A Stall Flutter Of Helicopter Rotor Blades.- A Special Case 01 The Dynamic Stall Phenomenon MASSACHUSETTS INSTITUTE CF TECHNOLOGY.

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$PROJECT 20014501B33C pi; r \^*. [ RECEIVED JUL2 6 1967 CFSTI U. S.$

1 - AD REPORT 1*81^.3 CO A Stall Flutter Of Helicpter Rtr Blades.- A Special Case 01 The Dynamic Stall Phenmenn NORMAN D. HAM MASSACHUSETTS INSTITUTE CF TECHNOLOGY MAY I967 D D C m Spa<iMf*d by n DA PROJECT B33C pi; r \^*. [ RECEIVED JUL CFSTI U. S. Army Research Office - Durham Distributin f this reprt is unlimited ^

STALL FLUTTER OF HELICOPTER ROTOR BLADES: A SPECIAL CASE OF THE DYNAMIC STALL PHENOMENON* Nrman D. Ham Maaaachusetts Institute f Technlgy I ntrductin Recent studies described in Refs.

The mechanism f the instability was shwn in Ref. 2 t cnsist f the adverse time phasing f the aerdynamic pitching mment assciated with the lss f blade bund vrticity as the dynamic stall ccurs.

2 STALL FLUTTER OF HELICOPTER ROTOR BLADES: A SPECIAL CASE OF THE DYNAMIC STALL PHENOMENON* Nrman D. Ham Maaaachusetts Institute f Technlgy I ntrductin Recent studies described in Refs. 1 and t have shwn that the negative damping in pitch assciated with airfils scillating at h ; gh mean angles f attack can lead t trsinal instability f helicpter rtr blades under certain cnditins. The mechanism f the instability was shwn in Ref. 2 t cnsist f the adverse time phasing f the aerdynamic pitching mment assciated with the lss f blade bund vrticity as the dynamic stall ccurs. Subsequent unpublished tests, Ref. 3, have indicated that the same mechanism is fund fr the case f airfil linear angle f attack change thrugh high angles f attack. The nature f these results indicates several imprtant cnclusins nt nly with respect t the analysis f rtr blade stall flutter inatalility, but als with respect t the general nature f the aerdynamic lading f an airfil experiencing transient angle f attack changes f large magnitude. Discussin A typical time histry f the pressure variatin acting at ne spanwise statin f a mdel helicpter rtr blade experiencing stall-induced scillatins while perating in the static thrust cnditin is shwn ir Fig. 1, reprduced frm Ref. 2. The figure shws the initiatin f a pressure disturbance in the regin f the leading edge as the blade sectin appraches maximum angle f attack, and the subsequent mtin f this disturbance in the chrdwise directin, at cnsiderably l^ss than free stream velcity. The characte- f the disturbance suggests that it cnsists f free vrticity intrduced int the blade flw field frm the neighbrhd f the blade leading edge as the blade lses bund vrticity during the dynamic stall prcess. The results indicate that the dynamic stall phenmenn has far different characteristics than thse assciated with the classic trailing edge separatin during static stall f a blunt airfil. Subsequent integratin f the pressures shwn in Fig. 1 led t the quarter chrd pitching mment time histry shwn in Fig. 2, This figure illustrates the rigin f the stall flutter * This research was spnsred by the U. 5. Army Research Office - Durham, Nrth Carlina. Paper presented befre the American Helicpter Sciety 23rd Annual Natinal Frum, Washingtn, D. C., May instability. The negative pressure peak generated by the pressure disturbance mving aft frm the leading edge leads t a nse dwn pitching mment cmpnent in phase with the nse dwn mtin f the airfil. Since this nse dwn mment is generated nce per pitching cycle, it is seen that the nnlinear aerdynamic mment variatin due t pitching mtin at high mean angles f attack is such as t sustain the mtin. This self-excited, self-limiting mtin is termed "stall flutter". Anther methd f presenting the pitching mment data is shwn in Fig. 3, in terms f pitching mment versus pitching angle. It is seen that the nse dwn mment ccurring in phase with the airfil mtin leads t the hysteresis lp in the mment curve. It can be shwn that the area f cunter-clckwise lps in the curve represents energy remved frm the mtin, while the area f clckwire lps represents energy added t the mtin. When the area f the tw lps is equal fr a given blade sectin, that sectin is capable f self-sustained mtin at the given amplitude and reduced frequency. The abve results suggested that the same stall mechanism wuld be fund in the general case f large transient blade angle f attack changes. Accrdingly, an experimental investigatin was cmmenced at MIT t study large linear angle f attack changes f a twdimensinal wing (Reference 3). Preliminary results are shwn in Figs. 4 t 6 fr the case f a nearly linear angle f attack change t a maximum angle f thirty degrees (subsequently held cnstant) at a maximum pitching velcity f ten radians per secnd. The variatin f airfil lift cefficient with time and assciated angle f attack is shwn in Fig. 4. Cmparisn f the dynamic lilt variatin with the crrespnding values f static lift at the same angles f attack indicates that the maximum dynamic lift achieved is substantially higher than the maximum static lift, and that a high value f lift is sustained fr several hundredths f a secnd (apprximately equivalent t ne eighth f a rtr revlutin) subsequently drpping ff t the steady state value. The crrespnding pitching mment variatin is shwn in Fig. 5. A very large, sustained, transient nse dewn mment is seen t ccur in the high angle f attack regin. The rigin f these effects is shwn in " s * iia =*, v**«*kws<i/

"i"^ tike crrespnding cbrdwiae pressure variatins shwn in Fig 6. Dynamic stall begins t ccur at cj = 20, a«indicated by the drp in leading edge suctin f the «* = <Ji. 6 * pressure trace.

2, the pressure disturbance has mved further aft, and is still f cnsiderable magnitude.

3 "i"^ tike crrespnding cbrdwiae pressure variatins shwn in Fig 6. Dynamic stall begins t ccur at cj = 20, a«indicated by the drp in leading edge suctin f the «* = <Ji. 6 * pressure trace. A negative pressure disturbance mving aft frm the leading edge simultaneusly increases the suctin in the midchrd regin. Subsequently, at a( = 29.2, the pressure disturbance has mved further aft, and is still f cnsiderable magnitude. The delay in the ccurrence f stall (as evidenced by lss f leading edge suctin) due t tiie high rate f change f angle f attack and the sustained upper surface suctin assciated with the chrdwise passage f the vticity shed during the stall prcess, bth cntribute t the high sustained lift values f Fig. 4. In additin, the increasingly aft center f pressure due t the aft mtin f the shed vrticity generates the extreme nse dwn pitching mment f Fig. 5. Finally, the nature f the pressure distributin is such as t indicate greatly increased pressure drag n the airfil. This drag may be a transient analg f the "vrtex drag*' due t the leading edge vrtex f a slender delta wing at lw speed and high angle f attack. Cudusins The abve results lead t several imprtant cnclusins with respect t stall flutter and airlad predictin f high speed and/r highly laded helicpter rtr blades. 1. The stall f an airfil sectin during rapid transient high angle f attack changes is delayed well abve the static stall angle and results in a large transient negative pressure disturbance leading t large transient lift and nse dwn pitching mment. i. The magnitude f the pitching mment ffi^ is such as t generate substantial nse dwn pitching displacements f the lade. These pitching displacements can substantially alter the angle f attack distributin f the rtr blade, based n steady state cnditins used in the stall flutter analyses f Refs. 1 and i. It appears that transient pitching dieplacement f the blade in respnse t the initial stall-induced pitching mment acting n the blade shuld be included in subsequent stall flutter analyses. 3. The dynamic stall phenmenn f a helicpter rtr blade can be separated int three majr phases: a, A delay in the lss f blade leading edge suctin t an argle f attack far abve the static stall angle, with assciated airlads f the type described by classical unsteady airfil thery. b. A subsequent lss cf leading edge suctin accmpanied by the frmatin f large negative pressure disturbance (due t the shedding f vrticity frm the vicinity f the blade leading edge) which mves aft ver the upper surface f the blade. Assciated with this phase are high transient lift, drag, and nse-dwn pitching mment assciated with the greatly altered pressure distributin n the airfil. e. Cmplete upper surface separatin f the classic static type, characterized by lw lift, high drag, and mderate nse-dwn pitching mment. References 1. Carte, F. O., "An Analysis f the Stall Flutter Instability f Helicpter Rtr Blades, " Prceedings f the 23rd Annual Natinal Frum, American Helicpter Sciety, May Ham, N. D. and Yung, M. I., "Trsinal Oscillatin f Helicpter Blades Due t Stall, " Jurnal f Aircraft, AIAA, 3, 3, May-June Carelick, M. May M. Thesis, M. I. T.

4 '. -- ' 4 SUCTION, d y?^ TORSION FIGURE i. PRESSURE DATA

5 z 5 a: Z ÜJ a. nl_ ^ TIME - SECONDS FIGURE 2 < CD CD 0.05 O X u H FIGURE 3

;-«f& iiiax'io RAD./SEC «23.4f, <n>a=25.3 0 NACA 0012 R.N.=3.

U 0» 132 FT/SEC «-^-«M 0*292' c = 5 IN. u. u. Ü u..03.

6 ;-«f& iiiax'io RAD./SEC «23.4f, <n>a= NACA 0012 R.N.=3.5 x I0 9 ROTATION AXIS 0.25C. c. U 0» 132 FT/SEC «-^-«M 0*292' c = 5 IN. u. u. Ü u TIME t-sec.07 FIGURE 4. LIFT COEFFICIENT vs TIME AND ANGLE OF ATTACK...*..,.-. --;.

7 TIME t-sec FIGURE 5. PITCHING MOMENT COEFFICIENT vs TIME AND ANGLE OF ATTACK.-.i«w* *. * v

< -t 6f a«l9.5 0 Z QC Ui U. U..2.3.4.5.6.7.8.

8 Will imm.!.ji.m *«Ml I r4 r "- sz&mt -J I I 60 amax-wrad./sec. NACA 0012 R.N. = 3.5 x I0 5 t«0-2912* c CO en a: a. < -t 6f a«l9.5 0 Z QC Ui U. U I CHORDWISE STATION FIGURE 6. DIFFERENTIAL PRESSURE DISTRIBUTION vs TIME

g**-^ uncxaasiriea Security a iflcatin f DOCUMENT CONTROL DATA R&D (SMurilp el IHeallan l lilt», bdr f abalnead Induing anntamn must 6«mnltrtd **i»n tht mra/l»prt fa el»»»lll»d) I.

A STALL FLUTPER OF HELICOPTER ROTOR BLADES: A SPECIAL CASE OF THE DYNAMIC. STALL PHENOMENON «OCSCRIPTIVC NOTEI (Typ» l «prt and Incluiln dt»») Reprint t. AUTMORfSJ rt-a»«nama. Hmt nm.

9 g**-^ uncxaasiriea Security a iflcatin f DOCUMENT CONTROL DATA R&D (SMurilp el IHeallan l lilt», bdr f abalnead Induing anntamn must 6«mnltrtd **i»n tht mra/l»prt fa el»»»lll»d) I. ONIOINATIN0 ACTIVITY (Cutptumt» authr} Massachusetts Institute f Technlgy $ NIPOftT TITLI 2a. MCPORT SCCURITV CLASSIFICATION Unclassified 2b «nu» J&. A STALL FLUTPER OF HELICOPTER ROTOR BLADES: A SPECIAL CASE OF THE DYNAMIC. STALL PHENOMENON «OCSCRIPTIVC NOTEI (Typ» l «prt and Incluiln dt»») Reprint t. AUTMORfSJ rt-a»«nama. Hmt nm. MUl) Ham, Nrman D. «. REFOUT UATC May 1967 a. CONTRACT OR SRANT NO. DA ARO-D-247 7a TOTAL NO. OP PACKS 7 9a ORIGINATOR'S REPORT NUMBERfS^ 7b. NO. r RCPS 3 *- PROJECT NO b. OTHER REPORT NOfSJ (Any tftar numbara (hat may t>»»»»ignd thi* rprt) 10. AVAILAaiLITY/UIMITATIOM NOTICES 48^6.3 Distributin f this reprt is unlimited. 11 SUPPLCMENTANV NOTES Hne 12 SPONSORING MILITARY ACTIVITY U.S. Army Research Office-Durham Bx CM, Duke Statin Durham, Nrth Carlina II ABSTRACT Recent studies have shwn that the negative damping in pitch assciated with airfils scillating at high mean angles f attack can lead t trsinal instability f helicpter rtr blades under certain cnditins. The mechanism f the instability was shwn t cnsist f the adverse time phasing f the aerdynamic pitching mment assciated with the lss f blade bund vrticity as the dynamic stall ccurs. Subsequent unpublished tests have indicated that the same mechanism is fund fr the- case f airfil linear angle f attack change thrugh high angles f attack.. The nature f these results indicates several Imprtant cnclusins nt, nly with respect t the analysis f rtr blade stall flutter instability, but als with respect t the general nature f the aerdynamic lading f an airfil experiencing transient angle f attack changes f large magnitude. in. KLY V::RD:: Ctall flutter helicpters rtr blades airfils angle f attack DD FOftM 1 JAN Unclassified Security Classificatin J

Introductory Thoughts

Introductory Thoughts Flw Similarity By using the Buckingham pi therem, we have reduced the number f independent variables frm five t tw If we wish t run a series f wind-tunnel tests fr a given bdy at a given angle f attack,