ALBERT

Beschreibung: The new genus Lianus of subfamily Doryctinae (Hymenoptera: Braconidae) is described and illustrated. The differences from other genera of Doryctinae are discussed. Both included species originate from the Atlantic forest at Campos do Jordão, São Paulo State, Brazil.

Beschreibung: The Pterophoridae (Ochyroticinae, Deuterocopinae, Pterophorinae (Tribus: Platyptiliini, Exelastini, Oxyptilini)) species of the Neotropical fauna are reviewed. The species are redescribed. Moths are illustrated in colour for the first time, their genitalia are illustrated in line drawings. The examination of type specimens revealed the presence of four new synonyms: Platyptilia juanvinas Gielis, 1999, is a junior synonym of Platyptilia gravior Meyrick, 1932; Platyptilia jonesi Gielis, 1996, is a junior synonym of Platyptilia semnopis Meyrick, 1931; Oxyptilus maleficus Meyrick, 1926, is a junior synonym of Leptodeuterocopus neales (Walsingham, 1915); Lioptilus parvus Walsingham, 1880, is a junior synonym of Lioptilodes albistriolatus (Zeller, 1871). During the study of much new material 40 new species were discovered: Leptodeuterocopus tungurahue, L. angulatus, L. panamaensis, L. duchicela, Sochchora mulinus, Melanoptilia nigra, Platyptilia spicula, Stenoptilodes maculatus, S. agricultura, S. heppneri, S. medius, S. altiaustralis, Postplatyptilia nebuloarbustum, P. antillae, P. caribica, P. uruguayensis, P. pluvia, P. seitetazas, P. transversus, P. carchi, P. boletus, P, ugartei, P. drechseli, P. corticis, P. zongoensis, P. vorbecki, Amblyptilia landryi, A. kosteri, Lioptilodes altivolans, L. arequipa, L. yungas, L. salarius, L. cocodrilo, Michaelophorus margaritae, M. bahiaensis, Geina integumentum, Capperia browni, Buckleria brasilia, Megalorhipida paraiso, M. dubiosa. In the larger genera the species are arranged in groups. For a distinct group of species a new genus: Melanoptilia is proposed. A comprehensive checklist of the species has been made. To facilitate identification a key is presented to the genus level.

Beschreibung: A bibliography of the family Braconidae/Hymenoptera: Ichneumonidae is given for the period 1964-2003. It is an addition to Shenefelt's bibliography (1965), which covers the period 1785-1963. In total 10,436 references are listed.

Beschreibung: The European species of the genus Helorus Latreille, 1802, are keyed and illustrated, with special reference to their distribution in The Netherlands. Three of the four European species are newly reported for The Netherlands. The description of a new species from Sulawesi (Indonesia) is added: H. celebensis spec. nov.

Beschreibung: Two new genera of Entedoninae are described from Afrotropical region. The genus Janicharis gen. nov. (type species: J. africana spec. nov.) is described from Cameroon, Nigeria and Madagascar. The genus Hakuna gen. nov. (type species: H. matata spec. nov.) is described from Uganda. Both genera have a rather characteristic habitus and a peculiar propodeum bearing large anterolateral strips. A new, but yet unnamed, species of the genus Trisecodes Delvare & LaSalle, 2000, is recorded from Cameroon. This is the fi rst Afrotropical record of this genus, originally described from the Neotropical region.

Beschreibung: In total 32 new species of Tenthredinidae are described from Vietnam and 18 new species from Indonesia: Nesoselandria albotonkinensis spec. nov., N. albeotegularissima spec. nov., N. devriesiana spec. nov., Anapeptamena achterbergiana spec. nov., Neostromboceros nigrogiganteus spec. nov., N. gracilioides spec. nov., N. alboclypeatus spec. nov., N. daoensis spec. nov., N. phuongensis spec. nov., Abusarbidea bicoloristigmata spec. nov., Eusunoxa alboapicalis spec. nov., Neopoppia tonkinometallica spec. nov., Abeleses vietnamensis spec. nov., A. metallotonkinensis spec. nov., Heptamelus tonkinensis spec. nov., Athlophorus devriesi spec. nov., Rhopographus vietnamensis spec. nov., Xenapatidea hematothoracica spec. nov., X. devriesi spec. nov., Allantidea achterbergiana spec. nov., Neothrinax nigrotonkinensis spec. nov., Indotaxonus flavissimus spec. nov., Darjilingia vietnamensis spec. nov., D. tonkinensis spec. nov., D. hoangliensis spec. nov., Ferna achterbergi spec. nov., Monophadnus bicoloritonkinensis spec. nov., Eutomostethus phongdiensis spec. nov., Eutomostethus albotegularissimus spec. nov., Tenthredo octomaculatus spec. nov., T. tricoloritonkinensis spec. nov. and T. danangiensis spec. nov. from Vietnam. Neostromboceros pleuronotatus spec. nov., N. albopedissimus spec. nov., N. flavopedis spec. nov., N. rubroguinealis spec. nov., N. rubromalayensis spec. nov., Eusunoxa malaya spec. nov., Emphytus danumiensis spec. nov., Hemibeleses sulawesiensis spec. nov., Neopoppia irregulata spec. nov., Tenthredo nigrosabahensis spec. nov., T. djampangensis spec. nov., Neothrinax kaindiensis spec. nov., N. excavata spec. nov., N. gedehensis spec. nov., Abeleses metallojavanus spec. nov., Caliroa nigrojavana spec. nov., Nesoselandria sulawesiensis spec. nov. and Athlophorus achterbergianus spec. nov. from Indonesia. Monophadnus glaucus Enderlein, 1919; Monophadnus sumatranus Enderlein, 1919; Amonophadnus nigripennis Benson, 1935; Amonophadnus nigrojavanus Haris, 2002; Corporaalinus azureus Forsius, 1925; Corporaalinus cyanescens Forsius, 1925; and Corporaalinus jacobsoni Forsius, 1929, are proposed as new synonyms of Monophadnus rivalis Konow, 1906. An identification key is added for the known Tenthredinidae of Borneo, Sumatra, Java, Sulawesi and New Guinea.

Beschreibung: Two new host records for Exasticolus fuscicornis (Cameron, 1887)(Hymenoptera: Braconidae: Homolobinae)are presented with detailed larvae morphology and other biological information.

Beschreibung: Thirty species belonging to 16 genera of the family Braconidae (Hymenoptera) are reported from Greenland. Seven are new species described and illustrated below: Dacnusa groenlandica spec. nov.; Aphidius tarsalis spec. nov.; Praon brevistigma spec. nov.; Blacus (B.) groenlandicus spec. nov.; Cotesia crassifemorata spec. nov.; C. fascifemorata spec. nov. and Microplitis lugubroides spec. nov.

Beschreibung: The little-known monotypic genus Paracyphocrania Redtenbacher, 1908 (Phasmatinae: Phasmatini) is reviewed and now comprises two species: P. lativentris Redtenbacher, 1908, and P. tecticollis (Redtenbacher, 1908) comb. nov., which are redescribed and illustrated. A neotype is designated for P. lativentris which is newly recorded from Sulawesi.

Beschreibung: Vestalaria vinnula spec.nov. (holotype male, southern Vietnam, Lam Dong province, Blao, 1962) is described in both sexes and compared with other species of Vestalaria May, 1935 (= the Vestalis smaragdina - group), which is ranked as valid genus.

Beschreibung: A new species of the genus Artocella van Achterberg is described from Spain. On the basis of its substantial sexual dimorphisim, the differences between the two previously known species of this Turanian-Mediterranean genus, which had each been described from specimens of only one (and differing) sex, are reassessed.

Beschreibung: Dasypoda intermedia spec. nov. from Iran is described. Its description fills a gap of our knowledge of the East Mediterranean fauna of the genus Dasypoda. The West Mediterranean Dasypoda species are well known but the eastern species lack convincing records. Moreover, D. intermedia spec. nov. is a very interesting species from a phylogenetic point of view. It shares some characters common to subgenera Dasypoda s. str. and Megadasypoda Michez, 2004, which provide further evidence for the close relationship of both subgenera.

Beschreibung: Thirty-one species of the family Platystictidae of the Philippines are revised, i.e. all species recognised, excluding the species of the Drepanosticta halterata-group. The following new taxa are described: 16 species in Drepanosticta Laidlaw: D. acuta spec. nov., D. aurita spec. nov., D. centrosaurus spec. nov., D. clados spec. nov., D. flavomaculata spec. nov., D. furcata spec. nov., D. hermes spec. nov., D. krios spec. nov., D. luzonica spec. nov., D. malleus spec. nov., D. myzouris spec. nov., D. paruatia spec. nov., D. pistor spec. nov., D. quadricornu spec. nov., D. rhamphis spec. nov., D. trachelocele spec. nov., two in Protosticta Selys, viz. P. lepteca spec. nov. and P. plicata spec. nov., and three in Sulcosticta gen. nov., viz. S. striata spec. nov., S. pallida spec. nov. and S. viticula spec. nov. The status of eleven previously described nominal taxa is established. One, D. septima Needham & Gyger, is doubtfully considered a synonym of D. mylitta Cowley. Based on a preliminary phylogenetic analysis, the species of Drepanosticta are divided into informal species groups. Most species of the Philippines have affinities to species of Sulawesi, the Moluccas and New Guinea. Several species confined to Palawan have sister-group relationships with species from Borneo. The affinities of various other species confined to the Sulu archipelago, are unsettled as yet. The species of Platystictidae here assigned to Protosticta Selys are presumably not closely related to the type species, P. simplicinervis Selys from Sulawesi. However, a better placement has to await a more detailed phylogenetic study of the family. For three species the new genus Sulcosticta gen. nov. is erected. These species are closely allied based on the structure of the appendages, but should have been assigned to different genera if based on the present generic definitions. Many species here described have small distributional ranges, a common phenomenon in Platystictidae. Since most forests in the Philippines are heavily under threat or have already disappeared in the last fifty years, several taxa described in this paper should be considered under threat of immediate extinction.

Beschreibung: In this part the remaining 78 species of the genus Pepsis, belonging to ten species-groups, are described and figured, and their phylogenetics and biogeography are discussed. 14 of the species are described as new: P. achterbergi spec. nov., P. adonta spec. nov., P. boharti spec. nov., P. caliente spec. nov., P. dayi spec. nov., P. esmeralda spec. nov., P. ianthoides spec. nov., P. jamaicensis spec. nov., P. krombeini spec. nov., P. martini spec. nov., P. multichroma spec. nov., P. nanoides spec. nov., P. wahisi spec. nov., and P. willinki spec. nov. Three species-names, P. infuscata Spinola, 1841, P. lampas Lucas, 1895, and P. thoreyi Dahlbom, 1845, are recalled from synonymy. The following 293 names are newly synonymized (the valid names are listed first): P. atalanta Mocsáry, 1885 = P. nitens Mocsáry, 1894, P. mocsaryi Lucas, 1895; P. inclyta Lepeletier, 1845 = P. mutabilis Lepeletier, 1845, P. vagabunda Lepeletier, 1845, P. cupripennis Taschenberg, 1869, P. violaceipennis Mocsáry, 1885, P. clotho Mocsáry, 1888, P. spengeli Mocsáry, 1888, P. sickmanni Mocsáry, 1888, P. nireus Mocsáry, 1894, P. atrovirens Lucas, 1895, P. cerastes Lucas, 1895, P. pygidialis Brèthes, 1908, P. guaranitica Brèthes, 1908, P. parca Lucas, 1919, P. atahualpa Banks, 1946, opimicornis, Haupt, 1952, atropos, Haupt, 1952, azurea Haupt, 1952; crassicornis Mocsáry, 1885 = P. sappho Brèthes, 1908, P. nitocris Brèthes, 1908, P. vivida Brèthes, 1908, P. arechavaletai Brèthes, 1908, P. lynchii Brèthes, 1908, P. operosa Brèthes, 1908, P. ataraqua Banks, 1946, P. splendida Haupt, 1952; P. sommeri Dahlbom, 1845 = P. azteca Cameron, 1893; P. xanthocera Dahlbom, 1843 = P. nigrescens Smith, 1855, P. fulgidipennis Mocsáry, 1885, P. juno Brèthes, 1908, P. ismare Banks, 1946, P. nigroprasina Haupt, 1952; P. seifferti Lucas, 1895 = P. cornuta Lucas, 1895, P. moebiusi Lucas, 1895, P. stygia Lucas, 1895; P. luteicornis Fabricius, 1804 = P. strenua Erichson, 1848, P. tinctipennis Smith, 1873, P. citreicornis Mocsáry, 1894, P. venosa Banks, 1945, P. alector Banks, 1946; P. asteria Mocsáry, 1894 = P. luridicornis Brèthes, 1926; P. convexa Lucas, 1895 = P. humeralis Brèthes, 1914; P. helvolicornis Lucas, 1895 = P. bahiae Brèthes, 1914; P. vitripennis Smith, 1855 = P. obscura Lepeletier, 1845, P. amabilis Mocsáry, 1885, P. centralis Cameron, 1893, P. margarete Lucas, 1895, P. venezuelae Kaye, 1913, P. aeneipennis Banks, 1946, P. helenae Haupt, 1952, P. coeruleoviridis Haupt, 1952; P. fumipennis Smith, 1855 = P. pallidicornis Mocsáry, 1885; P. amyntas Mocsáry, 1885 = P. vicina Lucas, 1895, P. clarinervis Brèthes, 1908, P. amyntoides Lucas, 1919, P. eurydice Lucas, 1919; P. dimidiata Fabricius, 1804 = P. vittigera Lucas, 1897, P. argentina Brèthes, 1908, P. sanctaeannae Brèthes, 1908, P. virgo Brèthes, 1908, P. externa Brèthes, 1908, P. transversa Brèthes, 1908, P. cordubensis Brèthes, 1908, P. banghaasi Lucas, 1919; P. menechma Lepeletier, 1845 = P. elegans Lepeletier, 1845, P. dubitata Cresson, 1867, P. prismatica Smith, 1855, P. advena Mocsáry, 1885, cinctipennis Mocsáry, 1885, P. guatemalensis Cameron, 1893, P. nestor Mocsáry, 1894, P. nigricornis Mocsáry, 1894, P. auranticornis Lucas, 1895, P. fruhstorferi Lucas, 1895, P. concolor Lucas, 1895, P. cerberus Lucas, 1895, P. euchroma Lucas, 1895, P. nigrocincta Lucas, 1895, P. mordax Lucas, 1895, P. inermis Fox, 1898, P. roberti Brèthes, 1908, P. janira Brèthes, 1908, P. cultrata Brèthes, 1908, P. novitia Banks, 1921; P. decipiens Lucas, 1895 = P. similis Lucas, 1895; P. minarum Brèthes, 1914 = P. pulchra Brèthes, 1914; P. basifusca Lucas, 1895 = P. angustimarginata Viereck, 1908; P. chrysoptera Burmeister, 1872 = P. exigua Lucas, 1895, P. smaragdinula Lucas, 1895, P. nebulosa Lucas, 1895, P. karschi Lucas, 1895, P. anisitsii Brèthes, 1908, P. indistincta Brèthes, 1908, P. dimidiatipennis Brèthes, 1908, P. chloroptera Brèthes, 1908, P. culta Brèthes, 1908, P. recta Brèthes, 1908, P. tornowii Brèthes, 1908, P. schrottkyi Brèthes, 1908, P. itinerata Brèthes, 1908, P. miniata Brèthes, 1908, P. spegazzinii Brèthes, 1908, P. paulistana Brèthes, 1914, P. chloe Brèthes, 1914, P. coronaria Brèthes, 1914, P. semilucana Haupt, 1952, P. bruneipes Haupt, 1952, P. brachynotus Haupt, 1952, P. diagonalis Haupt, 1952, P. discrepans Haupt, 1952; P. elongata Lepeletier, 1845 = P. purpurascens Smith, 1855, P. fuscipennis Smith, 1873, P. longula Banks, 1946; P. australis Saussure, 1867 = P. centaurus Lucas, 1897; P. cyanescens Lepeletier, 1845 = P. micans Mocsáry, 1885, P. jucunda Mocsáry, 1885, P. balloui Banks, 1946, P. diversa Haupt, 1952; P. lampas Lucas, 1895 = P. venturii Schrottky, 1902; P. nitida Lepeletier, 1845 = P. lucidula Smith, 1855, P. vaualba Smith, 1855, P. pruinosa Mocsáry, 1894, P. cylindrica Lucas, 1895, P. andina Brèthes, 1908, P. dilatata Brèthes, 1908, P. holmbergi Brèthes, 1908, P. concava Brèthes, 1908, P. ephebus Brèthes, 1908, P. vaga Brèthes, 1908, P. fuscobasalis Brèthes, 1908, P. cordata Brèthes, 1914, P. impatiens Brèthes, 1914, P. tricolor Brèthes, 1914, P. joergenseni Brèthes, 1914, P. cleone Brèthes, 1914, P. dorsata Brèthes, 1914, P. aretheas Brèthes, 1914, P. lassonis Lucas, 1819, P. consors Banks, 1946, P. interrupta Banks, 1946, P. analis Haupt, 1952; P. seladonica Dahlbom, 1843 = P. deuteroleuca Smith, 1855, P. kohli Lucas, 1895, P. venezolana Brèthes, 1908, P. burmeisteri Brèthes, 1908; P. cybele Banks, 1945 = P. weberi Banks, 1946; P. thoreyi Dahlbom, 1845 = P. lurida Lucas, 1895, P. euterpe Brèthes, 1908; P. flavescens Lucas, 1895 = P. periphetes Lucas, 1895, P. limbatella Brèthes, 1908, P. discoidalis Brèthes, 1914, P. limbatica Brèthes, 1914, P. militaris Brèthes, 1914, P. cavillatrix Haupt, 1952, P. arcuata Haupt, 1952, P. recterugosa Haupt, 1952, P. adversatrix Haupt, 1952; P. nigricans Lucas, 1895 = P. troglodytes Brèthes, 1908; P. montezuma Smith, 1855 = P. quitonensis Packard, 1869, P. sibylla Mocsáry, 1885, P. circe Mocsáry, 1885, P. occidentalis Cameron, 1893, P. peruanus Lucas, 1895, P. fulva Lucas, 1895, P. nessus Lucas, 1895, P. fusca Lucas, 1895, P. andicola Cameron, 1903, P. chilloensis Cameron, 1903, P. patagonica Brèthes, 1908, P. fasciculata Brèthes, 1908, P. pisoensis Strand, 1911, P. pacifica Brèthes, 1914, P. huascar Banks, 1946; P. completa Smith, 1855 = P. quichua Brèthes, 1908, P. comes Banks, 1946; P. smaragdina Dahlbom, 1843 = P. thunbergi Dahlbom, 1843, P. lara Mocsáry, 1888, P. satrapes Lucas, 1895, P. nupta Lucas, 1895, P. erynnis Lucas, 1895, P. fraterna Lucas, 1895, P. diabolus Lucas, 1895, P. mystica Lucas, 1895, P. thalia Brèthes, 1908, P. brasiliensis Brèthes, 1908, P. pallida Brèthes, 1908, P. iheringi Brèthes, 1908, P. dromeda Brèthes, 1908, P. sepultrix Lucas, 1919, P. strickeri Lucas, 1919; P. discolor Taschenberg, 1869 = P. sinnis Lucas, 1895, P. jujuyensis Brèthes, 1908, P. modesta Brèthes, 1908, P. comparata Brèthes, 1908, P. neutra Brèthes, 1908, P. terebrans Brèthes, 1908, P. procera Haupt, 1952, P. plaumanni Haupt, 1952, P. ogloblini Haupt, 1952, P. deletrix Haupt, 1952; P. limbata Guérin, 1831 = P. richteri Brèthes, 1908, P. polita Brèthes, 1908, P. limbella Haupt, 1952, P. artemis Haupt, 1952; P. basalis Mocsáry, 1885 = P. erdmanni Lucas, 1895, P. basinigra Haupt, 1952; P. infuscata Spinola, 1841 = P. niobe Mocsáry, 1885, P. sagana Mocsáry, 1894, P. incerta Banks, 1946; P. hyalinipennis Mocsáry, 1885 = P. subruficornis Haupt, 1952; P. festiva Fabricius, 1804 = P. pulchella Lepeletier, 1845, P. solitaria Smith, 1879, P. gallardoi Brèthes, 1908, P. hora Brèthes, 1914, P. amok Lucas, 1919, P. riojaneirensis Lucas, 1919; P. gracilis Lepeletier, 1845 = P. diana Mocsáry, 1885, P. hecate Mocsáry, 1885, P. spathulifera Lucas, 1895, P. sphinx Lucas, 1895, P. ierensis Banks, 1945, P. alceste Banks, 1946, P. scalaris Haupt, 1952; P. mildei Stål, 1857 = P. charon Mocsáry, 1885, P. cyanoptera Lucas, 1895, P. dryas Lucas, 1919; P. filiola Brèthes, 1914 = P. denserugosa Haupt, 1952; P. ruficornis Fabricius, 1804 = P. saphirus Palisot de Beauvois, 1805, P. violacea Mocsáry, 1885, P. hexamita Lucas, 1895, P. omniviolacea Haupt, 1952; P. brunneicornis Lucas, 1895 = P. glabripennis Lucas, 1895; P. purpurea Smith, 1873 = P. pan Mocsáry, 1885, P. parthenope Mocsáry, 1885, P. sagax Lucas, 1895, P. clypeata Brèthes, 1914, P. consimilis Banks, 1946, P. laconia Banks, 1946; P. viridisetosa Spinola, 1841 = P. eximia Smith, 1873; P. viridis Lepeletier, 1845 = P. errans Lepeletier, 1845, P. chlorotica Mocsáry, 1885, P. excelsa Lucas, 1895, P. selene Lucas, 1895, P. fimbriata Lucas, 1895, P. calypso Brèthes, 1908, P. fluminensis Brèthes, 1908, P. argentinicus Strand, 1910, P. mimetica Brèthes, 1914, P. garbei Brèthes, 1914, P. erecta Brèthes, 1914, P. tandilensis Brèthes, 1914, P. meridionalis Brèthes, 1914, P. minor Lucas, 1919, P. basifulgens Lucas, 1919, P. nebulosipennis Lucas, 1919, P. purpurea Lucas, 1919, P. koerberi Lucas, 1919, P. inimicissima Lucas, 1919, P. debilitans Lucas, 1919, P. itapaca Banks, 1946; P. aciculata Taschenberg, 1869 = P. nero Lucas, 1895; P. atripennis Fabricius, 1804 = P. flavilis Brèthes, 1908; P. ianthina Erichson, 1848 = P. fulvicornis Mocsáry, 1885, P. sirene Lucas, 1895, P. balboae Lucas, 1919, P. herodes Lucas, 1919, P. curti Lucas, 1919; P. nana Mocsáry, 1885 = P. mapiriensis Lucas, 1919, P. vinciens Lucas, 1919, P. ilione Banks, 1946, P. moesta Banks, 1946, P. orestes Banks, 1946, P. amautas Banks, 1946, P. inaequalis Haupt, 1952; P. hirtiventris Banks, 1946 = P. viridaurea Haupt, 1952, P. aequalis Haupt, 1952; P. auriguttata Burmeister, 1872 = P. aurimacula Mocsáry, 1885, P. flavicornis Mocsáry, 1894, P. guttata Lucas, 1895, P. incendiaria Lucas, 1895, P. pubiventris Lucas, 1895, P. planifrons Lucas, 1895, P. lestes Lucas, 1895, P. villosa Brèthes, 1908; P. sabina Mocsáry, 1885 = P. astioles Banks, 1946; and P. purpureipes Packard, 1869 = P. chlorana Mocsáry, 1885, P. antennalis Cameron, 1893, P. sulcifrons Cameron, 1903, P. carinata Brèthes, 1914, P. equatoriana Brèthes, 1914, P. angusta Banks, 1946. Keys to all forms are given. The mimicry-groups of P. atripennis Fabricius, 1804, and P. completa Smith, 1855, are defined and described and a comparative account of mimicry based on all four mimicrygroups in Pepsis is given. Lists of excluded species (with their current taxonomic placement and depository where ascertained), unplaced names, and a nomen nudum are given.

Beschreibung: Nonequilibrium phenomena in hypersonic flows are examined on the basis of theoretical models and selected experimental data, in an introduction intended for second-year graduate students of aerospace engineering. Chapters are devoted to the physical nature of gas atoms and molecules, transitions of internal states, the formulation of the master equation of aerothermodynamics, the conservation equations, chemical reactions in CFD, the behavior of air flows in nonequilibrium, experimental aspects of nonequilibrium flow, a review of experimental results, and gas-solid interaction. Diagrams, graphs, and tables of numerical data are provided.

Beschreibung: The introduction of transverse velocity fluctuations into a separated shear layer on an airfoil at high angles of attack is presently demonstrated to be an effective separation-control technique. Airfoil aerodynamic characteristics, including poststall lift and drag as well as maximum lift coefficient and stall angle, all exhibited improvements controlled forcing at 20 deg angle of attack led to an increased spreading of the mean velocity profile, together with increased turbulence activity; separation moved from the leading edge to about 80 percent of chord.

Beschreibung: Flow characteristics in the vicinity of the flap of a single-slotted airfoil are presented and analyzed. The flow remained attached over the model surfaces, except in the vicinity of the flap trailing edge where a small region of boundary-layer separation extended over the aft 7 percent of flap chord. The airfoil configuration was tested at a Mach number of 0.09 and a chord Reynolds number of 1.8 x 10 to the 6th in the NASA Ames Research Center 7- by 10-Foot Wind Tunnel. The flow was complicated by the presence of a strong, initially inviscid, jet, emanating from the slot between airfoil and flap, and a gradual merging of the main airfoil wake and flap suction-side boundary layer.

Beschreibung: The transonic aspect of helicopter flow analysis is addressed. The equations of motion and their implementations are examined, and the computation of real rotor flows is considered. Nonlifting rotor flows, high-speed hover, high advance ratio lifting rotor flows, and strong blade/vortex interaction computations are discussed.

Beschreibung: Some of the basic finite-difference schemes that can be used to solve the nonlinear equations that describe unsteady inviscid and viscous transonic flow are reviewed. Numerical schemes for solving the unsteady Euler and Navier-Stokes, boundary-layer, and nonlinear potential equations are described. Emphasis is given to the elementary ideas used in constructing various numerical procedures, not specific details of any one procedure.

Beschreibung: The flight testing conducted over the past 10 years in the NASA laminar-flow control (LFC) will be reviewed. The LFC program was directed towards the most challenging technology application, the high supersonic speed transport. To place these recent experiences in perspective, earlier important flight tests will first be reviewed to recall the lessons learned at that time.

Beschreibung: Although most of the laminar flow airfoils recently developed at the NASA Langley Research Center were intended for general aviation applications, low-drag airfoils were designed for transonic speeds and wind tunnel performance tested. The objective was to extend the technology of laminar flow to higher Mach and Reynolds numbers and to swept leading edge wings representative of transport aircraft to achieve lower drag and significantly improved operation costs. This research involves stabilizing the laminar boundary layer through geometric shaping (Natural Laminar Flow, NLF) and active control involving the removal of a portion of the laminar boundary layer (Laminar-Flow Control, LFC), either through discrete slots or perforated surface. Results show that extensive regions of laminar flow with large reductions in skin friction drag can be maintained through the application of passive NLF boundary-layer control technologies to unswept transonic wings. At even greater extent of laminar flow and reduction in the total drag level can be obtained on a swept supercritical airfoil with active boundary layer-control.

Beschreibung: Aerodynamic forces and moments for a slender wing-body configuration are summarized from an investigation in the Langley National Transonic Facility (NTF). The results include both longitudinal and lateral-directional aerodynamic properties as well as slideslip derivatives. Results were selected to emphasize Reynolds number effects at a transonic speed although some lower speed results are also presented for context. The data indicate nominal Reynolds number effects on the longitudinal aerodynamic coefficients and more pronounced effects for the lateral-directional aerodynamic coefficients. The Reynolds number sensitivities for the lateral-directional coefficients were limited to high angles of attack.

Beschreibung: The objective is to provide useful engineering formulations and to instill a modest degree of physical understanding of the phenomena governing convective aerodynamic heating at high flight speeds. Some physical insight is not only essential to the application of the information presented here, but also to the effective use of computer codes which may be available to the reader. Given first is a discussion of cold-wall, laminar boundary layer heating. A brief presentation of the complex boundary layer transition phenomenon follows. Next, cold-wall turbulent boundary layer heating is discussed. This topic is followed by a brief coverage of separated flow-region and shock-interaction heating. A review of heat protection methods follows, including the influence of mass addition on laminar and turbulent boundary layers. Next is a discussion of finite-difference computer codes and a comparison of some results from these codes. An extensive list of references is also provided from sources such as the various AIAA journals and NASA reports which are available in the open literature.

Beschreibung: A co-operative testing program is in progress between the Langley Research Center (NASA) and the National Aeronautical Establishment (NAE, Canada) to validate two different techniques of airfoil testing at transonic speeds. The procedure employed is to test the same airfoil model in the NAE two-dimensional tunnel and the Langley 0.3-m Transonic Cryogenic Tunnel (0.3-m TCT). The airfoil model used in testing was CAST-10-2/DOA-2 super-critical airfoil. The Langley 0.3-m TCT has a relatively small cross section of 13 in x 13 in, giving a (h/c) ratio of 1.44 for the same 9 in chord model. The approach employed in the 0.3-m TCT aims towards eliminating the wall effects by using active walls. The top and bottom walls are flexible. By changing the wall shapes during a test in an iterative manner, the wall interference effects are reduced. The method employed to change the wall shapes is the adaptive wall technique. The current test program provided an opportunity to validate the adaptive wall technique in the 0.3-m TCT. The relatively long chord airfoil represents a severe test case to test the efficacy of the adaptive wall technique under cryogenic conditions. The program also involved removal of side wall boundary-layer thus increasing the complexity of the wall adaptation technique. This paper deals with some salient results obtained regarding repeatability of test data and possible residual interference effects.

Beschreibung: The two-dimensional (2-D) and three-dimensional Navier-Stokes equations are solved for flow over a NAE CAST-10 airfoil model. Recently developed finite-volume codes that apply a multistage time stepping scheme in conjunction with steady state acceleration techniques are used to solve the equations. Two-dimensional results are shown for flow conditions uncorrected and corrected for wind tunnel wall interference effects. Predicted surface pressures from 3-D simulations are compared with those from 2-D calculations. The focus of the 3-D computations is the influence of the sidewall boundary layers. Topological features of the 3-D flow fields are indicated. Lift and drag results are compared with experimental measurements.

Beschreibung: An experimental Adaptive Wall Test Section (AWTS) process is described. Comparisons of the ONERA T2 and the 0.3-m TCT (transonic cryogenic tunnel) AWTS data for the ONERA CAST-10 airfoil are presented. Most of the 0.3-m TCT data is new and preliminary and no sidewall boundary layer control is involved. No conclusions are given.

Beschreibung: The transonic airfoil CAST 10-2/DOA 2 was investigated in several major transonic wind tunnels at Reynolds numbers ranging from Re=1.3 x 10(exp 6) to 45 x 10(exp 6) at ambient and cryogenic temperature conditions. The main objective was to study the degree and extent of the effects of Reynolds number on both the airfoil aerodynamic characteristics and the interference effects of various model-wind-tunnel systems. The initial analysis of the CAST 10-2 airfoil results revealed appreciable real Reynolds number effects on this airfoil and showed that wall interference can be significantly affected by changes in Reynolds number thus appearing as true Reynolds number effects.

Beschreibung: Hypersonic vehicles operate in a hostile aerothermal environment which has a significant impact on their aerothermostructural performance. Significant coupling occurs between the aerodynamic flow field, structural heat transfer, and structural response creating a multidisciplinary interaction. A long term goal of the Aerothermal Loads Branch at the NASA Langley Research Center is to develop a computational capability for integrated fluid, thermal and structural analysis of aerodynamically heated structures. The integrated analysis capability includes the coupling between the fluid and the structure which occurs primarily through the thermal response of the structure, because: (1) the surface temperature affects the external flow by changing the amount of energy absorbed by the structure, and (2) the temperature gradients in the structure result in structural deformations which alter the flow field and attendant surface pressures and heating rates. In the integrated analysis, a finite element method is used to solve: (1) the Navier-Stokes equations for the flow solution, (2) the energy equation of the structure for the temperature response, and (3) the equilibrium equations of the structure for the structural deformation and stresses.

Beschreibung: Wind tunnel wall interference assessment and correction (WIAC) concepts, applications, and typical results are discussed in terms of several nonlinear transonic codes and one panel method code developed for and being implemented at NASA-Langley. Contrasts between 2-D and 3-D transonic testing factors which affect WIAC procedures are illustrated using airfoil data from the 0.3 m Transonic Cryogenic Tunnel and Pathfinder 1 data from the National Transonic Facility. Initial results from the 3-D WIAC codes are encouraging; research on and implementation of WIAC concepts continue.

Beschreibung: The stability of compressible 2-D and 3-D boundary layers is reviewed. The stability of 2-D compressible flows differs from that of incompressible flows in two important features: There is more than one mode of instability contributing to the growth of disturbances in supersonic laminar boundary layers and the most unstable first mode wave is 3-D. Whereas viscosity has a destabilizing effect on incompressible flows, it is stabilizing for high supersonic Mach numbers. Whereas cooling stabilizes first mode waves, it destabilizes second mode waves. However, second order waves can be stabilized by suction and favorable pressure gradients. The influence of the nonparallelism on the spatial growth rate of disturbances is evaluated. The growth rate depends on the flow variable as well as the distance from the body. Floquet theory is used to investigate the subharmonic secondary instability.

Beschreibung: The treatment of turbulence effects on transonic shock/turbulent boundary layer interaction is addressed within the context of a triple deck approach valid for arbitrary practical Reynolds numbers between 1000 and 10 billion. The modeling of the eddy viscosity and basic turbulent boundary profile effects in each deck is examined in detail using Law-of-the-Wall/Law-of-the-Wake concepts as the foundation. Results of parametric studies showing how each of these turbulence model aspects influences typical interaction zone property distributions (wall pressure, displacement thickness and local skin friction) are presented and discussed.

Beschreibung: The aerodynamic characteristics for both single and twin-engine high-performance aircraft are significantly affected by shock induced flow interactions as well as other local flow interference effects which usually occur at transonic speeds. These adverse interactions can not only cause high drag, but also cause unusual aerodynamic loadings and/or severe stability and control problems. Many programs are under way to not only develop method for reducing the adverse effects, but also to develop an understanding of the basic flow conditions which are the primary contributors. It is anticipated that these programs will result in technologies which can reduce the aircraft cruise drag through improved integration as well as increase aircraft maneuverability through the application of thrust vectoring. Some of the primary integration problems for twin-engine aircraft at transonic speeds are identified, and several methods are demonstrated for reducing or eliminating the undersirable characteristics, while enhancing configuration effectiveness.

Beschreibung: Algorithms are described for the generation and adaptation of unstructured grids in two and three dimensions, as well as Euler solvers for unstructured grids. The main purpose is to demonstrate how unstructured grids may be employed advantageously for the economic simulation of both geometrically as well as physically complex flow fields.

Beschreibung: For many internal transonic flows of practical interest, some of the relevant nondimensional parameters typically are small enough that a perturbation scheme can be expected to give a useful level of numerical accuracy. A variety of steady and unsteady transonic channel and cascade flows is studied with the help of systematic perturbation methods which take advantage of this fact. Asymptotic representations are constructed for small changes in channel cross-section area, small flow deflection angles, small differences between the flow velocity and the sound speed, small amplitudes of imposed oscillations, and small reduced frequencies. Inside a channel the flow is nearly one-dimensional except in thin regions immediately downstream of a shock wave, at the channel entrance and exit, and near the channel throat. A study of two-dimensional cascade flow is extended to include a description of three-dimensional compressor-rotor flow which leads to analytical results except in thin edge regions which require numerical solution. For unsteady flow the qualitative nature of the shock-wave motion in a channel depends strongly on the orders of magnitude of the frequency and amplitude of impressed wall oscillations or fluctuations in back pressure. One example of supersonic flow is considered, for a channel with length large compared to its width, including the effect of separation bubbles and the possibility of self-sustained oscillations. The effect of viscosity on a weak shock wave in a channel is discussed.

Beschreibung: The 1980s may well be called the Euler era of applied aerodynamics. Computer codes based on discrete approximations of the Euler equations are now routinely used to obtain solutions of transonic flow problems in which the effects of entropy and vorticity production are significant. Such codes can even predict separation from a sharp edge, owing to the inclusion of artificial dissipation, intended to lend numerical stability to the calculation but at the same time enforcing the Kutta condition. One effect not correctly predictable by Euler codes is the separation from a smooth surface, and neither is viscous drag; for these some form of the Navier-Stokes equation is needed. It, therefore, comes as no surprise to observe that the Navier-Stokes has already begun before Euler solutions were fully exploited. Moreover, most numerical developments for the Euler equations are now constrained by the requirement that the techniques introduced, notably artificial dissipation, must not interfere with the new physics added when going from an Euler to a full Navier-Stokes approximation. In order to appreciate the contributions of Euler solvers to the understanding of transonic aerodynamics, it is useful to review the components of these computational tools. Space discretization, time- or pseudo-time marching and boundary procedures, the essential constituents are discussed. The subject of grid generation and grid adaptation to the solution are touched upon only where relevant. A list of unanswered questions and an outlook for the future are covered.

Beschreibung: The application of computational fluid dynamics (CFD) to fighter aircraft design and development is discussed. Methodology requirements for the aerodynamic design of fighter aircraft are briefly reviewed. The state-of-the-art of computational methods for transonic flows in the light of these requirements is assessed and the techniques found most adequate for the subject application are identified. Highlights from some proof-of-feasibility Euler and Navier-Stokes computations about a complete fighter aircraft configuration are presented. Finally, critical issues and opportunities for design application of CFD are discussed.

Beschreibung: A brief survey is given on the study of transonic shock/boundary layer effects in flight. Then the possibility of alleviating the adverse shock effects through passive shock control is discussed. A Swedish flight experiment on a swept wing attack aircraft is used to demonstrate how it is possible to reduce the extent of separated flow and increase the drag-rise Mach number significantly using a moderate amount of perforation of the surface.

Beschreibung: An iterative method for wall interference assessment and/or correction is presented for transonic flow conditions in wind tunnels equipped with two component velocity measurements on a single interface. The iterative method does not require modeling of the test article and tunnel wall boundary conditions. Analytical proof for the convergence and stability of the iterative method is shown in the subsonic flow regime. The numerical solutions are given for both 2-D and axisymmetrical cases at transonic speeds with the application of global Mach number correction.

Beschreibung: An intense research effort over the last few years has produced several competing and apparently diverse methods for generating meshes. Recent progress is reviewed and the central themes are emphasized which form a solid foundation for future developments in mesh generation.

Beschreibung: A cell-vertex scheme is outlined for solving the flow about a delta wing with M (sub infinity) is greater than 1. Embedded regions of mesh refinement allow solutions to be obtained which have much higher resolution than those achieved to date. Effects of mesh refinement and artificial viscosity on the solutions are studied, to determine at what point leading-edge vortex solutions are grid-converged. A macroscale and a microscale for the size of the vortex are defined, and it is shown that the macroscale (which includes the wing surface properties) is converged on a moderately refined grid, while the microscale is very sensitive to grid spacing. The level of numerical diffusion in the core of the vortex is found to be substantial. Comparisons with the experiment are made for two cases which have transonic cross-flow velocities.

Beschreibung: A computer analysis was developed for calculating steady (or unsteady) three-dimensional aircraft component flow fields. This algorithm, called ENS3D, can compute the flow field for the following configurations: diffuser duct/thrust nozzle, isolated wing, isolated fuselage, wing/fuselage with or without integrated inlet and exhaust, nacelle/inlet, nacelle (fuselage) afterbody/exhaust jet, complete transport engine installation, and multicomponent configurations using zonal grid generation technique. Solutions can be obtained for subsonic, transonic, or hypersonic freestream speeds. The algorithm can solve either the Euler equations for inviscid flow, the thin shear layer Navier-Stokes equations for viscous flow, or the full Navier-Stokes equations for viscous flow. The flow field solution is determined on a body-fitted computational grid. A fully-implicit alternating direction implicit method is employed for the solution of the finite difference equations. For viscous computations, either a two layer eddy-viscosity turbulence model or the k-epsilon two equation transport model can be used to achieve mathematical closure.

Beschreibung: The use of computational methods for three dimensional transonic flow design and analysis at the Boeing Company is presented. A range of computational tools consisting of production tools for every day use by project engineers, expert user tools for special applications by computational researchers, and an emerging tool which may see considerable use in the near future are described. These methods include full potential and Euler solvers, some coupled to three dimensional boundary layer analysis methods, for transonic flow analysis about nacelle, wing-body, wing-body-strut-nacelle, and complete aircraft configurations. As the examples presented show, such a toolbox of codes is necessary for the variety of applications typical of an industrial environment. Such a toolbox of codes makes possible aerodynamic advances not previously achievable in a timely manner, if at all.

Beschreibung: Progress in a recently started project aimed at the prediction of transition to turbulence in hypersonic flow is briefly discussed. The prediction of transition to turbulence is a very important issue in the design of space vessels. Two space vehicles currently under investigation, namely the aeroassisted transfer vehicle (AOTV) and the trans-atmospheric vehicle (TAV), suffer from strong aerodynamic heating. This heating is strongly influenced by the boundary layer structure. These aerospace vehicles fly in the upper atmospheric layer at a Mach number between 10 and 30 at very low atmospheric pressures. At very high altitudes the flow is laminar, but when the space vessel returns to a lower orbit, the flow becomes turbulent and the heating is dramatically increased. The prediction of this transition process is commonly done by means of experiments. The experimental facilities available nowadays cannot model the hypersonic flow field accurately enough by limitations in Mach and Reynolds number. These facilities also have a large free stream disturbance level which makes it very difficult to investigate transition accurately. An alternative approach is to study transition by theoretical means. Up to now numerical studies of hypersonic flow only discussed steady laminar or turbulent flow. This theoretical approach is extended to the study of transition in hypersonic flow by means of direct numerical simulations and additional theoretical investigations to explain the mechanisms leading to transition. A brief outline of how this research is to be performed is given.

Beschreibung: In the past decade, there has been much activity in the development of computational methods for the analysis of unsteady transonic aerodynamics about airfoils and wings. Significant features are illustrated which must be addressed in the treatment of computational transonic unsteady aerodynamics. The flow regimes for an aircraft on a plot of lift coefficient vs. Mach number are indicated. The sequence of events occurring in air combat maneuvers are illustrated. And further features of transonic flutter are illustrated. Also illustrated are several types of aeroelastic response which were encountered and which offer challenges for computational methods. The four cases illustrate problem areas encountered near the boundaries of aircraft envelopes, as operating condition change from high speed, low angle conditions to lower speed, higher angle conditions.

Beschreibung: A unified formulation is presented based on the full potential framework coupled with an appropriate structural model to compute steady and unsteady flows over rigid and flexible configurations across the Mach number range. The unsteady form of the full potential equation in conservation form is solved using an implicit scheme maintaining time accuracy through internal Newton iterations. A flux biasing procedure based on the unsteady sonic reference conditions is implemented to compute hyperbolic regions with moving sonic and shock surfaces. The wake behind a trailing edge is modeled using a mathematical cut across which the pressure is satisfied to be continuous by solving an appropriate vorticity convection equation. An aeroelastic model based on the generalized modal deflection approach interacts with the nonlinear aerodynamics and includes both static as well as dynamic structural analyses capability. Results are presented for rigid and flexible configurations at different Mach numbers ranging from subsonic to supersonic conditions. The dynamic response of a flexible wing below and above its flutter point is demonstrated.

Beschreibung: One of the most important uses of method that calculate unsteady aerodynamic loads is to predict and analyze the aeroelastic responses of flight vehicles. Currently, methods based on transonic small disturbance potential aerodynamics are the primary tools for aeroelastic analysis. Flow solutions obtained using isentropic potential theory can be highly inaccurate and even multivalued, because they do not model the effects of entropy that is produced when shock waves are in the flow field. From the results that are presented, it is concluded that nonisentropic potential methods more accurately model Euler solutions than do isentropic methods. The primary effects of modeling shock generated entropy are: (1) to eliminate mulitple flow solutions when strong shock waves are in the flow field; and (2) to bring the strengths and locations of computed shock waves into better agreement with those calculated using Euler method and those measured during experiments.

Beschreibung: A finite difference technique is used to solve the transonic small disturbance flow equation making use of shock capturing to treat wave discontinuities. Thus the nonlinear effects of thickness and angle of attack are considered. Such an approach is made feasible by the development of a new code called CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance), and is based on a fully implicit approximate factorization (AF) finite difference method to solve the time dependent transonic small disturbance equation. The application of the CAP-TSD code to the calculation of low to moderate supersonic steady and unsteady flows is presented. In particular, comparisons with exact linear theory solutions are made for steady and unsteady cases to evaluate shock capturing and other features of the current method. In addition, steady solutions obtained from an Euler code are used to evaluate the small disturbance aspects of the code. Steady and unsteady pressure comparisons are made with measurements for an F-15 wing model and for the RAE tailplane model.

Beschreibung: Two wind tunnel investigations were conducted to assess two different wall interference alleviation/correction techniques: adaptive test section walls and classical analytical corrections. The same airfoil model has been tested in the adaptive wall test section of the NASA-Langley 0.3 m Transonic Cryogenic Tunnel (TCT) and in the National Aeronautical Establishment (NAE) High Reynolds Number 2-D facility. The model has a 9 in. chord and a CAST 10-2/DOA 2 airfoil section. The 0.3 m TCT adaptive wall test section has four solid walls with flexible top and bottom walls. The NAE test section has porous top and bottom walls and solid side walls. The aerodynamic results corrected for top and bottom wall interference at Mach numbers from 0.3 to 0.8 at a Reynolds number of 10 by 1,000,000. Movement of the adaptive walls was used to alleviate the top and bottom wall interference in the test results from the NASA tunnel.

Beschreibung: Nonintrusive measurements were made of a normal shock wave/boundary layer interaction. Two dimensional measurements were made throughout the interaction region while 3-D measurements were made in the vicinity of the shock wave. The measurements were made in the corner of the test section of a continuous supersonic wind tunnel in which a normal shock wave had been stabilized. Laser Doppler Anemometry, surface pressure measurement and flow visualization techniques were employed for two freestream Mach number test cases: 1.6 and 1.3. The former contained separated flow regions and a system of shock waves. The latter was found to be far less complicated. The results define the flow field structure in detail for each case.

Beschreibung: Three dimensional linear secondary instability theory is extended for compressible boundary layers on a flat plate in the presence of finite amplitude Tollmien-Schlichting waves. The focus is on principal parametric resonance responsible for strong growth of subharmonics in low disturbance environment.

Beschreibung: A review is made of the performance of a variety of turbulence models in the evaluation of a particular well documented transonic flow. This is done to supplement a previous attempt to calibrate and verify transonic airfoil codes by including many more turbulence models than used in the earlier work and applying the calculations to an experiment that did not suffer from uncertainties in angle of attack and was free of wind tunnel interference. It is found from this work, as well as in the earlier study, that the Johnson-King turbulence model is superior for transonic flows over simple aerodynamic surfaces, including moderate separation. It is also shown that some field equation models with wall function boundary conditions can be competitive with it.

Beschreibung: Computational fluid dynamics has an increasingly important role in the design and analysis of aircraft as computer hardware becomes faster and algorithms become more efficient. Progress is being made in two directions: more complex and realistic configurations are being treated and algorithms based on higher approximations to the complete Navier-Stokes equations are being developed. The literature indicates that linear panel methods can model detailed, realistic aircraft geometries in flow regimes where this approximation is valid. As algorithms including higher approximations to the Navier-Stokes equations are developed, computer resource requirements increase rapidly. Generation of suitable grids become more difficult and the number of grid points required to resolve flow features of interest increases. Recently, the development of large vector computers has enabled researchers to attempt more complex geometries with Euler and Navier-Stokes algorithms. The results of calculations for transonic flow about a typical transport and fighter wing-body configuration using thin layer Navier-Stokes equations are described along with flow about helicopter rotor blades using both Euler/Navier-Stokes equations.

Beschreibung: Computational results are presented for three advanced configurations: the F-16A with wing tip missiles and under wing fuel tanks, the Oblique Wing Research Aircraft, and an Advanced Turboprop research model. These results were generated by the latest version of the TranAir full potential code, which solves for transonic flow over complex configurations. TranAir embeds a surface paneled geometry definition in a uniform rectangular flow field grid, thus avoiding the use of surface conforming grids, and decoupling the grid generation process from the definition of the configuration. The new version of the code locally refines the uniform grid near the surface of the geometry, based on local panel size and/or user input. This method distributes the flow field grid points much more efficiently than the previous version of the code, which solved for a grid that was uniform everywhere in the flow field. TranAir results are presented for the three configurations and are compared with wind tunnel data.

Beschreibung: Vector potential and related methods, for the simulation of both inviscid and viscous flows over aerodynamic configurations, are briefly reviewed. The advantages and disadvantages of several formulations are discussed and alternate strategies are recommended. Scalar potential, modified potential, alternate formulations of Euler equations, least-squares formulation, variational principles, iterative techniques and related methods, and viscous flow simulation are discussed.

Beschreibung: Advancements have occurred in transonic numerical simulation that place aerodynamic performance design into a relatively well developed status. Efficient broad band operating characteristics can be reliably developed at the conceptual design level. Recent aeroelastic and separated flow simulation results indicate that systematic consideration of an increased range of design problems appears promising. This emerging capability addresses static and dynamic structural/aerodynamic coupling and nonlinearities associated with viscous dominated flows.

Beschreibung: A review of several applications of Computational Fluid Dynamics (CFD) to various aspects of aerodynamic design recently carried out at Grumman is presented. The emphasis is placed on project-oriented applications where the ease of use of the methods and short start-to-completion times are required. Applications cover transonic wing design/optimization, wing mounted stores load prediction, transonic buffet alleviation, fuselage loads estimation, and compact offset diffuser design for advanced aircraft configurations. Computational methods employed include extended transonic small disturbance (automatic grid embedding) formulation for analysis/design/optimization and a thin layer Navier-Stokes formulation for both external and internal flow analyses.

Beschreibung: Numerous computational fluid dynamics (CFD) codes are available that solve any of several variations of the transonic flow equations from small disturbance to full Navier-Stokes. The design philosophy at General Dynamics Fort Worth Division involves use of all these levels of codes, depending on the stage of configuration development. Throughout this process, drag calculation is a central issue. An overview is provided for several transonic codes and representative test-to-theory comparisons for fighter-type configurations are presented. Correlations are shown for lift, drag, pitching moment, and pressure distributions. The future of applied CFD is also discussed, including the important task of code validation. With the progress being made in code development and the continued evolution in computer hardware, the routine application of these codes for increasingly more complex geometries and flow conditions seems apparent.

Beschreibung: Flight research and testing form a critical link in the aeronautic research and development chain. Brilliant concepts, elegant theories, and even sophisticated ground tests of flight vehicles are not sufficient to prove beyond a doubt that an unproven aeronautical concept will actually perform as predicted. Flight research and testing provide the ultimate proof that an idea or concept performs as expected. Ever since the Wright brothers, flight research and testing were the crucible in which aeronautical concepts were advanced and proven to the point that engineers and companies are willing to stake their future to produce and design aircraft. This is still true today, as shown by the development of the experimental X-30 aerospace plane. The Dryden Flight Research Center (Ames-Dryden) continues to be involved in a number of flight research programs that require understanding and characterization of the total airplane in all the aeronautical disciplines, for example the X-29. Other programs such as the F-14 variable-sweep transition flight experiment have focused on a single concept or discipline. Ames-Dryden also continues to conduct flight and ground based experiments to improve and expand the ability to test and evaluate advanced aeronautical concepts. A review of significant aeronautical flight research programs and experiments is presented to illustrate both the progress being made and the challenges to come.

Beschreibung: Present flutter analysis methods do not accurately predict the flutter speeds in the transonic flow region for wings with supercritical airfoils. Aerodynamic programs using computational fluid dynamic (CFD) methods are being developed, but these programs need to be verified before they can be used with confidence. A wind tunnel test was performed to obtain all types of data necessary for correlating with CFD programs to validate them for use on high aspect ratio wings. The data include steady state and unsteady aerodynamic measurements on a nominal stiffness wing and a wing four times that stiffness. There is data during forced oscillations and during flutter at several angles of attack, Mach numbers, and tunnel densities.

Beschreibung: The model wing consists of a set of fiberglass panels mounted on a steel spar that spans the 8 ft test section of the UTRC Large Subsonic Wind Tunnel. The first use of this system was to measure surface pressures and flow conditions for a series of constant pitch rate ramps and sinusoidal oscillations a Mach number range, a Reynolds number range, and a pitch angle range. It is concluded that an increased pitch rate causes stall events to be delayed, strengthening of the stall vortex, increase in vortex propagation, and increase in unsteady airloads. The Mach number range causes a supersonic zone near the leading edge, stall vortex to be weaker, and a reduction of unsteady airloads.

Beschreibung: Steady and unsteady pressures were measured on a 14 percent supercritical airfoil at transonic Mach numbers at Reynolds numbers from 6,000,000 to 35,000,000. Instrumentation techniques were developed to measure unsteady pressures in a cryogenic tunnel at flight Reynolds numbers. Experimental steady data, corrected for wall effects show very good agreement with calculations from a full potential code with an interacted boundary layer. The steady and unsteady pressures both show a shock position that is dependent on Reynolds number. For a supercritical pressure distribution at a chord Reynolds number of 35,000,000 laminar flow was observed between the leading edge and the shock wave at 45 percent chord.

Beschreibung: In transonic flutter problems where shock motion plays an important part, it is believed that accurate predictions of the flutter boundaries will require the use of codes based on the Euler equations. Only Euler codes can obtain the correct shock location and shock strength, and the crucially important shock excursion amplitude and phase lag. The present study is based on the finite volume scheme developed by Jameson and Venkatakrishnan for the 2-D unsteady Euler equations. The equations are solved in integral form on a moving grid. The variable are pressure, density, Cartesian velocity components, and total energy.

Beschreibung: The initial application of the CAP-TSD computer program for wing flutter analysis is presented. Computational Aeroelasticity Program - Transonic Small Disturbance (CAP-TSD) is based on an approximate factorization (AF) algorithm that is stable and efficient on supercomputers with vector arithmetic. CAP-TSD was used to calculate steady and unsteady pressures on wings and configurations at subsonic, transonic, and supersonic Mach numbers. However, the CAP-TSD code has been developed primarily for aeroelastic analysis. The initial efforts for validation of the aeroelastic analysis capability is presented. The initial applications include two series of symmetric, planar wing planforms. Well defined modal properties are available for these wings. In addition, transonic flutter boundaries are available for evaluation of the transonic capabilities of CAP-TSD.

Beschreibung: The DAST Aeroelastic Research Wing had been previously in the NASA Langley TDT and an unusual instability boundary was predicted based upon supercritical response data. Contrary to the predictions, no instability was found during the present test. Instead a region of high dynamic wing response was observed which reached a maximum value between Mach numbers 0.92 and 0.93. The amplitude of the dynamic response increased directly with dynamic pressure. The reponse appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on the upper and lower wing surfaces. The onset of flow separation coincided with the occurrence of strong shocks on a surface. A controller was designed to suppress the wing response. The control law attenuated the response as compared with the uncontrolled case and added a small but significant amount of damping for the lower density condition.

Beschreibung: One aircraft configuration that shows great promise in achieving high performance is that of an asymmetrically swept wing. When compared to conventional swept wings, these advantages include higher lift to drag ratios and reduced takeoff and landing speeds, which translate into greater performance in terms of fuel comsumption, loiter time, and range. However, the oblique wing has a number of disadvantages because of its asymmetric configuration. The question is how to best achieve maximum stability and roll equilibrium without compromising performance. Using aeroelastic tailoring to enhance aeroelastic stability and control has been demonstrated in several analyses, especially for the forward swept wing. The advantages and disadvantages for the oblique wing configuration are discussed.

Beschreibung: The advantages of oblique wings have been the subject of numerous theoretical studies, wind tunnel tests, low speed flight models, and finally a low speed manned demonstrator, the AD-1. The specific objectives of the OWRA program are: (1) to establish the necessary technology base required to translate theoretical and experimental results into practical mission oriented designs; (2) to design, fabricate and flight test an oblique wing aircraft throughout a realistic flight envelope, and (3) to develop and validate design and analysis tools for asymmetric aircraft configurations. The preliminary design phase of the project is complete and has resulted in a wing configuration for which construction is ready to be initiated.

Beschreibung: A finite volume implicit approximate factorization method which solves the thin layer Navier-Stokes equations was used to predict unsteady turbulent flow airfoil behavior. At a constant angle of attack of 16 deg, the NACA 0012 airfoil exhibits an unsteady periodic flow field with the lift coefficient oscillating between 0.89 and 1.60. The Strouhal number is 0.028. Results are similar at 18 deg, with a Strouhal number of 0.033. A leading edge vortex is shed periodically near maximum lift. Dynamic mesh solutions for unstalled airfoil flows show general agreement with experimental pressure coefficients. However, moment coefficients and the maximum lift value are underpredicted. The deep stall case shows some agreement with experiment for increasing angle of attack, but is only qualitatively comparable past stall and for decreasing angle of attack.

Beschreibung: The application of unsteady 3-D Euler and Navier-Stokes equations to transonic flow past rotor blades, and wing-alone configurations is described. A promising approach for the numerical solution of these equations is examined. Additional work is needed for improving the efficiency of the present procedure. It is hoped that the techniques presented will find use in fixed and rotary wing aircraft analysis.

Beschreibung: The Euler code is used extensively for computation of transonic unsteady aerodynamics. The boundary layer code solves the 3-D, compressible, unsteady, mean flow kinetic energy integral boundary layer equations in the direct mode. Inviscid-viscous coupling is handled using porosity boundary conditions. Some of the advantages and disadvantages of using the Euler and boundary layer equations for investigating unsteady viscous-inviscid interaction is examined.

Beschreibung: Recent experience in calculating unsteady transonic flow by means of viscous-inviscid interactions with the XTRAN2L computer code is examined. The boundary layer method for attached flows is based upon the work of Rizzetta. The nonisentropic corrections of Fuglsang and Williams are also incorporated along with the viscous interaction for some cases and initial results are presented. For unsteady flows, the inverse boundary layer equations developed by Vatsa and Carter are used in a quasi-steady manner and preliminary results are presented.

Beschreibung: Since emphasis is on the transonic speed range, special importance is placed on configurations for which available data are sufficient to define accurately a transonic flutter boundary. Only configurations with clean, smooth surfaces are considered suitable. Segmented models or models with surface-slope discontinuities are inappropriate. Excluded also, in general, are configurations and data sets that involve behavior that is uncertain or not well understood, uncertain model properties, or know sensitivities to small variations in model properties. In order to assess the suitability of configurations already tested and the associated data for designation as standard, a survey of AGARD member countries was conducted to seek candidates for the prospective set. The results of that survey are given and summarized along with the initial selection of a standard configuration.

Beschreibung: An implicit, two factor, split flux, finite volume Euler equations solution algorithms is applied to the time accurate solution of transonic flow about an NACA 0012 airfoil and a rectangular planform supercritical wing undergoing pitch oscillations. Accuracy for Courant numbers greater than one is analyzed. Freezing the flux Jacobians can result in significant savings for steady state solutions; the accuracy of freezing flux Jacobians for unsteady results is investigated. The Euler algorithm results are compared with experimental results for an NACA 0012 and a rectangular planform supercritical wing.

Beschreibung: The method of flux vector splitting used is that of Van Leer. The fluxes split in this manner have the advantage of being continuously differentiable at eigenvalue sign changes and this allows normal shocks to be captured with at most two interior zones, although in practice only one zone is usually observed. The fluxes as originally derived, however did not include the necessary terms appropriate for calculations on a dynamic mesh. The extension of the splitting to include these terms while retaining the advantages of the original splitting is the main purpose of this investigation. In addition, the use of multiple grids to reduce the computer time is investigated. A subiterative procedure to eliminate factorization and linearization error so that larger time steps can be used is also investigated.

Beschreibung: An efficient and accurate transonic unsteady aerodynamic method is needed for predicting flight loads, flutter and aeroelastic stability for advanced aircraft. There have been new developments and many improvements to older codes. XTRAN3S was improved and is a useful code for the near term. However, to predict the unsteady aerodynamics for high performance maneuvering aircraft, the Euler/Navier-Stokes codes must be extended and improved for complex 3-D configurations. The long term goal is development of Euler/Navier-Stokes unsteady aerodynamic methods for aeroelastic analysis.

Beschreibung: Since ground based flow simulations are presently unable to model flight conditions expected for AOTVs (Aeroassist Orbital Transfer Vehicle) and other hypersonic space vehicles, computer codes are being developed to provide design parameters necessary for structure, guidance, and control aspects. Over the past four years, VRFLO (Viscous Reactive Flow) has been written to model finite-rate chemistry and viscous effects for a variety of aerobrake bodies. VRFLO includes a number of unique features that are summarized as follows: (1) Grid generation is an integral part of the code for several aerobrake configurations which includes the wake flow region; (2) The formulation is valid for three air chemical models; (3) An ADI central difference technique is used to solve the Navier-Stokes and species continuity equations in split groups; and (4) Grid density and numerical damping are minimized by shock-fitting and conformal mapping of body points.

Beschreibung: As the capability of the space transportation vehicles (STV's) expand to meet the requirements for future space exploration and utilization, the effects of rarefied hypersonic flows will play a more significant role in defining the aerodynamic and aerothermodynamic performance of STV's. This is particularly true of the low lift/drag aeroassisted STV's where aerobraking occurs at relatively high altitudes and high velocity. Because of the limitations of the continuum description as expressed by the Navier-Stokes equations and the difficulties of solving the Boltzmann equation, the particle of molecular approach has been developed over the last three decades for modeling rarefied gas effects. The direct simulation Monte Carlo (DSMC) method of Bird is the most used method today for simulating rarefied flows. The DSMC method provides a direct physical simulation as opposed to a numerical solution of a set of model equations. This is accomplished by developing phenomenological models of the relevant physical events. The DSMC method accounts for translational, thermal, chemical, and radiative nonequilibrium effects. The general features of the DSMC method, the numerical requirements for obtaining meaningful results, the modeling used to simulate high temperature gas effects, and applications of the method to calculate the flow about an aeroassist flight experiment vehicle (AFE) are reviewed. The AFE simulates a geosynchronous return while entering the Earth's upper atmosphere at approximately 10 km/s. Results obtained using a general 3-D code are presented for the more rarefied portion of the atmospheric encounter (altitudes of 200 to 100 km) emphasizing surface, flowfield, and aerodynamic characteristics of the AFE. Finally, results obtained using axisymmetric and 1-D versions of the code are presented for lower altitude conditions.

Beschreibung: The code development and application program for the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA), with emphasis directed toward support of the Aeroassist Flight Experiment (AFE) in the near term and Aeroassisted Space Transfer Vehicle (ASTV) design in the long term is reviewed. LAURA is an upwind-biased, point-implicit relaxation algorithm for obtaining the numerical solution to the governing equations for 3-D, viscous, hypersonic flows in chemical and thermal nonequilibrium. The algorithm is derived using a finite volume formulation in which the inviscid components of flux across cell walls are described with Roe's averaging and Harten's entropy fix with second-order corrections based on Yee's Symmetric Total Variation Diminishing scheme. Because of the point-implicit relaxation strategy, the algorithm remains stable at large Courant numbers without the necessity of solving large, block tri-diagonal systems. A single relaxation step depends only on information from nearest neighbors. Predictions for pressure distributions, surface heating, and aerodynamic coefficients compare well with experimental data for Mach 10 flow over an AFE wind tunnel model. Predictions for the hypersonic flow of air in chemical and thermal nonequilibrium over the full scale AFE configuration obtained on a multi-domain grid are discussed.

Beschreibung: Several conceptual designs for vehicles that would fly in the atmosphere at hypersonic speeds have been developed recently. For the proposed flight conditions the air in the shock layer that envelops the body is at a sufficiently high temperature to cause chemical reaction, vibrational excitation, and ionization. However, these processes occur at finite rates which, when coupled with large convection speeds, cause the gas to be removed from thermo-chemical equilibrium. This non-ideal behavior affects the aerothermal loading on the vehicle and has ramifications in its design. A numerical method to solve the equations that describe these types of flows in 2-D was developed. The state of the gas is represented with seven chemical species, a separate vibrational temperature for each diatomic species, an electron translational temperature, and a mass-average translational-rotational temperature for the heavy particles. The equations for this gas model are solved numerically in a fully coupled fashion using an implicit finite volume time-marching technique. Gauss-Seidel line-relaxation is used to reduce the cost of the solution and flux-dependent differencing is employed to maintain stability. The numerical method was tested against several experiments. The calculated bow shock wave detachment on a sphere and two cones was compared to those measured in ground testing facilities. The computed peak electron number density on a sphere-cone was compared to that measured in a flight test. In each case the results from the numerical method were in excellent agreement with experiment. The technique was used to predict the aerothermal loads on an Aeroassisted Orbital Transfer Vehicle including radiative heating. These results indicate that the current physical model of high temperature air is appropriate and that the numerical algorithm is capable of treating this class of flows.

Beschreibung: The Roe flux difference splitting method was extended to treat 2-D viscous flows with nonequilibrium chemistry. The derivations have avoided unnecessary assumptions or approximations. For spatial discretization, the second-order Roe upwind differencing is used for the convective terms and central differencing for the viscous terms. An upwind-based TVD scheme is applied to eliminate oscillations and obtain a sharp representation of discontinuities. A two-state Runge-Kutta method is used to time integrate the discretized Navier-Stokes and species transport equations for the asymptotic steady solutions. The present method is then applied to two types of flows: the shock wave/boundary layer interaction problems and the jet in cross flows.

Beschreibung: Mesh generation procedures as well as solution algorithms for solving the Euler and Navier-Stokes equations on unstructured meshes are presented. The solution algorithms discussed utilize approximate Riemann solver, upwind differencing to achieve high spatial accuracy. Numerical results for Euler flow over single and multi-element airfoils are presented.

Beschreibung: The advancing front technique is being used to develop a code to generate grids around complex 3-D configurations for use in computing the invisid flow solutions by the Euler equations. By the advancing front technique points are introduced concurrently with the connectivity information so that a separate library is not required. The generation of a 3-D grid is accomplished in several steps. First the boundaries of the domain to be gridded must be described by two-, three- or four-sided surface patches. Next, a background mesh is required to control the grid spacing and stretching throughout the domain. This coarse tetrahedral grid is not required to conform to any of the boundaries. Next, each of the patches is mapped to 2-D, triangulated by the advancing front technique and mapped back to 3-D. These triangles form the initial front for the generation of the final tetrahedral mesh.

Beschreibung: A method for designing wings and airfoils at transonic speeds using a predictor/corrector approach was developed. The procedure iterates between an aerodynamic code, which predicts the flow about a given geometry, and the design module, which compares the calculated and target pressure distributions and modifies the geometry using an algorithm that relates differences in pressure to a change in surface curvature. The modular nature of the design method makes it relatively simple to couple it to any analysis method. The iterative approach allows the design process and aerodynamic analysis to converge in parallel, significantly reducing the time required to reach a final design. Viscous and static aeroelastic effects can also be accounted for during the design or as a post-design correction. Results from several pilot design codes indicated that the method accurately reproduced pressure distributions as well as the coordinates of a given airfoil or wing by modifying an initial contour. The codes were applied to supercritical as well as conventional airfoils, forward- and aft-swept transport wings, and moderate-to-highly swept fighter wings. The design method was found to be robust and efficient, even for cases having fairly strong shocks.

Beschreibung: The increased national interest in high speed flight has increased research for high speed propulsion components. The highly 3-D flows present in supersonic/hypersonic inlets are currently being studied at NASA-Lewis both experimentally and computationally using a family of steady Parabolized Navier-Stokes (PNS) and Navier-Stokes (NS) solvers and unsteady NS solvers. Some of the results of these efforts are presented with an emphasis on the comparison of the computational and experimental results. The flow in high speed inlets typically involves the interaction of compression shock waves and boundary layers on the internal surfaces. The fundamentals of these interactions have been studied experimentally for many years, while more recently, computations have been used to study these complex 3-D flow fields. Attempts to control the flow through boundary layer bleed are being investigated computationally prior to wind tunnel experiments. The ultimate goal is the higher performing inlets required for high speed flight.

Beschreibung: The feasibility was determined of incorporating the Navier-Stokes computational code, CFL3D, into the supersonic wing design process. The approach taken is of two steps. The first step was to calibrate CFL3D against existing experimental data sets obtained on thin sharp edged delta wings. The experimental data identified six flow types which are dependent on the similarity parameters of Mach number and angle of attack normal to the leading edge. The calibration showed CFL3D capable of simulating these various separated and attached flow conditions. The second step was to use CFL3D to study the initial formation of leading edge separation over delta wings at supersonic speeds. This consisted of examining solutions obtained on a 65 deg delta wing at Mach number of 1.6 with varying cross sectional shapes. Reynolds number was held constant at 1000000 and the Baldwin-Lomax turbulence model was used. The study showed that through the use of leading edge radius and/or camber, the onset of leading edge separation can be delayed to a higher angle of attack than observed on a flat sharp edged wing. Based on the geometries studied, three wind tunnel models are being designed to verify these results.

Beschreibung: An embedded grid algorithm for the Euler and/or Navier-Stokes equations is developed and applied to delta wings at high angles of attack in low speed flow. The Navier-Stokes code is an implicit, finite volume algorithm, using flux difference splitting for the convective and pressure terms and central differencing for the viscous and heat transfer terms. Calculations are compared with detailed experimental results over an angle of attack range up to and beyond the maximum lift coefficient, corresponding to vortex breakdown at the trailing edge, for a delta wing nominally of unit aspect ratio. The results indicate that the overall flowfield, including surface pressures, surface streamlines, and vortex trajectories, can be simulated accurately with the global grid version of the present algorithm. However, comparison of computed velocities and vorticity with experimentally measured off-body values at an angle of attack of 20.5 deg indicates the core region is substantially more diffuse in the computations than that measured with either a five-hole probe or a laser velocimeter. Embedded grids, used to improve the numerical discretization in the core region, are formulated within the framework of the implicit, upwind-biased multi-grid algorithm. Structured levels of local nested refinements are made. Three-dimensional results for both Euler and Navier-Stokes calculations are shown, with up to 3 levels of embedded refinement. The embedding procedure was effective in eliminating a crossflow secondary separation produced in the Euler solutions on coarse grids.

Beschreibung: The development of Short Takeoff Vertical Landing (STOVL) aircraft has historically been an empirical- and experienced-based technology. A 3-D turbulent flow CFD code was used to calculate the hot gas environment around an STOVL aircraft operating in ground proximity. Preliminary calculations are reported for a typical STOVL aircraft configuration to identify key features of the flow field, and to demonstrate and assess the capability of current 3-D CFD codes to calculate the temperature of the gases ingested at the engine inlet as a function of flow and geometric conditions.

Beschreibung: The primary objective is to expose government, industry, and academic scientists to work underway at NASA-Ames towards the application of CFD to the powered lift area. One goal is to produce the technologies which will be required in the application of numerical techniques to, for example, the Supersonic STOVL program. The progress to date on the following specific projects is presented: Jet in ground effect with crossflow; Jet in a crossflow; Delta planform with multiple jets in ground effect; Integration of CFD with thermal and acoustic analyses; Improved flow visualization techniques for unsteady flows; YAV-8B Harrier simulation program; and E-7 simulation program.

Beschreibung: A 3-D Navier-Stokes code was developed for analysis of turbomachinery blade rows and other internal flows. The Navier-Stokes equations are written in a Cartesian coordinate system rotating about the x-axis, and then mapped to a general body-fitted coordinate system. Streamwise viscous terms are neglected using the thin layer assumption, and turbulence effects are modelled using the Baldwin-Lomax turbulence model. The equations are discretized using finite differences on stacked C-type grids and are solved using a multistage Runge-Kutta algorithm with a spatially varying time step and implicit residual smoothing. Calculations were made of the flow around a supersonic throughflow fan blade. The fan was designed as a key component in a supersonic cruise engine. The 3-D calculations were done on a 129x29x33 grid and took 50 minutes of cpu time. Comparisons with the quasi-3-D results show minor differences in loading due to 3-D effects. Particle traces show nearly 2-D flows near the pressure surface, but large secondary flows within the suction surface boundary layer. The horseshoe vortex ahead of the leading edge is clearly seen.