across the range of flight Mach numbers examined (6 ≤ M ≤ 12). However, jet-to-jet spacings of 4-6 total jet diameters displayed universal performance enhancements over single fuel jets. With penetration improvements of 30-40%, and mixing improvements of 40-70%, the simple fixed geometry, passive technique is ideal for use within an accelerating access-to-space scramjet.While cascaded injectors improved performance within uniform flows, the flows ingested by airframe-integrated scramjets are far less homogeneous. Non-uniform compression fields combine with thick boundary layers developed over the vehicle forebody to deliver density stratified flow to the combustor. This thesis ...

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Performance of high mach number scramjets - Tunnel vs flight

et al. 2018

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While typically analysed through ground-based impulse facilities, scramjets experience significant heating loads in flight, raising engine wall temperatures and the fuel used to cool them beyond standard laboratory conditions. Hence, the present work numerically compares an access-to-space scramjet's performance at both these conditions. The Mach 12 Rectangular-to-Elliptical Shape-Transitioning scramjet flow path is examined via three-dimensional and chemically reacting Reynolds-averaged Navier-Stokes solutions. Flight operation is modelled through 800 K and 1800 K inlet and combustor walls respectively, while fuel is injected at both inlet-and combustor-based stations at 1000 K stagnation temperature. Room temperature walls and fuel plena model shock tunnel conditions. Mixing and combustion performance indicates that while flight conditions promote rapid mixing, high combustor temperatures inhibit the completion of reaction pathways, with reactant dissociation reducing chemical heat release by 16%. However, the heated walls in flight ensured 28% less energy was absorbed by the walls. While inlet fuel injection promotes robust burning of combustor-injected fuel, premature ignition upon the inlet in flight suggests these injectors should be moved further downstream. Coupled with counteracting differences in heat release and loss to the walls, the optimal engine design for flight may differ considerably from that which gives the best performance in the tunnel.

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Reusable Launch of Small Satellites Using Scramjets

Cited by 37 publications

References 22 publications

Adjoint-based aerodynamic design optimisation in hypersonic flow

Adjoint-based aerodynamic design optimisation in hypersonic flow

Improving performance of high Mach number scramjets: fuelling strategies and combustor design

Performance of high mach number scramjets - Tunnel vs flight

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