Ka-band payloads are becoming more and more popular for satellite communication. The wider band width in Ka-band allows a better satisfaction of the increasing demand for capacity. In addition to the use of more resources, a more efficient use of the available resources will become key for a successful development of satellite communication services. Modern antenna concepts allow a high frequency reuse scheme and, therefore, an extreme efficient use of the most rare resource in satellite communication, the frequency band width. In this paper, we describe the design and the use of different types of such antennas.
Design challenges of four basic hypersonic flight vehicle classes are discussed: non-winged (capsules) and winged re-entry vehicles, airbreathing cruise/acceleration and ascent/re-entry vehicles. Basically flight in the Earth atmosphere is considered, no systematic review is given, propulsion systems are not considered. Three topics are treated in detail: (1) The concept of the thermal state of the vehicle surface, which encompasses both thermal surface effects and thermal loads. Thermal surface effects are important especially for hypersonic airbreathing flight vehicles. (2) The aerodynamic performance (L/D) of non-winged and winged re-entry vehicles as important mission parameter. It is sketched, how the shapes of the vehicles evolved and how L/D increases were accomplished. (3) Air-vehicle engineering issues of large airbreathing hypersonic vehicles. High design sensitivities and small payload fractions of such vehicles make necessary new development and test approaches, with numerical multidisciplinary simulation and optimization as well as experimental vehicles playing a deciding role.
Future earth observation satellites call for GEO relay links to make their data immediately available to the user. While RF communication limits the GEO relay’s data rate to roughly 1 Gbps (Gigabit per second) laser communication will extend its capacity into the 10 Gbps range as is required for the future systems. Laser communication will be applied for EDRS, the European data GEO relay system. At first, it is foreseen to provide the RF LEO-to-GEO link with an additional laser communication channel, still limiting, of course, the GEO relay’s performance to the RF bottleneck. However besides the operational service, the EDRS laser terminals shall also demonstrate the performance of high data rate links for both, the inter-satellite link and the GEO-to-ground link.
During their operational life-time, actively cooled liners of cryogenic combustion chambers are known to exhibit a characteristic so-called doghouse deformation, pursued by formation of axial cracks. The present work aims at developing a model that quantitatively accounts for this failure mechanism. High-temperature material behaviour is characterised in a test programme and it is shown that stress relaxation, strain rate dependence, isotropic and kinematic hardening as well as material ageing have to be taken into account in the model formulation. From fracture surface analyses of a thrust chamber it is concluded that the failure mode of the hot wall ligament at the tip of the doghouse is related to ductile rupture. A material model is proposed that captures all stated effects. Basing on the concept of continuum damage mechanics, the model is further extended to incorporate softening effects due to material degradation. The model is assessed on experimental data and quantitative agreement is established for all tests available. A 3D finite element thermo-mechanical analysis is performed on a representative thrust chamber applying the developed material-damage model. The simulation successfully captures the observed accrued thinning of the hot wall and quantitatively reproduces the doghouse deformation.
The interplanetary space probe Cassini/Huygens reached Saturn in July 2004 after 7 years of cruise phase. The German cosmic dust analyser (CDA) was developed under the leadership of the Max Planck Institute for Nuclear Physics in Heidelberg under the support of the DLR e.V. This instrument measures the interplanetary, interstellar and planetary dust in our solar system since 1999 and provided unique discoveries. In 1999, CDA detected interstellar dust in the inner solar system followed by the detection of electrical charges of interplanetary dust grains during the cruise phase between Earth and Jupiter. The instrument determined the composition of interplanetary dust and the nanometre-sized dust streams originating from Jupiter’s moon Io. During the approach to Saturn in 2004, similar streams of submicron grains with speeds in the order of 100 km/s were detected from Saturn’s inner and outer ring system and are released to the interplanetary magnetic field. Since 2004 CDA measured more than one million dust impacts characterising the dust environment of Saturn. The instrument is one of the three experiments which discovered the active ice geysers located at the south pole of Saturn’s moon Enceladus in 2005. Later, a detailed compositional analysis of the water ice grains in Saturn’s E ring system led to the discovery of large reservoirs of liquid water (oceans) below the icy crust of Enceladus. Finally, the determination of the dust-magnetosphere interaction and the discovery of the extended E ring (at least twice as large as predicted) allowed the definition of a dynamical dust model of Saturn’s E ring describing the observed properties. This paper summarizes the discoveries of a 10-year story of success based on reliable measurements with the most advanced dust detector flown in space until today. This paper focuses on cruise results and findings achieved at Saturn with a focus on flux and density measurements. CDA discoveries related to the detailed dust stream dynamics, E ring dynamics, its vertical profile and E ring compositional analysis are published elsewhere (see Hus et al. in AIP Conference Proccedings 1216:510–513, 2010; Hsu et al. in Icarus 206:653–661, 2010; Kempf et al. in Icarus 193:420, 2008; 206(2):446, 2010; Postberg et al. in Icarus 193(2):438, 2008; Nature 459:1098, 2009; Nature, 2011, doi: 10.1038/nature10175 ).
A stress wave internal force balance for the High Enthalpy Shock Tunnel Göttingen (HEG) of the German Aerospace Center (DLR) to measure lift, pitching moment and drag was designed, calibrated and tested. The balance is designed to measure forces in ground based test facilities with test times in the order of milliseconds on models additionally instrumented with surface pressure and wall heat flux gauges from angles of attack of −40° to 20°. Experiments in HEG were performed on a 303 mm long, 10° half angle blunt cone at angles of attack from −20° to 0°. The tests were conducted utilizing two different operating conditions at total specific enthalpies of 3.0 and 3.5 MJ/kg and dynamic pressures of 30 and 72 kPa. The performance of the balance was assessed by comparing the measured force and moment coefficients with computational fluid dynamics (CFD) predictions.
Down-scaled solid propellant motors are a valuable tool in the study of thrust oscillations and the underlying, vortex-shedding-induced, pressure instabilities. These fluctuations, observed in large segmented solid rocket motors such as the Ariane 5 P230, impose a serious constraint on both structure and payload. This paper contains a survey of the numerous configurations tested at ONERA over the last 20 years. Presented are the phenomena searched to reproduce and the successes and failures of the different approaches tried. The results of over 130 experiments have contributed to numerous studies aimed at understanding the complicated physics behind this thorny problem, in order to pave the way to pressure instability reduction measures. Slowly but surely our understanding of what makes large segmented solid boosters exhibit this type of instabilities will lead to realistic modifications that will allow for a reduction of pressure oscillations. A “quieter” launcher will be an important advantage in an ever more competitive market, giving a easier ride to payload and designers alike.
Concurrent engineering (CE) has been in use within the space industry since the mid-1990s for the development of robust, effective design solutions within a reduced period of time; to date, however, such applications have focussed on Phase 0/A feasibility studies, with the potential for application in later phases not yet demonstrated. Applications at the DLR Institute of Space Systems have addressed this gap with practical attempts made on three satellite projects. The use of Phase 0/A CE techniques, such as dedicated CE sessions, online trade-offs, and design iterations and consolidation, was taken and augmented with more novel practices such as online requirements engineering. Underlying these practices was a suite of tools coming from both external and internal sources. While it is noted that the traditional time and cost benefits expected from Phase 0/A use are less likely to be achieved for Phase B applications, the resulting solutions demonstrated an increased robustness and performance.
While Network Coding has been extensively studied from a theoretical point of view, the number of practical implementations is still very limited. We have performed an actual demonstration of this technique in a real satellite network, where Network Coding was applied on the streams of a videoconference between two terminals. The design challenges are analyzed along with the main ideas that enabled the proposed protocol to achieve more than 90% of the theoretical gains.
Emission spectra during re-entry have been measured in 2006 for the STARDUST capsule and in 2008 for the ATV1 “Jules Verne” re-entry. This paper summarizes the approach to design the airborne UV spectroscopic setup and its modifications with respect to the missions. For the STARDUST mission, results of data analysis of data presented in 2008 are given while for the ATV1 observation first spectra of the main disruption are exemplary presented. The surface radiation during the STARDUST re-entry is used to estimate convective and radiative heat flux using different analytical models. A first look at the spectroscopic footprint of ATV1 shows that during the first explosive event, a severe break-up of the main ATV1 structure occurs. However, a correlation with an explosion of fuel could not be observed.
Concerning the requirements of future rocket technologies, providing a cost-efficient access to orbit as well as an increase in system reliability, a deeper insight into the unsteady phenomena during ascent of modern launchers is essential. Unsteady interactions and resonances of the turbulent separated launcher wake and the nozzle structure play an important role for the design of future main stage propulsion systems. The so-called buffeting coupling phenomenon is one of the main challenges during ascent. In the present study, a coupled simulation of the afterbody of the Ariane-5 launcher with a realistic structural and aerodynamic representation of different nozzle configurations is carried out. On the computational fluid dynamics side, unsteady detached eddy simulations are coupled with structural computations for different nozzle configurations. The essential features of the interaction process are well captured. The coupling algorithm is validated by a nonlinear supersonic panel flutter simulation with highly transient mechanical interaction.
In view of future film cooling tests at the Institute for Flight Propulsion (LFA) at Technische Universität München, the Astrium in-house spray combustion CFD tool Rocflam-II was validated against first test data gained from this rocket test bench without film cooling. The subscale rocket combustion chamber uses GOX and kerosene as propellants which are injected through a single double swirl element. Especially the modeling of the double swirl element and the measured wall roughness were adapted on the LFA hardware. Additionally, new liquid kerosene fluid properties were implemented and verified in Rocflam-II. Also the influences of soot deposition and hot gas radiation on the wall heat flux were analytically and numerically estimated. In context of reviewing the implemented evaporation model in Rocflam-II, the binary diffusion coefficient and its pressure dependency were analyzed. Finally simulations have been performed for different load points with Rocflam-II showing a good agreement compared to test data.
In the space business, there is typically a quite long period from the design and development of new technologies to their commercial use. Even comprehensive tests on ground cannot replace long-lasting experiments and tests in space. Such IOV (In Orbit Verification) activities provide the scientific basis for the introduction and application of new technologies and the necessary heritage for commercial satellite programs. The Heinrich-Hertz mission of a geostationary communication satellite with a planned life time of 15 years lead by the German Space Administration (DLR) establishes a valuable basis to verify new technologies scientifically in orbit over a long period of time and to gain heritage regarding their performance in space . In addition, research institutes and the industry are enabled to perform numerous scientific and technological experiments over the full life time of 15 years. With this approach of the mission, the German Space Administration offers to the German satellite industry an outstanding advantage and gain in knowledge for the development of new communication technologies and their applications. The launch of the satellite is envisaged for 2016. The technical feasibility of the overall program was successfully demonstrated within a Phase A study. The major tasks for the payload responsible during Phase A have been: (1) survey and assessment of all proposed IOV-technologies, (2) development of a payload concept for the scientific-technical verification of the IOV-technologies.
The hot fire test strategy for liquid rocket engines has always been a concern of space industry and agency alike because no recognized standard exists. Previous hot fire test plans focused on the verification of performance requirements but did not explicitly include reliability as a dimensioning variable. The stakeholders are, however, concerned about a hot fire test strategy that balances reliability, schedule, and affordability. A multiple criteria test planning model is presented that provides a framework to optimize the hot fire test strategy with respect to stakeholder concerns. The Staged Combustion Rocket Engine Demonstrator, a program of the European Space Agency, is used as example to provide the quantitative answer to the claim that a reduced thrust scale demonstrator is cost beneficial for a subsequent flight engine development. Scalability aspects of major subsystems are considered in the prior information definition inside the Bayesian framework. The model is also applied to assess the impact of an increase of the demonstrated reliability level on schedule and affordability.
Research efforts are currently underway at the German Aerospace Center (DLR) Lampoldshausen, which aim to understand the mechanisms by which self-sustaining oscillations in combustion chamber pressure, known as high frequency combustion instabilities, are driven. Testing has been conducted in the rectangular combustor ‘BKH’, running cryogenic oxygen and hydrogen propellants under pressure and injection conditions which are representative of real rocket engines and with acoustic forcing. For the first time, such tests with LOx/H2 propellants and acoustic forcing have been conducted at combustion chamber pressures above 10 bar, the reported results herein from a test at 42 bar. Optical access to the combustor allowed the application of high speed hydroxyl radical (OH*) chemiluminescence imaging of the flame during periods of forced excitation of acoustic resonance modes of the combustion chamber. This paper reports the investigation of flame response to acoustic excitation. Both fluctuation in OH* emission intensity and deflection of the flame at frequencies corresponding to the excitation frequency have been observed. These responses are then discussed as potential indicators of driving mechanisms for combustion instabilities.
The axisymmetric concave body is a typical configuration about which shock/shock interactions appear. Various shapes of axisymmetric concave bodies are used in a variety of applications in aeronautics, for example, axisymmetric jet inlets with conical centerbody, ballistic missiles drag reduction by spike, plasma or hot gas injection, parachutes for pilot-ejection capsules. However, it is well known that two distinct modes of instability appear around a concave body in the high-speed flow regime for a certain range of geometric parameters. These instabilities can cause undesirable effects such as severe vibration of the structure, heating and pressure loads. According to the experimental evidence, the unsteady flow is characterised by periodic radial inflation and collapse of the conical separation bubble formed around the forebody (pulsation). Various explanations have been given for the driving mechanism of the instabilities. In the present, merging of the leading explanations is done, and basic rules for the passive suppression of the instabilities are applied, in order to enforce their proposed driving. In addition, the effect of the flow initialisation method on the flow structure predicted by numerical simulations is examined. For certain configurations, bifurcation of the time-dependent flow has been found. This behaviour is explained with recourse to the phenomenon of hysteresis, which is an inherent feature of the examined flows.